Repair Manual: Trainer Airplanes, Series AT-6A, AT ... - SprueMaster

REPORT NO. NA.5614J-N]i^^ ..f^--^CJi^TRAINERAIRPLANESSERIESAT-6A, AT.6B, AT.6CSNJ-3andSNJ-4NORTH AMERICAN AVIATION, INDALLAS iNGLEWOOO

Digitized by tliein 2010Internet Arciiivelittp://www.arcliive.org/details/repairmanualtraiOOnort

TO THE HOLDER of THIS BOOKIfat anytime contact cannot be madewith the localN. A. A. representative,write to the field service department,NORTH AMERICAN AVIATION, INC.,INGLEWOOD, CALIFORNIA,regarding the representatives location.

RestrictedREPORT NO. NA-5614TITLEHandbook ofInstructionsfor the-Structural /v^epuirJfL "^of^ tne ^exunTRAINER AIRPLANESSERIES: AT-6A, AT-6B, AT-6C, SNJ-3 AND SNJ-4CONTRACTSAC-12969 CHANGE NO. 8AC-15977W535 AC-19192W535 AC-29319DA W535 AC-8DA-W535 AC-2799NORTH AMERICAN AVIATION. INCDALLAS INGLEWOOD KANSAS CITYAanuarii 1 8y1^43

NGRESTRICTEDCONTENTSTABLEOF CONTENTSPar. Page Par. PageSECTION IGENERAL INSTRUCTIONS1 AIRPLANE CONSTRUCTION 12 RIVET TYPES 13 GENERAL RIVETING EQUIPMENT 14 DR'LLING 15 STANDARD HOLES FOR RIVETS 56 DRILL SIZES 57 RIVET COUNTERSINKING DIMENSIONS .... 68 CUT COUNTERSINKING 69 DIMPLE COUNTERSINKING 610 PRERIVETING FASTENERS AND CLAMPS ... 611 BUCKING BARS 712 RIVET SETS 713 PNEUMATIC RIVET GUN . 814 HAMMERS 815 RIVET SQUEEZER 816 REMOVAL OF SOLID RIVETS 817 RIVET AND BOLT EDGE DISTANCES 918 RIVETING 919 FITTING OF BOLTS AND SCREWS 1120 TYPES OF BOLTS, SCREWS, AND NUTS ... 1221 STANDARD HOLES FOR BOLTS AND SCREWS . . 1222 ALUMINUM CASTINGS 1223 ALUMINUM SHEET MARKINGS 1224 ALUMINUM WROUGHT STOCK - GENERAL ... 1725 ALUMINUM SHEET, 2S0 1726 ALUMINUM SHEET, 2S-1/2H 1727 ALUMINUM SHEET, 3S0 1728 ALUMINUM SHEET, 3S-1/2H 1729 ALUMINUM SHEET AND TUBE, 52S0 1730 ALUMINUM SHEET, 52S-1/2H 1831 ANNEALING ALUMINUM, 2S, 3S, AND 52S . . 1832 ALUMINUM EXTRUSIONS, 24S0 1833 ALUMINUM EXTRUSIONS, 24ST 1834 ALUMINUM SHEET, 24S0 ALCLAD 1835 ALUMINUM SHEET, 24ST ALCLAD 1936 HEAT TREATING ALUMINUM, 24S0 1937 QUENCHING 1938 AGE HARDENING 2039 ANNEALING ALUMINUM, 24ST 2040 BENDING 24S ALCLAD 2041 EXTRUSION EQUIVALENT DIE NUMBERS ... 2242 SCRATCHES, DENTS, AND CRACKS 2243 SUPPORT OF STRUCTURE DURING REPAIR . . 2244 EXTENT OF DAMAGE 2245 CORROSION PROTECTION 2246 TOOLS - GENERAL 2247 METAL SHRINKER OR STRETCHER 2348 HAND ROLLER 2349 HAND BRAKE 2350 HAND METHODS OF BENDING 2551 ARBOR PRESS 2552 FORMING BLOCKS 2553 TRIMMING MACHINE 2654 SPIRAL REAMER 2655 STANDARD POWER AND HAND TOOLS .... 2656 STANDARD METAL- WORK I HAND TOOLS . . 3057 ALUMINUM ALLOY MATERIAL SPECIFICA-TION EQUIVALENTS 3258 STEEL ALLOY MATERIAL SPECIFICATIONEQUIVALENTS 3259 COPPER ALLOY MATERIAL SPECIFICATIONEQUIVALENTS 3360 WIRE AND CABLE MATERIAL SPECIFICATIONEQUIVALENTS 33SECTION 2FUSELAGE1 FUSELAGE GENERAL CONSTRUCTION .... 342 CONSTRUCTION OF ENGINE MOUNT ANDFRONT FUSELAGE TRUSS 343 WELDING STEEL PARTS" GENERAL .... 364 ELECTRIC ARC WELDING OF STEEL .... 365 OXYACETYLENE WELDING OF STEEL .... 376 CONDITION OF COMPLETED WELDS 387 GENERAL REPAIR OF STEEL TUBES .... 398 NEGLIGIBLE DAMAGE TO STEEL TUBES ... 409 ESTIMATING EXTENT OF DAMAGE TO STEELTUBES 4010 SMOOTH DENTS IN STEEL TUBES 4011 STEEL TUBE CIRCUMFERENCE BENT TO ANOVAL SHAPE 4112 BOWED STEEL TUBES 4113 SMALL CRACKS AT STEEL TUBING CLUSTERJOINTS 4314 SHARP DENTS AT A STEEL TUBE CLUSTERJOINT 4315 SHARP DENTS OR CRACKS IN STEEL TUBELENGTH 4416 SPLICING STEEL TUBES - GENERAL .... 4417 SPLICING STEEL TUBE BY INNER SLEEVEMETHOD 4618 SPLICING STEEL TUBE BY OUTER SLEEVEMETHOD 4719 SPLICING STEEL TUBE USING LARGERDIAMETER REPLACEMENT TUBE 4720 REPLACING STEEL TUBES 4921 TESTING WELDED STEEL JOINTS 5122 PROTECTI NG STEEL TUBE AGAINST CORROSION 5123 ALIGNMENT OF ENGINE MOUNT AND FRONTFUSELAGE TRUSS 5124 EXHAUST MANIFOLD ASSEMBLY CONSTRUCTION 51[ i ]RESTRICTED

CONTENTSRESTRICTEDPar. Page Par. Page25 EXHAUST MANIFOLD REPAIR 5226 EXHAUST MANIFOLD COLLAR REPLACEMENTS 5.227 EXHAUST MANIFOLD REASSEMBLY 5228 ENGINE FIREWALL CONSTRUCTION .... 5229 FIREWALL WEB REPAIR 5330 FIREWALL STIFFENER REPAIR 5431 COWLING ATTACHING FIREWALL ANGLEREPAIR 5432 REPAIR MATERIALS FOR STEEL STRUCTURES 5433 REPAIR TOOLS FOR STEEL STRUCTURES . . 5534 ENGINE COWLING CONSTRUCTION 5535 ENGINE COWLING REPAIR 5736 CONSTRUCTION OF FRONT FUSELAGE SIDEPANELS 5737 FORWARD FORMER OF FRONT FUSELAGE SIDEPANELS 5738 REAR FOUR FORMERS OF FRONT FUSELAGESIDE PANELS 6039 ALUMINUM REAR FUSELAGE CONSTRUCTION . 6040 ALUMINUM REAR FUSELAGE FORMERS ... 6141 REPLACING ALUMINUM FORMERS 6142 ALUMINUM STRINGER TYPE C107LT-40 . . 6243 ALUMINUM STRINGER TYPE C180T .... 6344 ALUMINUM STRINGER TYPE C234LT .... 6345 ALUMINUM STRINGER TYPE C364LT .... 6346 ALUMINUM STRINGER TYPE C366T .... 6447 ALUMINUM REAR FUSELAGE BAGGAGE COM-PARTMENT INTERCOSTAL ANGLE .... 6448 ALUMINUM REAR FUSELAGE UPPER LONGERONS 6449 ALUMINUM REAR FUSELAGE LOWER LONGERONS 6650 ALUMINUM SKIN AND BULKHEAD WEBS ... 6851 FAIRINGS AND FILLETS 6852 EXTRUDED ALUMINUM REPAIR MATERIAL . . 6853 ALCLAD SHEET REPAIR MATERIAL .... 7054 RIVETS REQUIRED FOR ALUMINUM REPAIRS 7055 MISCELLANEOUS ALUMINUM REPAIR MATERIAL 7156 REPAIR TOOLS FOR ALUMINUM STRUCTURES 7157 WOODEN REAR FUSELAGE CONSTRUCTION . . 7158 REPAIR WOOD MOISTURE CONTROL .... 7359 SUBSTITUTES FOR SPECIFIED WOODS ... 7460 CASEIN GLUE FOR WOOD REPAIRS .... 7561 MIXING CASEIN GLUE FOR WOOD REPAIRS . 7562 ASSEMBLING GLUED JOINTS OF WOOD ... 7563 APPLICATION OF WOOD-GLUING PRESSUREWITH CLAMPS 7664 APPLICATION OF WOOD-GLUING PRESSUREWITH NAI LING STRIPS 7665 WOODEN REAR FUSELAGE STRINGERS ... 7766 WOODEN REAR FUSELAGE BAGGAGE COMPART-MENT INTERCOSTAL 7967 WOODEN REAR FUSELAGE UPPER LONGERONS 8068 WOODEN REAR FUSELAGE LOWER LONGERONS 8269 FLUSH PATCH FOR PLYWOOD SKIN HOLESLESS THAN 2 INCHES WIDE 8470 EXTERNAL PATCH FOR PLYWOOD SKIN HOLESLESS THAN 2 INCHES WIDE 8471 REPLACING DAMAGED PLYWOOD SKIN SECTION72 BULKHEAD PLYWOOD WEBS73 WOODEN REAR FUSELAGE FORMERS ....74 GENERAL FINISH REQUIREMENTS FOR WOOD75 WOOD FINISH MATERIALS76 WOOD REPAIR MATERIALS77 WOOD REPAIR TOOLSSECTION 3FIXED SURFACES1 WING CONSTRUCTION 912 HORIZONTAL STABILIZER CONSTRUCTION . 913 VERTICAL STABILIZER CONSTRUCTION . . 914 STRINGERS - GENERAL 975 STRINGER REPAIRS - GENERAL 976 STRINGER TYPE C107LT-20 997 STRINGER TYPE CI23LT 1008 STRINGER TYPE C148T 1019 DOUBLED C148T TYPE STRINGERS .... 10210 STRINGER TYPE C180T 10411 STRINGER TYPE C204T 10412 DOUBLED C204T TYPE STRINGERS .... 10513 STRINGER TYPE C250T 10514 DOUBLED C250T TYPE STRINGERS .... 10615 COMBINATION C250T~C366T STRINGERS . . 10716 STRINGER TYPE C265T 10817 STRINGER TYPE C266T 10918 STRINGER TYPE C274T 10919 STRINGER TYPE C366T 11020 DOUBLED C366T TYPE STRINGERS .... Ill21 STRINGER TYPE C3731-T Ill22 STRINGER TYPE K77A 11223 DOUBLED STRINGER TYPES K77A 11224 WING TRAILING EDGE II325 SPAR REPAIRS - GENERAL II326 OUTER WING MAIN SPAR - ROOT TO 106INCHES OUTBOARD 11427 OUTER WING MAIN SPAR - 106 TO 151INCHES OUTBOARD OF ROOT 11428 OUTER WING MAIN SPAR - 151 INCHESOUTBOARD TO TIP 11629 OUTER WING FLAP SPAR 11930 OUTER WING AILERON SPAR 11931 CENTERSECTION FLAP SPAR 11932 CENTERSECTION REAR SPAR 12233 CENTERSECTION FRONT SPAR UPPER CAP . 12434 CENTERSECTION FRONT SPAR LOWER CAP . 12635 PATCHING CENTERSECTION FRONT SPAR WEB 12836 SPLICING CENTERSECTION FRONT SPAR WEB 13O37 VERTICAL STABILIZER FRONT SPAR ... 13O38 VERTICAL STABILIZER REAR SPAR AREAABOVE ROOT RIB I3O39 VERTICAL STABILIZER REAR SPAR AREABELOW ROOT RIB 13285868690RESTRICTED[ii]

RESTRICTEDCONTENTSPar.PagePar.Page40 REPAIR OF WOODEN HORIZONTAL STABILIZER 13241 HORIZONTAL STABILIZER FRONT SPAR ... I3242 HORIZONTAL STABILIZER REAR SPAR AREAOUTBOARD OF CENTER HINGE FITTING . . I3543 HORIZONTAL STABILIZER REAR SPAR AREAINBOARD OF CENTER HINGE FITTING- . • 13644 SKIN - GENERAL I3645 SMALL HOLES IN SKIN LESS THAN 1/2-fNCH DIAMETER I3746 SMALL HOLES IN SKIN 1/2- TO 1-INCHDIAMETERI3747 ISOLATED SKIN CRACKS 13848 PREPARING LARGE HOLES IN SKIN FORPATCHING 13849 PATCHING LARGE HOLES IN SKIN I3950 LOCATING BLIND RIVET HOLES IN SKIN . . 14051 SPLICING SKIN PANELS 14152 PREPARING WING LEADING EDGE SKIN FORSPLICING 14353 SPLICING WING LEADING EDGE SKIN. ... 14554 REMOVING AND REPLACING THE OUTER WINGCLOSING STRIP SKIN 14755 REMOVING AND REPLACING SKIN PANELS • . 14856 REMOVING DENTS IN THE HORIZONTAL STABI-LIZER TIP SKIN 14957 FUEL TANK COMPARTMENT COVER FORMERS- - 14958 CENTERSECTION INTERMEDIATE RIBS- - - . 15059 WING TRAILING EDGE RIBS AND CENTER-SECTION LEADING EDGE RIBS 15160 DAMAGED RIB BEADS 15261 DAMAGED RIB CUTOUTS 15262 BUCKLED RIB WEBS 15363 DAMAGED RIB FLANGES 15364 SPLICING TYPICAL RIBS 15465 REMOVING AND REPLACING DAMAGED RIBS- • 15566 WING JOINT BOLTING ANGLES 15567 WING JOINT COVERS 15768 EXTRUDED ALUMINUM REPAIR MATERIAL- - - 15869 ALCLAD SHEET REPAIR MATERIAL 15870 RIVETS, BOLTS, SCREWS, AND NUTS RE-QUIRED FOR REPAIR 15871 REPAIR TOOLS 159SECTION 4MOVABLE SURFACES1 GENERAL 1612 LANDING FLAP CONSTRUCTION 1613 RUDDER CONSTRUCTION 1614 ELEVATOR CONSTRUCTION 1635 AILERON CONSTRUCTION 1646 RUDDER, ELEVATOR, AND AILERON ORIGINALFABRIC COVERING 1647 CONTROL SURFACE TRIM AND BOOSTER TABS- 1668 FLAP REPAIR, GENERAL 1669 FLAP SKIN REPAIR 16610 FLAP CHANNEL SPAR 16711 FLAP LEADING EDGE 16812 FLAP TRAILING EDGE 16813 STRUCTURAL REPAIR OF RUDDER, ELEVA-TORS, AND AILERONS, GENERAL 16914 ELEVATOR AND RUDDER RIB REMOVAL. ... 16915 ELEVATOR AND RUDDER RIB REPLACEMENT. - 17016 ELEVATOR AND RUDDER LEADING EDGE SKINREPAIR 17017 RUDDER, ELEVATOR, AND AILERON TRAILINGEDGE STRIP REPLACEMENT 17118 AILERON SPAR SPLICE 17119 AILERON TRAILING EDGE RIB REPLACEMENT- 17220 AILERON NOSE RIB REPLACEMENT 17321 AILERON LEADING EDGE SKIN REPAIR ... 17522 GENERAL FABRIC PATCHING 17523 TESTING FABRIC FLEXIBILITY 17524 SMALL FABRIC TEARS 17625 V-SHAPED FABRIC TEARS 17626 LARGE TEARS AND HOLES IN FABRIC- - - - 17727 FABRIC SECTION REPLACEMENT 17728 PARTIAL FABRIC RE-COVERING 17829 REMOVAL OF ENTIRE FABRIC COVERING- - - 17930 PREPARING NEW FABRIC ENVELOPE 18331 ATTACHING OF THE NEW FABRIC ENVELOPECOVER 18332 HAND SEWING FABRIC ENVELOPE AFTERATTACHMENT18U33 DOPING FABRIC - GENERAL 18534 FABRIC DOPING PROCEDURE 18635 CONTROL SURFACE STATIC BALANCE .... 18636 DETERMINING CONTROL SURFACE UNBALANCE- 18737 CORRECTING UNBALANCE WHEN ABOVE LIMIT- 18838 METAL AND FABRIC REPAIR TOOLS 18939 METAL STRUCTURE REPAIR MATERIALS - - - 19140 RUDDER COVERING MATERIALS 19141 ELEVATOR COVERING MATERIALS 19242 AILERON COVERING REQUIREMENTS 193SECTION 5FUEL AND OIL SYSTEMS1 GENERAL 19U2 STRUCTURAL DESCRIPTION 19U3 GENERAL REPAIR 1944 cOCATiNG SMALL CRACKS 1955 SEALING SMALL CRACKS 1976 REMOVING DENTS 1987 SPLIT SEAM REPAIR 1988 LARGE CRACK AND HOLE REPAIR 1999 HYDRAULIC OIL TANK REPAIRS 19910 TESTING 19911 REPAIR TOOLS AND MATERIALS 199[iii ]RESTRICTED

CONTENTSRESTRICTEDPar. Page Par. PageOIL COOLERS12 GENERAL 20 213 OIL COOLER CONSTRUCTION 20214 CLEANING OIL COOLERS 20215 TESTING OIL COOLER FOR LEAKS .... 20216 REMOVING SINGLE TUBES 20317 REPLACING TUBES 20318 REMOVING LARGE SECTIONS OF CORE ... 20319 LEAKS ON CORE SURFACE 20320 LEAKS AROUND SHELL SEAMS 2O321 DENTS IN SHELL 20322 HOLES IN SHELL 20323 TESTING OIL COOLERS 2O32U OIL COOLER REPAIR MATERIALS 20525 OIL COOLER REPAIR TOOLS 205SECTION 6LANDING GEAR1 GENERAL 2062 MAIN LANDING GEAR 2063 TAIL WHEEL 2064 REMOVAL OF WHEEL ASSEMBLY 2075 REMOVAL AND INSPECTION OF TIRE ANDTUBE 2076 WHEEL INSPECTION AND REPAIR 2087 NEW CASING REPAIR 2088 CASING REPAIRS 2089 TREADING 20910 MARKINGS 20911 TUBE REPAIR 20912 TIRE AND TUBE BALANCE 20913 TUBE BALANCING 21014 TIRE BALANCING 21015 ASSEMBLING TIRE AND TUBE ON WHEEL . . 21116 MOUNTING TAIL WHEEL TIRE AND TUBE . . 21217 TOOLS 21518 MATERIALS 215SECTION 7CONTROL CABLES1 GENERAL 2162 NEGLIGIBLE DAMAGE 2163 CABLE REPLACEMENT 2164 CABLE FABRICATION 2165 MATERIAL SPECIFICATIONS 2166 CUTTING CABLES 2167 RUST PREVENTION 2168 SWAGED TERMINALS 2189 SWEAT-SOLDERED TERMINALS 22010 WOVEN SPLICED TERMINALS 22011 WRAPPED-SOLDERED TERMINALS 22112 TESTING 22213 COLOR BANDING 22214 TOOLS 222SECTION 8ELECTRICAL SYSTEM1 GENERAL 2232 WIRING 2233 WIRE NUMBERING 2234 WIRE REPLACEMENT 2235 WIRE TERMINALS 2246 ATTACHING CRIMPED TERMINALS 2247 CONDUIT 2258 BONDING 2259 ELECTRICAL REPAIR TOOLS AND MATERIALS 227SECTION 9MISCELLANEOUSTUBES AND TUBING REPAIRS1 GENERAL 2292 TUBE JOINTS 2293 FORMING FLARED TUBE JOINTS 2294 FORMING BEADED TUBE JOINTS 2295 GENERAL TUBING REPAIR 2306 REPAIR OF HIGH-PRESSURE TUBING ... 2307 LOW-PRESSURE LINE REPAIR 2318 CONDUIT REPAIR 2329 MISCELLANEOUS 23210 TOOLS 23211 MATERIALS 233CHERRY BLIND RIVETS12 GENERAL 23313 SELF-PLUGGING CHERRY RIVETS 23314 TYPES OF SELF-PLUGGING CHERRY RIVETS 23415 DETERMINING CHERRY RIVET GRIP LENGTH 23516 DRILLING PREPARATIONS 23517 DIMPLING PREPARATIONS 23518 SHEET CLAMPING 23519 CHERRY RIVET INSERTION 23520 TYPES OF CHERRY RIVET GUNS 23621 USE OF CHERRY RIVET GUNS 23622 CHERRY RIVET GUN MAINTENANCE .... 23723 TRIMMING CHERRY BLIND RIVETS .... 238DU PONT EXPLOSIVE RIVETS24 GENERAL 23825 RIVETING PRECAUTIONS 23826 STORAGE AND HANDLING PRECAUTIONS • . 238RESTRICTED[iv]

RESTRICTEDCONTENTSPar.PagePar.Page27 RIVET INSTALLATION > . . 23928 DU PONT RIVETING IRONS 239RIVNUTS60616263NEGLIGIBLE DAMAGESPLICING CONTROL RODSFABRICATING REPLACEMENT RODS . . ,REPLACING DAMAGED ROD END FITTINGS25125125225229 GENERAL 23930 TYPES OF RIVNUTS 23931 GRIP RANGE 23932 RIVNUT TYPE NUMBERS 24033 STRENGTH OF RIVNUTS 2403^ PREPARATIONS FOR INSTALLATION .... 24135 RIVNUT INSTALLATION 241ARMAMENT INSTALLATION REPAIRS64 GENERAL 253SECTION 10REBUSHING SPEC IF ICATIOHSDZUSFASTENERSGENERAL 25436 GENERAL 24237 REMOVING LIGHT-DUTY DZUS FASTENERS . 24238 REPLACING LIGHT-DUTY DZUS FASTENERS . 24239 REMOVING HEAVY-DUTY DZUS FASTENERS . 24440 INSERTING NEW FASTENER ....... 244TRANSPARENT PLASTICPANELS41 CLEANING 24442 MINOR SCRATCHES 24443 DEEP SCRATCHES 24444 REPAIRS - GENERAL 24545 ISOLATED CRACKS 24546 REPAIRING HOLES 24647 PREPARATION OF NEW CURVED PANELS . . 24648 ROUTING EDGES OF PANELS 24749 ROUTING INSIDE CURVES 24750 COCKPIT ROOF PANEL REPLACEMENT ... 24851 SIDE PANEL REPLACEMENT 248ALUMINUM CASTINGS AND FORGINGS52 GENERAL 24953 MINOR SURFACE CRACKS 24954 MINOR WELDING OF LOWLY STRESSED MEN^BERS 249PROPELLER55 Ml NOR REPAI RS 250WOODENACCESSORIES56 WOODEN COCKPIT SEATS 25057 WOODEN FLOOR BOARDS 25058 MINOR WOODEN ACCESSORIES 251CONTROLRODS59 GENERAL 251SECTION IIFINISH SPECIFICATION1 GENERAL REQUIREMENTS 2572 FINISH AND FINISH MATERIAL SPECIFI-CATIONS 2573 FI NISH CODE 258ALUMINUM AND ALUMINUM ALLOYS4 GENERAL FA-0 2585 INTERIOR CLOSED MEMBERS FA-20 .... 2586 INTERIOR SURFACES, PARTS, AND OPENMEMBERS FA-21 2587 ENGINE COMPARTMENT SURFACES FA-23 . • 2588 EXTERIOR SURFACES, PARTS, AND MEMBERSFA-25 2589 INSTRUMENT PANELS FA-28 25810 ELECTRICAL AND RADIO JUNCTION BOXESAND CONDIUT FA-29 25811 GENERAL FC-G 26012 COPPER LINES FC-20 26013 DISSIMILAR METALS FC-23 260GENERALFINISHES14 CHROMIUM PLATE FG-0 26015 DISSIMILAR METALS FG-5 26016 STEPS FG-7 26017 WALKWAYS FG-8 26018 COIL SPRINGS FG-10 26019 FLAT SPRINGS FQ-ll 26120 ARMAMENT PROTECTION FG" 13 2-^121 MARKINGS: LINES FG-21 26122 MARKINGS: CABLES, CRANKS, AND HORNSFG-22 26123 MARKINGS: MISCELLANEOUS FG-23 ... 26124 MARKINGS: ENGINE DISCONNECT POINTSFG-24 261[v]RESTRICTED

ILLUSTRATIONSRESTRICTEDPar. Page Par. PageMAGNESIUM AND MAGNESIUM ALLOYS TEXTILES25 GENERAL FM-0 26326 EXTERIOR SURFACES, PARTS, AND MEMBERSFM-20 26327 ENGINE COMPARTMENT SURFACES, PARTS,AND OPEN MEMBERS FM-23 26328 INTERIOR SURFACES, PARTS, AND MEMBERSFM-25 263PLASTICS29 INTERIOR SURFACES, PARTS, AND MEMBERSFP-21 26330 EXTERIOR SURFACES, PARTS, AND MEMBERSFP-23 26331 INTERIOR OF CLOSED MEMBERS FP-24. . . 263STEEL32 GENERAL FS-0 26333 EXTERIOR SURFACES, PARTS, AND OPENMEMBERS THAT CAN BE PLATED FS-20. . 26334 EXTERIOR CLOSED MEMBERS THAT CAN BEPLATED FS-21 26335 EXTERIOR CLOSED MEMBERS THAT CANNOTBE PLATED FS-22 26436 ENGINE COMPARTMENT SURFACES FS-23 • • 26437 INTERIOR SURFACES, PARTS, AND OPENMEMBERS THAT CAN BE PLATED FS-26. • 26438 INTERIOR CLOSED MEMBERS THAT CANNOTBE PLATED FS-27 26439 INTERIOR CLOSED MEMBERS THAT CAN BEPLATED FS-28 26440 TANK PADS AND STRAPS FT-20 26441 ELEVATORS AND AILERONS FT-21RUDDER FT-23 26442 ANTENNA MAST FT-24 26443 COTTON WEBBING AND FABRIC FT-25 ... 264WOOD44 GENERAL REQUIREMENTS 26545 TYPES OF WOOD SURFACES 26546 DETAIL REQUIREMENTS 26547 APPLICATION OF SEALER 265I II I I48 APPLICATION OF FILLER 26549 APPLICATION OF SURFACER 26550 APPLICATION OF FINISH COATS TO TYPESURFACES 26551 APPLICATION OF FtNISH COATS TO TYPESURFACES 26552 FINISHING EXPOSED END-GRAIN SURFACESNOT SUITABLE FOR TAPfNG 26653 FINISHING EXPOSED END-GRAIN SURFACESSUITABLE FOR TAPING 26654 SANDING S.EALER, FILLER, AND SURFACERCOATS 266SECTION 12HEAT-TREATED PARTS1 GENERAL 2672 MATERIAL CODE 2673 HEAT-TREATED PARTS LIST 268LIST OFILLUSTRATIONSFig.Page Fig. PageSECTION IGENERAL INSTRUCTIONS12345678910111213TYPES OF RIVETSGENERAL AIRPLANE DATAAIRPLANE THREE-VIEW DIMENSIONS.NAA PART NUMBERSTYPICAL CENTER-PUNCH MARKS- • .USING RIGHT-ANGLED DRILL. . . .RIVET COUNTERSINKINGCUT COUNTERSINKINGDIMPLE COUNTERSINKINGSKIN FASTENERSCLECO COLOR CODINGTYPICAL RIVET SETSRIVET BUCKING BARS

RESTRICTEDLLUSTRATI0N3Fi^'Page^i^'Page29A EXTRUSION EQUIVALENT CHART (CONTD.) .... 2130 STRESS CONCENTRATIONS 2231 PRIMING ALUMINUM SHEET 2232 STANDARD TOOL KIT 2333 METAL SHRINKER AND STRETCHER 2334 HAND ROLLER 2435 HAND BRAKE 2U36 BENDING SHEET STOCK LOCALLY 2437 ARBOR PRESS 2538 FORMING BLOCK 2539 TRIMMING MACHINE 2640 USE OF SPIRAL REAMER 2641 STANDARD POWER AND HAND TOOLS 2942 STANDARD METAL-WORKING HAND TOOLS 3I91011121314151617181920212223242526272829303132SECTIONFUSELAGEFUSELAGE STATIONS . . .34ENGINE MOUNT & FRONT FUSELAGE TUBE SIZES .35ENGINE MOUNT TRUSS ASSEMBLY 36FRONT FUSELAGE TRUSS ASSEMBLY 37UPPER TUBE CLUSTER JOINT AT STA. 55 ... .38ARC WELDING GENERATOR WITHCRATER ELIMINATOR 38ARC WELDING ELECTRODE REQUIREMENTS .... 39TYPICAL WELD ENDS 39TYPICAL ARC-WELDED ENGINE MOUNT JOINT ... 39OXYACETYLENE WELDING EQUIPMENT 40TYPICAL WELDING FAILURES 4OCORRECTING OVAL-SHAPED STEELTUBE DISTORTION 41STRAIGHTENING BOWED STEEL TUBES 42APPLICATION OF FINISH TO STEEL TUBES ... 43IIREINFORCING A DENT AT A STEELTUBE CLUSTER JOINT 43REINFORCING A DENT OR CRACK INSTEEL TUBE LENGTH 44STEEL TUBE INNER SLEEVE SPLICE 45CENTERING INNER SLEEVE IN STEEL TUBE ... 48STEEL TUBE OUTER SLEEVE SPLICE 49STEEL TUBE FISHMOUTH SPLICE USINGLARGER DIAMETER REPLACEMENT TUBE .... 50STEEL TUBE DIAGONAL SPLICE USINGLARGER DIAMETER REPLACEMENT TUBE .... 51EXHAUST MANIFOLD ASSEMBLY 51EXHAUST MANIFOLD ASSEMBLY EXPLODED .... 52EXHAUST MANIFOLD SLIP JOINT 52ENGINE FIREWALL ASSEMBLY 53FIREWALL WEB REPAIR 54FIREWALL WIDE-FLANGED STIFFENER SPLICE . . 54FIREWALL NARROW-FLANGED STIFFENER SPLICE . 55ENGINE COWLING CONSTRUCTION 55ENGINE COWLING ASSEMBLY 56FRONT FUSELAGE SIDE PANELS 57FRONT FUSELAGE SIDE PANEL FORWARDFORMER SPLICE 5733 ALUMINUM REAR FUSELAGE COVERED ASSEMBLY . . 5834 ALUMINUM REAR FUSELAGE PART NUMBERS .... 5935 SPLICE OF REAR FOUR FORMERS OFFRONT FUSELAGE SIDE PANELS 6036 ATTACHMENT OF FRONT TRUSS TOALUMINUM REAR FUSELAGE 6037 ALUMINUM REAR FUSELAGE INTERIOR 6138 ALUMINUM REAR FUSELAGE INTERIOR EXPLODED . 6139 ALUMINUM REAR FUSELAGE FORMER SPLICE ... 6240 FABRICATING REPLACEMENT FORMERS 6241 SPLICE FOR STRINGER TYPE C107LT-40 .... 6342 SPLICE FOR STRINGER TYPE C234LT 6443 ALUMINUM REAR FUSELAGE BAGGAGECOMPARTMENT INTERCOSTAL ANGLE SPLICE . • 6544 SPLICE FOR STRINGER TYPE C364LT 6645 ALUMINUM REAR FUSELAGE DOORS 6646 ALUMINUM REAR FUSELAGE UPPERLONGERON SPLICE 6747 ALUMINUM REAR FUSELAGE LOWERLONGERON SPLICE 6848 FUSELAGE SKIN ARRANGEMENT 6949 PLASTIC FAIRINGS AND FILLETS 7050 INTERIOR VIEW OF WOODEN FUSELAGE 7151 ALUMINUM ALLOY REINFORCEMENT OFWOODEN FITTINGS 7152 WOODEN REAR FUSELAGE FRAME ASSEMBLY .... 7253 INTERIOR VIEW OF FUSELAGE SHOWINGSKIN GRAIN DIRECTION 7354 ATTACHMENT OF LOWER STEEL LONGERONTO REAR WOODEN FUSELAGE 7355 MIXING CASEIN GLUE 7456 ATTACHMENT OF UPPER STEEL LONGERONTO REAR WOODEN FUSELAGE 7457 APPLYING CASEIN GLUE WITH BRUSH 7558 CORRECT EQUALIZED APPLICATIONOF GLUING PRESSURE 7559 INCORRECT CONCENTRATED APPLICATIONOF GLUING PRESSURE 7660 USING NAILING STRIPS TO APPLYGLUING PRESSURE TO SKIN 7761 TYPES OF CLAMPS FOR WOOD 7762 WOODEN REAR FUSELAGE STRINGER SPLICE ... 7863 REMOVING PORTION OF DAMAGED WOODSTRINGER ADJACENT TO SCARF CUT 7864 BAGGAGE COMPARTMENT WOODENINTERCOSTAL SPLICE 79^5 SMOOTHING SCARF CUT WITH A WOOD PLANE ... 7966 WOODEN REAR FUSELAGE UPPER LONGERON.SPLICE 8067 CHECKING SCARF CUT WITH STRAIGHTEDGE ... 8068 CUTTING WOOD SCARF JOINT 8169 WOODEN REAR FUSELAGE SKIN ARRANGEMENT ... 8170 WOODEN REAR FUSELAGE LOWER LONGERONSPLICE 8271 FLUSH PATCH FOR PLYWOOD SKIN HOLESLESS THAN 2 INCHES WIDE 8372 EXTERNAL PATCH FOR PLYWOOD SKINHOLES LESS THAN 2 INCHES WIDE 84RESTRICTED[vii ]

ILLUSTRATIONSRESTRICTEDFig. Page Fig. Page73 SCARFING PLYWOOD SKIN WITH SPOKESHAVE ... 857ii PLYWOOD SKIN SECTION REPLACEMENT 8675 WOODEN REAR FUSELAGE FORMER SPLICE .... 8776 TYPES OF SAWS 8777 TYPES OF FILES FOR WOOD 8878 TYPES OF PLANES FOR WOOD 8879 FINISH REQUIREMENTS FOR WOOD 89SECTION 3FIXED SURFACES1 CENTERSECTION RIB AND SPAR PART NUMBERS . • 922 OUTER WING RIB AND SPAR PART NUMBERS . . » 933 CENTERSECTION FUEL TANK COMPARTMENT DOORS • 94U ALUMINUM HORIZONTAL STABILIZER STRUCTURE • 955 WOODEN HORIZONTAL STABILIZER STRUCTURE • • 966 REAR SPAR JOINT OF WOODEN HORIZONTALSTABILIZER 977 WOODEN HORIZONTAL STABILIZER FRONTSPAR FITTING 978 VERTICAL STABILIZER RIB AND SPARPART NUMBERS 989 AT-6A WING STRINGER ARRANGEMENTAND STATIONS 9910 AT-6B AND AT-6C WING STRINGERARRANGEMENT AND STATIONS 10011 AT-6B AND AT-6C RIGHT OUTER WINGUPPER SURFACE STRINGER DETAILS 10112 AT-6B AND AT-6C RIGHT OUTER WINGLOWER SURFACE STRINGER DETAILS 10113 AT-6A STABILIZER STRINGER ARRANGEMENT ... 102lU AT-6B AND AT-6C STABILIZERSTRINGER ARRANGEMENT 10215 SUMMATION OF STRINGER TYPESUSED IN FIXED SURFACES 10316 SPLICE FOR STRINGER TYPE C107LT-20 .... lOU17 SPLICE FOR STRINGER TYPE C123LT ...... 10418 SPLICE FOR STRINGER TYPE C148T 10519 SPLICE FOR DOUBLED C148T TYPE STRINGERS . . 10520 SPLICE FOR STRINGER TYPE C180T 10621 SPLICE FOR STRINGER TYPE C204T 10622 SPLICE FOR DOUBLED C204T TYPE STRINGERS . . 10723 SPLICE FOR STRINGER TYPE C250T 10724 SPLICE FOR DOUBLED C250T TYPE STRINGERS . • 10825 SPLICE FOR COMBINATION C250T-C366TSTRINGERS 10826 SPLICE FOR STRINGER TYPE C265T 10827 SPLICE FOR STRINGER TYPE C266T 10928 SPLICE FOR STRINGER TYPE C274T HO29 SPLICE FOR STRINGER TYPE C366T HO30 SPLICE FOR DOUBLED C366T TYPE STRINGERS . • 11131 SPLICE FOR STRINGER TYPE C373LT 11232 SPLICE FOR STRINGER TYPE K77A 11233 SPLICE FOR DOUBLED K77A TYPE STRINGERS • . 11334 SPLICE FOR WING TRAILING EDGE 11335 OUTER WING MAIN SPAR 11436 OUTER WING MAIN SPAR SPLICE-ROOT TO 106 INCHES OUTBOARD 11537 OUTER WING MAIN SPAR SPLICE"106 TO 151 INCHES OUTBOARD OF ROOT ... 11738 OUTER WING MAIN SPAR SPLICE-151 INCHES OUTBOARD TO TIP 11839 OUTER WING FLAP SPAR SPLICE 12040 OUTER WING AILERON SPAR SPLICE 12141 CENTERSECTION FLAP SPAR SPLICE 12342 CENTERSECTION REAR SPAR 12443 CENTERSECTION REAR SPAR SPLICE 12544 CENTERSECTION FRONT SPAR 12645 CENTERSECTION FRONT SPAR UPPER CAP SPLICE 12746 CENTERSECTION FRONT SPAR LOWER CAP SPLICE 12847 CENTERSECTION FRONT SPAR WEB REPAIR ... 12948 VERTICAL STABILIZER FRONT SPAR SPLICE . . 13149 VERTICAL STABILIZER REAR SPAR SPLICE FORAREA ABOVE ROOT RIB 13350 VERTICAL STABILIZER REAR SPAR SPLICEFOR AREA BELOW ROOT RIB 13^51 HORIZONTAL STABILIZER FRONT SPAR SPLICE . 13552 HORIZONTAL STABILIZER REAR SPAR SPLICEFOR AREA OUTBOARD OF CENTER HINGEFITTING 13653 WING SKIN ARRANGEMENT 13754 STABhLIZER SKIN ARRANGEMENT 13855 SKIN REPAIR FOR HOLES LESS THAN1/2-INCH DIAMETER 13956 USING HOLE SAW TO TRIM SKIN DAMAGE .... 13957 PATCH FOR SKIN HOLES l/2-TO 1-INCHDIAMETER 14058 PATCH FOR 1-INCH DIAMETER SKIN HOLESNEAR ADJACENT STRUCTURE 14159 ARRESTING GROWTH OF CRACK 14160 TYPICAL FLUSH PATCH FOR LARGE SKIN HOLES . 14261 USING SPIRAL REAMER TO TRIM SKIN DAMAGE • 14262 TYPICAL EXTERNAL PATCH FOR LARGESKIN HOLES 14363 TEMPORARILY SECURING SKIN PATCHPRIOR TO RIVETING 14364 LOCATING BLIND RIVET HOLES BYIMPROVISED METHOD 14365 DRILLING OUT RIVETS COVERED BY SKIN PATCH 14366 BLIND RIVET HOLE LOCATING TOOL 14467 USING BLIND RIVET HOLE LOCATING TOOL ... 14468 BLIND RIVET HOLE LOCATING TOOLWITH PILOT BUSHING 14469 TYPICAL SKIN SPLICE 14470 WING LEADING EDGE SKIN SPLICE 14571 OUTER WING LEADING EDGE CONTOURS 14672 GAINING ACCESS TO THE OUTER WINGLEADING EDGE 14773 USE OF DRIFT PUNCH WHEN RIVETINGCLOSING STRIP SKIN 14774 USE OF DRILL AT CLOSING STRIP SKIN .... 14775 WELDING A ROD TO A DENT IN THEHORIZONTAL STABILIZER TIP 148RESTRICTED[ V i i i]

NRESTRICTEDLLUSTRATIONSFig. Page Fig. Page76 REMOVING DENT IN HOR. STAB. TIP 14877 FUEL TANK COMP. COVER FORMER SPLICE ... 1U978 CENTERSECTION INTERMEDIATE CENTER RIB . . 14979 CENTERSECTION END RIB 15080 CENTERSECTION INTERMEDIATE RIB SPLICE . . 15181 TRAILING EDGE RIB SPLICE 15182 VIEW OF OUTER WING SECTION J5283 TYPICAL RIBS 15284 TYPICAL REPAIR OF BROKEN RIB BEADS .... 15385 TYPICAL REPAIR OF BROKEN RIB CUTOUTS ... 15386 TYPICAL REPAIR OF BUCKLED RIB WEBS .... 15487 TYPICAL RIB FLANGE REPAIR WHERE FLANGEIS RIVETED TO SKIN 15488 TYPICAL RIB FLANGE REPAIR WHERE FLANGEIS NOT RIVETED TO SK I15589 TYPICAL RIB SPLICE 15590 CUTTING REPLACEMENT RIB FORM BLOCKS ... 15691 FORMING REPLACEMENT RIB 15692 CENTERSECTION BOLTING ANGLE PART NUMBERS . 15693 OUTER WING BOLTING ANGLE PART NUMBERS . . 15694 INSTALLING LOWER PORTION OF WING JOINTCOVER,15795 INSTALLING UPPER PORTION OF WING JOINTCOVER 15796 LOCATING BRACKET HOLE IN WING JOINTCOVER 15797 ASSEMBLING WING JOINT COVER 157SECTION 1MOVABLE SURFACES1 LANDING FLAPS INSTALLED ON WING 1612 LANDING FLAP PART NUMBERS 1623 RUDDER FRAME PART NUMBERS 1634 RUDDER INSTALLED ON VERTICAL STABILIZER . 1645 ELEVATOR FRAME PART NUMBERS 1656 ELEVATORS INSTALLED ON HOR. STAB 1667 AILERON FRAME PART NUMBERS 1678 LEFT AILERON INSTALLED ON WING 1689 REPAIR OF ONE-INCH DIAMETER HOLE IN SKINUNDER FLAP CHANNEL SPAR 16810 FLAP CHANNEL SPAR SPLICE 16911 FLAP LEADING EDGE SPLICE 17012 RUDDER AND ELEVATOR RIB REPLACEMENT ... 17013 FLAP TRAILING EDGE SPLICE 17114 AILERON SPAR SPLICE 17215 SEWING SMALL FABRIC TEAR 17316 CUTTING OUT DAMAGED FABRIC 17317 PATCHING V-SHAPED FABRIC TEAR 17418 DAMAGE REQUIRING FABRIC SECTIONREPLACEMENT 17519 DAMAGED FABRIC SECTION REMOVED 17620 FINAL FOLDING, PINNING, AND SEWING OFFABRIC SECTION 17621 DETAIL OF FABRIC SECTION STITCHING .... 17722 RUBBING DOWN REINFORCING TAPE AT SEAMS . . 17723 APPLYING FINISHING TAPE ON FABRICSECTION SEAMS 17824 TYPICAL REMOVAL OF FABRIC FROM TWOOR MORE SECTIONS 17825 REMOVING FABRIC RETAINING SCREWS .... 17926 PULLING ON PARTIAL FABRIC ENVELOPE COVER 17927 RUDDER FABRIC COVERING REQUIREMENTS ... 18028 ELEVATOR FABRIC COVERING REQUIREMENTS . . 13129 AILERON FABRIC COVERING REQUIREMENTS . . 18230 FABRIC DOUBLE-STITCHING MACHINE 18331 PULLING NEW FABRIC COVER OVER CONTROLSURFACE STRUCTURE 18332 PINNING UNSEWED END OF FABRIC ENVELOPE . 18433 TRIMMING THE UNSEWED END OF FABRICENVELOPE AFTER PINNING 1843^ FABRIC PINNED AT TRIM TAB CUTOUT .... 18535 CUTTING FABRIC AT TRIM TAB CUTOUT .... 18536 ATTACHING FABRIC ALONG TRIM TAB CHANNEL . 18637 CUTTING TAPE WITH HAND PINKING MACHINE . 18638 ATTACHING FABRIC TO RIBS 18739 HAND SEWING AT TRIM TAB CUTOUT 187W USE OF PNEUMATIC SCREWDRIVER TO ATTACHFABRIC TO RIBS 18841 HAND SEWING FABRIC AT HINGE CUTOUT ... 18842 APPLYING BRUSH COAT OF DOPE TO FABRIC . . 18843 HAND STITCHING THE UNSEWED END OF FABRICENVELOPE 18844 APPLYING FINISHING TAPE TO FABRIC .... 18945 FABRIC ENVELOPE PRIOR TO FIRST ANDSECOND COATS OF DOPE 18946 APPLYING SPRAY COATS OF ALUMlNlZED DOPETO FABRIC 18947 APPLYING BALANCE PATCH TO AILERON .... 18948 DETERMINING CONTROL SURFACE UNBALANCE . • 19049 RUDDER UNBALANCE LIMITS 19150 AILERON UNBALANCE LIMITS 19251 ELEVATOR UNBALANCE LIMITS 19252 APPLYING BALANCE PATCH TO ELEVATOR ANDRUDDER 193SECTION 5FUEL AND OIL SYSTEM1 FUEL TANK SUPPORTS 1942 FUEL TANK CONSTRUCTION 1943 OIL TANK CONSTRUCTION 1954 OIL TANK ASSEMBLY 1955 OIL TANK AND OIL COOLER INSTALLATION . . 1966 FUEL TANK TESTING JIG 1977 SEALING SMALL CRACKS 1988 WELDING SPLIT SEAMS 1989 HYDRAULIC OIL TANK 19910 FORMING AND WELDING INSERT 20011 COMPLETED FUEL AND OIL TANK REPAIRS ... 20112 OIL COOLER ASSEMBLY 20213 OIL COOLER TUBE REPLACEMENT 204[ix]RESTRICTED

ILLUSTRATIONSRESTRICTEDFig. Page Fig, PageSECTION 6LANDING GEAR1 MAIN LANDING GEAR SHOCK STRUT 2062 MAIN LANDING GEAR 2063 TAIL WHEEL ASSEMBLY 2074 TAIL WHEEL COMPONENTS 2075 HYDRAULIC JACK CRADLE 2076 REMOVING MAIN LANDING GEAR WHEEL 2087 REMOVING TAIL WHEEL 2098 BALANCE WHEEL 2109 APPLYING CEMENT 21010 APPLYING BALANCING DOUGH 21111 APPLYING TIRE TALC TO BALANCING DOUGH . . 2ll12 APPLYING TIRE TALC TO INFLATED TUBE ... 21213 VALVE EXTENSION 21214 TIRE AND TUBE MARKINGS 21215 SPECIAL TIRE TOOLS 21216 MOUNTING MAIN GEAR TIRE AND TUBE 21317 MOUNTING TAIL WHEEL TIRE AND TUBE .... 21418 APPLYING TIRE TALC TO TAIL WHEEL TUBE . . 215SECTION 7CONTROL CABLES1 TYPES OF CABLES 2172 CABLE DATA 2183 PREPARATION OF A WOVEN CABLE SPLICE ... 2194 PREPARATION OF A WOVEN CABLE SPLICE ... 2195 TYPES OF CABLE TERMINALS 2206 CABLE CLAMP FOR WOVEN SPLICE 2207 WRAPPED SOLDERED SPLICE 221SECTION 8ELECTRICAL SYSTEM1 TYPICAL WIRE ROUTING 2232 SLIPPING INSULATION OVER WIRE 2243 CABLE TERMINAL SWAGING 2184 CRIMPING THE TERMINAL 2255 SLIPPING INSULATION INTO POSITION .... 2256 TYPICAL SWITCH BOX WIRING 2267 TYPES OF TERMINALS 2278 CONDUIT BONDING METHODS 2279 ELECTRICAL REPAIR TOOLS 22810 ELECTRICAL REPAIR MATERIALS 228SECTION 9MISCELLANEOUS1 TUBE FLARING TOOLS 2292 BEADING SMALL TUBING 2303 BEADING LARGE TUBING 2304 HIGH-PRESSURE TUBING REPAIR OF MAJORDAMAGE 2315 HIGH-PRESSURE TUBING REPAIR OF MINORDAMAGE 2316 LOW-PRESSURE TUBING REPAIR 2327 CONDUIT REPAIR > 2328 SPECIAL WRENCHES FOR HIGH-PRESSURECONNECTIONS 2339 CHERRY RIVET GRIP LENGTH 23410 INSERTING CHERRY RIVETS 23611 EXPANDING CHERRY RIVETS 23612 UNTRIMMED CHERRY RIVETS 23713 TRIMMING CHERRY RIVETS 23714 CHERRY RIVET HAND GUN 23715 EXPLOSIVE RIVET CROSS SECTION 23816 USE OF EXPANDING IRON 23817 USE OF KEYSEATING TOOL 24018 INSERTING RIVNUT 24019 SQUEEZING RIVNUT 24120 REMOVING MANDREL 24121 TYPICAL DAMAGE TO DZUS FASTENERS 24222 DZUS FASTENER TOOL 24223 LIGHT-DUTY DZUS FASTENER REPLACEMENT ... 24324 REMOVING GROMMET 24425 INSERTING GROMMET AND FASTENERS 24426 REMOVING DEEP SCRATCHES IN TRANSPARENTPLASTIC PANELS 24527 APPLYING HOT-AIR BLAST TO CRACKEDPLASTIC PANEL 24528 APPLYING TORCH FLAME TO A CRACKEDTRANSPARENT PLASTIC PANEL 24629 PATCHING HOLES IN PLASTIC PANELS 24630 MODIFIED DRILLS FOR TRANSPARENT PLASTICPANELS 24631 PANEL ROUTING JIG 24732 ROUTING INSIDE CURVES 24733 CURVE ROUTING DETAIL 24834 COCKPIT ENCLOSURE ASSEMBLY 24835 PROPELLER INSTALLED 25036 WOODEN COCKPIT SEAT 25137 CONTROL ROD REPAIR COMPONENTS 25138 CONTROL ROD SPLICE 25239 CONTROL ROD FITTING REPLACEMENT 25240 ARMAMENT INSTALLATIONS 253SECTION 11FINISH SPECIFICATION1 ENGINE DISCONNECT POINTS 2592 LINE COLOR CODING CHART 2603 AIRPLANE INSIGNIA LOCATIONS 262RESTRICTEDI - ]

EESTRICTEDINTRODUCTIONTO THE REPAIRMAN:In the preparation of this Manual, an attempt has been made to include extensive and complete informationfor the repair of damaged structural and nonstructural parts of the airplane. It is intended thatthis Repair Manual will facilitate the restoration of the airplane to active service within a minimumof elapsed time and with a minimum of replacement parts. If spare or replacement parts are required,they ma^y be ordered as out I ined below.The repairs outlined in this Manual provide for the repair of any daniaye to any part of the aircraftand are not based upon any preconceived idea as to the extent of damage. Since the scope of possibledamage is so broad it would be impossible to outline any exact procedure, only general repa i methodsare described and illustrated. These methods may have to be supplemented and altered to suit each specificsituation. The material requirements, rivet specifications, etc., are the minimum al lowable; andfor the repair of any important structural components, personnel are warned against the use of any methodof repair not described herein.A complete explanation of each repair is outlined in the text. The illustrations should not be construedas being self-explanatory. The scheme of arrangement of each section is similar, deal Tng firstwith the description, then with repair procedures, materials, and tools required for making ttie variousrepairs. A thumb index has been inserted at the outside top of each page for quick identification ofthe section title and paragraph number.For convenience, all common U. S. material specifications together with the British equivalents ofthese material specifications are listed at the end of Section I.Your cooperation in reporting your views on this N-lanual and the material contained herein will helpsubstantial ly in the further development of a more adequate manual. Opinions on the matter are sol icitedand may be communicated to the company through the North American Service Representative at yourfield or sent directly to the Field Service Department at Inglewood, California.Field Sen/ ice DepartmentNorth American Aviation, Inc.SPAREPARTSIn many instances, incorrect part numbers, lack of complete infometion, and lack of part identificationmay cause delay in the filling of orders for spare parts. Requests by the company- for additionalinformation, necessitated tiy these discrepancies, involve loss of time; and the need for such can beeliminated if the correct part number, name, and location are clearly set forth in the original request.To facilitate the prompt fulfillment of requests, observe the following rules:(I) Give correct part number and full title. (Refer to the applicable structures diagram in thisManual or to the Parts Catalog, T.O. No. 01-60FE-4JExample: 54-14015 left nose rib, station 15.087 (left wing, AT-6B)88-I40I5-I right nose rib, station 15.087 (right wing, AT-6C)Note the prefix change for the AT-6C and that the part number ends with -I signifying a right-hand part.[xi]RESmiCTED

I beAN,NTRODUCTIONrestricted(2) Give the serial number and iiDdei of the airplane for which the part is intended.Example: Serial No. AC4-I-322I8, ^todel AT-6C (This infom^tion is stenciled in the upper left-handcorner on the left-hand front fuselage sTde panel and stamped on a name plate on the righthandside of the front cockpit. )(3) If the part number or name is not available, furnish the location and complete description of thepart and the unit or system in which the part wi I used.Example: The second nose rib outboard of the bolting angle of the left-hand wing assembly, AT-6C,Serial No. AC4I-322I8.(4) Give conplete dash numbers and title of al IAC, B, C, and NAF Standard Parts.Example: AN3-4A bolt, AC36^832 nut, BI289- 1(^10 screw, C384->5 neoprene extrusion, NAF314LH cable terminal.1062 -S2 1[>-RESTRICTED[xii]

RESTRICTEDFRONTISPIECE.^^•nrD-^Three Viewsof Complete Airplane[ xiii ]RESTRICTED

specifiesumIRESTRICTEDGENERALPAR. I TO^SECTIONGENERAL INSTRUCTIONSII. AIRPLANE CONSTRUCTION.The AT-6 Series Airplane is a single-engine,low-wing, two-place advanced trainer. Generalairplane data are tabulated in this Section.(See Figures 2 and ^.) The airplane is equippedwith a Pratt and Whitney, nine-cylinder, radial,air-cooled engine on which is mounted a HamiltonStandard, controllable pitch, constant speed,two-blade propeller. The main landing gear isfully retractable and the nonretractab le tailwheel is full-pivoting. With the exception ofthe chrome molybdenum steel tubing front fuselagetruss and engine mount, the airplane is primarilyconstructed of 24ST ale lad sheet and 24ST a I u-minum alloy extruded members. The front andrear fuselage sections, the outer wing panels,and the empennage are constructed to facilitateremoval and replacement in the event of dafnage.All structural members are readily accessibleand are fabricated in a manner as to providefor ease of repair. It is to be noted thatmultiplicity of types and sizes of componentparts is confined to an absolute minimum inoriginal construction of the airplane and alsoin repair. North American Aviation, Inc., majorcomponent part numbers are listed and illustratedin this Section. (See Figure 4.)rivet" signifies a flat head rivet of A|7S material.The last number of the rivet notationalways signifies the rivet diameter in 1/52-inch. Therefore, the rivet diameter in thiscase is 5/32-inch.FLATE^SSJ D^DIA.100*.^c'sunk'"ft -BRAZIERAN442-AD BI2I9-AD BI227-ADDIA. 3/32(AN426-AD)APPROXIMATEDIMENSIONS2. RIVET TYPES.There are three different types of rivets usedin the construction of the airplane (see Figure1) . All a re made of A 1 7S a I i num a I Ioy material(rivet designation AD), and thus no heattreatment is required prior to driving. Rivetsof AITS material are marked with an identifyingdent or depression on the center of the manufacturedrivet head. These three types of rivetsemployed in the construction of the airplaneare as fol lows: ( I ) the Type AN442-AD flat headrivet, used for the assembly of a I interiorstructure; (2) the Type BI2I9-AD (supersededby Type AN426-AD for all future procurement)l(DO-degree countersunk head rivet, used to reducedrag on the forward skin of the fuselage andfixed surfaces; (3) the Type BI227-AD brazierhead rivet, used on the aft skin of the fixedsurfaces and fuselage. In making repairs, thesetypes of rivets should be used in their originalfunction. The rivet coding used throughoutthis ^.'lanua I the type of rivet andthe size. For example, a notation "AN442-AD5

GENERALFIGURE 2RESTRICTEDTABLE OF SPECIFICATIONSOVERALL W ING SPANOVERALL LENGTHOVERALL HE IGHT ANTEN NA MAST, THRUST LINE LEVELOVERALL HE IGHT ANTEN NA MAST, THREE-POINTHE IGHT, PROPELLER HU B, THRUST LINE LEVELCLEARANCE, PROPELLER T IPS, THRUST LINE LEVELDIAMETER OF PROPELLE RNORMAL C.G. TO WING LEADING EDGE - WHEELS DOWNNORMAL C.G. TO W I NG LEAD ING EDGE - WHEELS UPC.G. LIMITS TO W ING LEADING EDGEELEVATOR H I NGE TO Wl NG LEADING EDGEMINIMUM CLEARANCE OF PROPELLER AND R ING COWLANGLE OF NORMAL C.G. WITH WHEEL (Side Elevation)ANGLE OF REARWARD C. G. WITH WHEEL (Side Elevation)ANGLE OF WING TIP AN D WHEEL BOTTOM (Front Elevation)ANGLE OF THRUST LINE WITH GROUND (3-Point Position)42

RESmiCTEDGENERALFIGURE 3t STA. I tELEVATOR HINGE ^1 RUDDER HINGEFigure 3— Airplane Three-view Dimensions[3]RESTRICTED

GENERALFIGURE ^RESTRICTED{«T-6») («T.6«) (»T-SO(«T-St) (tr-e» (1T-«C)77 an M 00002AltrLME ASSUBLYCIMEKAL77 77 88 3II0IFKAIIE ASSEUBLY. fVSELACE COHrtElE77 >« 88 IWOI55 9H 88 I "400955 IW)M78 l'tM28>t'. 66 80 ISOOl55 1600555 180066 1800277 2000155 2000255 2000355 2000155 2000555 2000855 20O0977 8* 8» 2100166 2200155 220 1855 SU 8«....2»0I77 2W)0I55 211036mSlALLAlIOH - BinesFIUET, KKG TO FUSanOt, FWIITFILLET, milO TO FUSELtGE, mTERHEOI >TE tUtFILLH tSSEMBLY, MING TO FUSELAGE, I1E»IIMING «SSE>l6Lr, CENTER SECTION COHFLETECOVERED JSSEMBLY, OUTER MINGWGLE INSTtLLtTlON, OUTER MING BOLTINGTIP tSSL-IBLT, MINGDOOR ISSEMBLT, OUTER MING LW01N3 LIGHTIILERON tSSLISLTT»B «SSEHBLr, ilLERON BOOSTERFLU> tSSEMBLT, OUTER MINGFLAP ASSEMBLY, MING CENTER SECTIONIHSIALLAIION. ItFEHNACEFILLET ASSEMBLY, HORIZONTAL STABILIZER LEFT FRONTFILLH ASSEMBLY, VERTICAL STABILIZER TO FUSELAGEFILLET, HORIZONTAL STABILIZER, INTERMEOl ATE LOMERFILLET, HORIZONTAL STABILIZER, LEFT REARFILLET, HORIZONTAL STABILIZER, »1 GHT REARFILLET ASSEMBLY, HORIZONTAL STABILIZER, RIGHT FRONTSTABILIZER ASSEMBLY, HORIZONTALELEVATOR ASSEMBLY, COVERECTAB ASSEMBLY, ELEVATOR TRIMSTABILIZER ASSEMBLY, VERTICALRUDDER ASSEMBLY, COVEREDTAB ASSEMBLY, RUDDER TRIM77 78 88 3110477 3110577 3110655 3123855 78 78 3121177 77 88 3180155 55 88....3I80277 77 88 3182655 55 88 3182719 19 88 318 2855 55 88 3182977 31901n 3300177 78 78 3310256 33»555 55 88 3100158 58 88 3100277 78 88 5100177 78 88 5100277 78 88 51003FIREMALL ASSEMBLY, FUSELAGEFRAME ASSEMBLY, FUSELAGE FRONTFRAME ASSEMBLY, FUSELAGE REAR SECTION COVEREDCOVER ASSEMBLY, FUSELAGE REAR SECTION BODY ACCESSCOVER ASSEMBLY, FUSELAGE REAR SECTION FLEXIBLE GUN TROUGHIKSIALLATIOn, COCKFIT EHCLOSVSEHOOD ASSEMBLY,GUNNER'SMINDSHIELD ASSEMBLY, COCKPIT ENCLOSUREPAfEL ASSEMBLY, COCKPIT ENCLOSURE, FRONTPANEL ASSEMBLY, COCKPIT ENCLOSURE, FIXEDPANa ASSEMBLY, COCKPIT ENCLOSURE, REARMOUNT ASSEMBLY,ENGINEinSTALLAH OK, LAUDING OEARINS7ALLATI0N , TAIL WHEELPOST AND HOUSING ASSEMBLY, TAIL MHEELIHSIALLAIION. INSIKVUENIS77 77 88 31001FVSELAGl ASSEtBLr COVEKED55 55 88.. ..53001INSTALLATION, FUKMISHINCS77 77 88 31002-1777 77 88 31002-1877 77 S8 31002- 1977 78 78 3101077 78 78 3101377 77 88 3102777 77 88 3102877 77 88 3103077 3103277 310 3377 3103H77 3103577 310»277 3105077 77 88. ...3105155 3106177 3106577 3106677 3107577 3107677 31077COMLING ASSEMBLY, ENGINE RING, LOMERCOMLING ASSEMBLY, ENGINE RING, RIGHT HAND UPPERCOMLING ASSEiHBLY, EKGINE RING, LEFT HAND UPPERCOMLING ASSEMBLY, FUSELAGE FLEXIBLE GUN CUTOUTCOMLING ASSEMBLY, FUSELAGE GUNNER'S HOODCOMLING ASSEMBLY, ENGINE REMOVABLE, CARBURETOR AIR S(COMLING ASSEMBLY, ENGINE REMOVABLE, UPPER RIGHTCOMLING ASSEMBLY, ENGINE REMOVABLE, BOTTOMFAIRING ASSEMBLY, FUSELAGE FIREMALL UPPER LEFT SIDECOVER ASSEMBLY, FUSELAGE MAIN FUSE BOX ASSEMBLYFAIRING ASSEMBLY, FUSELAGE FIREMALL LOMER LEFT SIDEFAIRING ASSEMBLY, FUSELAGE FIREMALL UPPER RIGHT SIDEFAIRING ASSEMBLY, FUSELAGE FIREMALL LOMER RIGHT SIDEPANEL ASSEMBLY, FUSELAGE LEFT SIDE FAIRINGPANEL ASSEMBLY, FUSELAGE RIGHT SIDE FAIRINGCOMLING ASSEMBLY, FUSELAGE TAllMHEEL CUTOUTCOMLING ASSEMBLY, ENGINE REMOVABLE TOPCOMLING ASSEMBLY, ENGINE REMOVABLE LOMERDOOR ASSEMBLY, FUSELAGE FIREMALL TO MINDSHIELO LEFTDOOR ASSEMBLY, FUSELAGE BAGGAGE COMPARTMENTDOOR ASSEMBLY, FUSELAGE FIREMALL TO MINOSHIELO RIGHT55 55 88 5301055 55 88. ...53 10055 7310155 55 88 73106INSTALLATION , FHOTOCRAFH IC EQVIFMEMTDOOR ASSEMBLY, CAMERA^^^'

I fromI siRESTRICTEDGENERALPAR. H TO 6should be center punched and the actual drillingdone with a light power drill. The centerpunchmark should be large enough to preventthe dri I si ipping out of position, yetmust not be made with enough force to dent thesurrounding material (see Figure ^}. The drillingcan be done with a hand drill if no powertool is available, but considerably more timewill be taken. Also, as the drill speed isslower when the work is done by hand, there is atendency to apply more pressure; consequently,a large burr is often formed as the dri I I piercesthe material. All burrs must be removed beforeriveting by utilizing a metal countersink, or5. STANDARD HOLES FOR RIVETS.The following chart specifies the size ofhole to drill for the application of the varioussizes of rivets. See the following paragraphfor dri I zes.RIVETFigure S-'Typical Center-punch Marksby filing or scraping. In some cases, the useof a right-angled drill is particularly advantageousin drilling through restricted access,small fittings, clips, etc. (see Figure 6).

GENERALPAR. 7 TORESTRICTED7. RIVET COUNTERSINKING DIMENSIONS.The following dimensions should be adheredto when cut countersinking or dimpling sheetmaterial for the application of BI2I9-AD (orthe AN426-AD) lOO-degree countersunk rivets.Use the proper degree and diameter countersinkand cut just deep enough so the rivet head andthe metal will forma flush surface. (See Figure7JRIVET SIZE

RESTRICTEDGENERALPAR. 10 TO 12drilled through, several varieties of skin fastenersare availaole to line up holes. (SeeFigure 10.) Cleco skin fasteners are commonlyr-SKINFASTENERpRlVETGUN>^' 5SFigure 10—Skin Fastenersused throughout the aircraft industry in productionand repair work. Used in conjunctionwith a Cleco safety gun (see Figure 41), Clecoshold two pieces of metal firmly together untilriveting is accomplished. Inasmuch as Clecosare used only through drilled holes, it is importantthat the proper size oe used. The varioussized Clecos are easily identified by their colorcoding. A Cleco for a 3/32-inch diameter holeis cadmium colored, copper colored for a 1/8-inch hole, Diack for a 5/32-inch hole, and brasscolored for a 3/16-inch hole. (See Figure 11.)11. BUCKING BARS.Sucking oars may assume any numoer of givenshapes in order to facilitate rivet forming ininaccessible areas. (See Figure 13.) The rivetgun pressure applied to the manufactured headof the rivet forms the rivet shank against theDucking oar. In similarity to rivet sets, rivetDucking Dars must oe kept clean, smooth, andwell polished at a I I times.Figure 11 —Cleco Color CodingliHBRAZIER HEAD TYPE4*AA^FLUSHTYPE12. RIVET SETS.A rivet set is a punch-like tool with a flatFigure 12— Typical Rivet Sets[7]RESTRICTED

)GENERALPAR. 12 TO 16RESTRICTEDliJX i U^TJ600-I TJ600-2 TJ600-5JLJTJ600-6 TJ 600-10 TJ600-28aTJ 60 0-32 TJ600-42 TJ600-60m. HAMMERS.Hartniers sometimes may have to be used to driverivets, but they should be used only in emergencieswhere other equipment cannot be obtained.The use of hammers to drive rivets for the repairof primary structure is not recommended.On minor secondary structure, however, ballpeenhammers may sonietimes be employed.15. RIVET SQUEEZER.A rivet squeezer consists of a large pair offorcers which are used over the edge of theTRIGGERV]i f»TJ600-64 TJ600-65 TJ600-86J^TJ600-90 TJ600-99 TJ600-106Figure 13~-'Rivet Bucking Barsor recessed die for driving flat, countersunk,and brazier head rivets, respectively. Theymay assume a variety of shapes to facilitaterivet driving in restricted areas (see Figure1^). Rivet sets should be kept smooth and wellpel ished at al I times.13. PNEUMATIC RIVET GUN.The pneumatic rivet gun used with a rivet setis applied to the manufactured head of the rivet;and the pressure, applied in a series ofvibrations, forms the rivet shank against abucking bar. Different types are available tomeet various requirements. (See Figure 14'\The period of time during which pressure isapplied to fonn the head must be short enoughto prevent the metal from strain hardening nearthe end of the rivet. The pneumatic equipmentmay be used either with a flat or recessed rivetset, depending upon the type rivet being driven."Because of the uniformity of driving rivetspneumatically, all riveting should require thisequipment. Both the rivet gun and the buckingbar are separate units requiring separate operators.(See Figure lo.Figure 14-^ Pneumatic Rivet Gunsmaterial to head the rivet. Pneumatic rivetsqueezers may be used with alligator, offsetalligator, and C-type jaws. Inasmuch as arivet squeezer can be used only over the edgeof the sheet where conditions permit, its applicationis necessarily limited. (See Figure 25.J16. REMOVAL OF SOLID RIVETS.Rivet removal should be accomplished by drilling,not shearing. Use drill one size smallerthan diameter of rivet, and place point intodepression in rivet head. If power tool isused, locate drill first; then apply power inshort bursts. Drill must be applied carefullyand small cuts taken in a series of applications.Drill until rivet head twists from shank. (SeeFigure 16-) Drift remainder of rivet free.RESTRICTED[8]

)RESTRICTEDGENERALPAR. 16 TO 18PNEUMATICSQUEEZERDISTANCE "E" SHOULD EQUALTWICE THE RIVET DIAMETERh'f[W'MINCORRECT -TOOCLOSE TO EDGESECT. AA^CORRECTE»2D1RESULTANT CRACKSAFEFigure15—Rivet SqueezerFigure 17 — Rivet Edge DistanceIt is recommended that the drill be equippedwith a stop of some soft nonabrasive materialto prevent the drill chuck from plunging intoand marring the surrounding metal.17. RIVET AND BOLT EDGE DISTANCES.The necessity for providing adequate edgedistances from the rivet or bolt to the edgeof the affected material is of extreme importancein repair. Generally, this distance shouldnot be less than two times the diameter of therivet or bolt, measured from the centerlineof the rivet or bolt hole to the edge of thematerial involved. (See Figure ly-DRILL UNTILRIVET HEADTWISTS FROMSHANK, THENPUNCH OUTRIVETFigure 16 '—Removal of Solid Rivets18. RIVETING.Inasmuch as riveting is the major means ofsecuring airplane parts, the quality of workmanshipemployed in driving rivets must be high.The structural efficiency of the completed productis greatly increased where power-driventools are available. For this reason, the repairof primary structures should demand pneumaticallyoperated rivet-driving tools. However,for the repair of secondary structures, manuallyoperated rivet-squeezing tools will sufficewhere conditions permit. The riveting of minorparts sometimes may be accomplished by meansof a simple rivet set and bounce ball-peenhamrrer or the equivalent. Examples of correctlyand incorrectly driven rivets are illustratedon the following diagram. (See Figure i8.)The following comments are made with referenceto th is figure.a. Rivet (A) is driven correctly. The properlength has been determined and the rivet hasbeen driven slowly and compressed evenly. Theupset end of an und riven rivet of proper lengthshould measure approximately 1-1/2 times thediameter of the rivet from the surface of thematerial being riveted.b. Rivet (B) is driven incorrectly. The headof the rivet has been cut by insecure and/orimproperly held driving tool. Rivets damagedin this manner should be replaced. Damage ofthis nature may be prevented by holding thetool securely and as nearly in alignment withthe rivet shank as possiole.[9]RESTRICTED

I equalGENERALPARAGRAPH 18RESTRICTEDc. Rivet (C) is driven incorrectly. The rivethas been driven to excess and /or too much pressurehas been applied to the bucking bar. Asa result, the head has flattened excessivelyand cracks are attendant. In the majority ofsuch cases, the cracks extend well into theshank of the rivet. Such conditions are acceptablecause for failure of the rivet uponapplication of load. Rivets damaged in thismanner must oe replaced. In the thinner sheetmaterial, the rivet holes become enlarged tothe extent that the sheet must be replaced.d. Rivet (D) is driven incorrectly. Thiscondition is the result of failure to securethe sheets or parts together prior to riveting,thus causing creeping or distortion of the sheet.In this case, the shank of the rivet has swol I eninto the area produced oy separation of thesheets. The rivet must be replaced. Dri M therivet free oy means of a drill considerablysmaller than its shank diameter; then followwith a dri I in diameter to the estimated(A) (B) (C)diameter of the swollen portion. Special careshould be exercised to prevent such occurrence,as it is most difficult to remove drill chipsfrom between the sheets. Upon reriveting, asmooth surface cannot be obtained. Such conditionsgreatly lower the value of compressiveload resistance of the joint. Various skinclamps can be used in connection with rivetingto prevent this occurrence. (See Figure lo.)e. Rivet (E) is driven incorrectly. An unbalancedamount of head material is locatedabout the centerl ine of the rivet and, as such,the efficiency of the rivet is impaired. Thiscondition is a result of any one or a combinationof three things: the rivet bucking barhas not been held securely; it has been forcedto one side oy pressing too hard without adequatesupport; or it has been permitted to bounce andslide carelessly over the rivet. This is amost common error and one which should receivemuch cond iderat ion. The rivet should be replaced.(D) (£) (F)DRIVEN UNSTEADY DRIVEN SEPARATION UNSTEADY EXCESSIVECORRECTLY RIVET SET EXCESSIVELY OF SHEETS TOOL SHANK LENGTHBOTTOM VIEWFigure18^ Rivet Driving PracticesRESTRICTEDC 10]

RESTRICTEDGENERALPAR. 18 TO 19f. Rivet (F) is driven incorrectly. The rivethas failed laterally. This condition is a resultof improper selection of rivet length,combined with any one or all of the faults ofrivet (E). Sufficient care must be exercisedin proper selection of rivet length. A rivettoo short in length does not permit the requiredamount of head material to be upset. Too muchlength results in a buckled shank. Driving arivet of excessive length also presents possibilitiesof hole enlargement.19. FITTING OF BOLTS AND SCREWS.All bolt holes should be the smallest consistentwith design, and all bolts must be suitablylocked. Always use the first drill sizelarger than the nominal diameter of the bolt.When in doubt, it is advisable to drill thehole undersize and ream for a close fit.. Boltholes must not be oversized or elongated. Abolt in an oversized or an elongated hole takesnone of the applied shear load until the metalit is holding in shear has slipped enough, andthe remaining bolts sheared enough, to allowthe bearing surface to come in contact withthe bolt. (See Figure ig. ) This shearingaction is a definite failure in these boltsand they are no longer capable of carryingtheir design loads. In some cases of oversizedor elongated holes, it is permissibleto drill or reani the hole until large enoughto permit the insertion of the next size largerbolt. However, before resorting to this method,obtain permission from an authorized engineeringofficer. When in doubt, it is advisable toreplace the fitting, or fittings, which containsthe oversized hole with one in which theholes are the proper size. To make a bolt installationstructurally sound, bolt threadsmust never bear in any part of a fitting. Allbolt holes must be normal to the surface involvedand must provide complete bearing surfaceto the head and shank of the bolt. Also, anybolt entering an anchor nut must be free enoughto engage the threads in the nut by hand manipulation.Although bolt threads must cut completelythrough the fiber of self- locking nuts,there should be no more than two bolt threadsshowing beyond the fiber. To meet this requirement,washers may be used under nuts, and shimsmay be added under nut plates. The bolts mustbe sufficiently tight, but too much twist onthe wrench must also be avoided. The inchpoundstwist required should be within a certainrange for the bolt diameter used. The followingchart specifies the twist required. Inch-poundsis obtained by multiplying the pull by the wrenchlength or by the use of a toraue wrench whichis calibrated in inch-pounds (see Figure 20).-'^'-rN^BOLT SIZE TWIST (INCH-POUNDS)AN3A NUAN5AN6AN7AN8(1032(1/U(5/16(3/8(7/16(1/22 0-2550-75100-lUO1 60-190u 50-500U80-690Figure 19—Bolted Joint With Oversized Hole Figure 20 — Bolt Torque Measur ementLii]RESTRICTED

RESTRICTEDGENERALFIGURE 22THREAD 1/4- 28NF 3« SPEC. AN-GG6-SI26.4375 —ioioMARK FORBRONZEDRILL *48 (.076)CHAMFER 45* X gyOR ROUND END OP-TIONAL- END RAD. TMAXIMUM 2IMPERFECT TH'DS.CHAMFER ON BOTTOM FACEIS*' TO DIM. 7/16 OPTIONALDASHNOS.

GENERALFIGURE 23RESTRICTEDHEAD ANGLE 8 DIM. B'ARE RELATED DIM'S.DIM. V TO BE USEDAS REFERENCE ONLYSTAMP "X"INDENTEDON STEELONLYioo*t rTOP FLAT UNSHAVEDSHAVE CONCENTRICWITH SHANK WITHINTOTAL RUNOUT OF .005THREAD CLASS 3MAXIMUM 2IMPERFECT TH'DSCHAMFER1/32 X 45*ROLLED OR CUT THREADS OPTIONALSIZE

RESTRICTEDGENERALFIGURE 2^MEASURED TOSHARP CORNERS.015 R. MAX. FOR 5/16 a 3/8 SIZES.OIOR. MAX. FOR * 8, ** 10 8 1/4THREAD CLASS 3MAXIMUM 2IMPERFECT TH'DSSTAMP "X"INDENTED ONSTEEL ONLYROLLED OR CUT THREADS OPTIONALCHAMFER1/32 X 45**SIZE

GENERALFIGURE 25RESTRICTEDDIMENSIONSSIZE aTHREAD

loysloyseRESTRICTED24. ALUMINUM WROUGHT STOCK - GENERAL.The symbol S is used after the formula numberof wrought aI to distinguish them from thecast alloys. In the case of heat- treatab 1aluminum, the symbol fol lowing the S indicatesan annealed material, but the symbol T followingthe S indicates a heat-treated, hardened material.Alloy 24ST in the form of sheet and extrudedshapes is used almost entirely in theconstruction of the aircraft; but because ofbetter forming qualities, parts are often fabricatedof 24S0 and subsequently heat treated.Alloys I4ST and I7ST are used for forgings only.Material of I4ST, possessing high mechanicalproperties, is most prevalent in forgings.Material of I7ST finds use where an increasein corrosion resistance is required, with buta slight sacrifice in strength. See Section12 for the required strength of forgings. Thesymbol W appl ies only to a requiring artificialaging and is used principally with theIalloy 53SW to signify a heat-treated, quenched,but not completely age- hardened material. Itis to be noted that the use of the symbol T todesignate a hard alloy is applicable only tothe heat-treatable alloys, principally 14ST,I7ST, and 24ST. However, of the wrought alloys,theri are a few which are nonheat-treatable,principally 2S, 3S, and 52S. These nonheattreatablealloys are also nonstructural; butbecause of their malleability, they find particularuse in fairing and filleting. Coldworking hardens these al loys to a certain extentand this condition is noted not by the symbolT but by the symbol H. Intermediate conditionsof hardness are noted as I/4H, I/2H, etc. ThefoMoA/ing chart gives general particulars pertinentto wrought alloy uses.ALLOYFORMUSE14ST17STFORGINGSFITTINGS2USTSHEET,ALCLADSKIN, SPARS, LONGERONSRIBS, FRAMES, FLOORS,FUSELAGE STRINGERS24STEXTRUDEDSHAPESSPAR CAPS,STRINGERS,WINGSTIFFENERS53 STEXTRUDEDSHAPESENCLOSURE & WINDOWFRAMESONLY53SWTUBEPLUMBING,LINES2SSHEETFAIR.ING-,FILLETING

etheGENERALPAR. 29 TO 311RESTRICTEDoy vi Drat ion, and does not require annealingafter prolonged service.30. ALUMINUM SHEET, 52S-I/2H.This alloy possesses the same chemical andmechanical properties as those of 52S0; Dutin similarity to the 2S-I/2H and 3S- I /2H alloys,it is hardened to an intermediate temper oetweenthe soft and hard oy cold working. The materialfinds general use in fillets and fairing whereintermediate values of strength, hardness, andductility are required.31. ANNEALING ALUMINUM, 2S, 33, AND 523.The nonheat-t reataD I alloys, 2S, 3S, and52S, are sufficiently malleaole for cold workingin their original state. If these materialshave already oeen strain hardened oy cold working,annealing is instantaneous upon reachingthe proper temperature, and the material shouldoe profdptly removed from the furnace and allowedto cool at normal room temperature. Hardnessis accomplished only oy cold working. Thefollowing chart specifies annealing temperaturepart icu lars.ALLOY2S52SANNEALING TEMPERATURE329-371385-U13329-37132. ALUMINUM EXTRUSIONS, 2US0.625-700725-775625-700This material has values of corrosion resistanceconsideraoly inferior to 2S or 5S, Dutit possesses excellent working and forming qualities.The material is heat-treatable, Dut itdoes not work harden. Heat treatment oringsits strength and hardness to a level fully equalto that of 24ST material. All 24S0 materialmust be heat treated after forming.Figure 26— Enlarged Cross Sectionof Ale ladSheetFigure 27— Typical Extruded Sections33. ALUMINUM EXTRUSIONS, 2»IST.This material is the hardened form of 24S0aluminum alloy. Extruded shapes of 24ST areextensively employed as stringers, longerons,and spar caps. (See Figure 21.) The 24ST aluminumalloy composition is the strongest andhardest of al I a! leys used, but the corrosionresistance of the material is consideraDly inferiorto 24S alclad sheet. To protect thematerial against corrosion, zinc chromate primermust De applied to all surfaces.34. ALUMINUM SHEET, 2i+S0 ALCLAD.This material is identical to 24S0 aluminumalloy; Dut in order to provide greater protectionagainst corrosion, a thin layer of purealuminum 2S is bonded to each side of the sheet.This pure aluminum coating has a yery high resistanceto direct contact with corrosive agentsand also, Dy electrolytic action, it servesto protect the cut edges of the sheet. Thematerial is slightly inferior in strength to24S0 aluminum alloy, but it possesses excellentforming qualities. Parts requiring severe workingand bending should be formed of 24S0 alcladand then heat treated to the full hardness of24ST.RESTRICTED[18]

theedloysRESTRICTEDGENERALPAR. 35 TO 3735. ALUMINUM SHEET, 2^ST ALCLAD.SALT BATH METHODThe heat-treated form of 24S0 ale lad is knownas 24ST. This material is the most generallyused of alI aluminum a I and is uti I I zedfor ribs, frames, spars, longerons, stringers,skin, and formers. Although slightly inferiorto 24ST aluminum alloy in strength, 24ST ale ladpossesses far greater corrosion resistance. Oneach face of the high-strength 24ST core is anintegrally bonded layer of relatively high purityaluminum known as the coating. (See Figure26.) This coating protects the core in twoways. It protects the core elect rolyt ical lyand covers the core to prevent contact withcorrosive agents. The material should neverbe fully annea I .36. HEAT TREATING ALUMINUM, 24S0.The heat treating of 24S0 aluminum or 24S0alclad should be accomplished in either an airfurnace or salt bath furnace. Methods of treatmentin which the flame may come in direct contactwith the material are not recommended. Theheat-treating process consists essentially of:( I) heating the material to a prescribed temperature;(2) holding the material at this temperaturefor a specified length of time; (3)promptly quenching the material; and (4) allowingthe material to age harden at room temperaturefor a specified period. The actual heattreatingtemperature of 24S0 aluminum and 24S0alclad must remain constant within a range of4g\-4gq°C (9I5-930°F) for the time specifiedon the following charts. However, periodsspecified should be timed after the charge hasreached uniform heat-treating temperaturesthroughout the furnace. AJso, when heat treatingparts of varied thickness simultaneously,use the thickest sheet as a timing criterion.For example, when heat treating a sheet of .040- in.and a sheet of .064^ in. use a minimum heat-treatingtime of 30 minutes and maximum heat-treatingtime of 40 minutes.AIRFURNACE METHODMATERIAL

cadloy,ordinaryI onlyGENERALPAR. 37 TO ^0RESTRICTED(due to the quenching) should not cause rejectionif the part can De properly straightenedduring assembly or within one hour after quenching.In fact, if any rework is necessary, itis advisable to accomplish it always withinone hour after quenching before the aging hasprogressed too far. During this period, themetal may be worked with ease and without dangerof cracking. Aging may be retarded byplacing the material in a cold atmosphere.38. AGE HARDENING.After quenching, the strength of the 24Saluminum or 24S aid ad gradual ly increases atroom temperature until, after approximately24 hours aging, the material reaches a fullyhardened condition. When this occurs, thematerial is ready for installation.39. ANNEALING ALUMINUM, 24ST.When annealed 24S or 24S a materialsI Iare required in the course of repair, theyshould be obtained in the annealed condition/VAA STANDARD ALCOADIE NO.PRINCIPALDIMENSIONSfrom the metal manufacturer. However, if 24S0material is not available, it is sometimes permissibleto anneal 24ST aluminum al loy fullyand to anneal 24ST alclad partially. However,this is not recommended where the material isto be used in location of high stress. Of thetemperature range specified for anneal ing 24STaluminum a I the lower temperature (343°C,650°F) will anneal the material sufficientlyto permit removal of cofd-working strains asrequired for al I types of repair fabrication.The upper temperature (427°C, 800°F)specified below for 24ST aluminum al loy wi IIpermit maximum annealing; however, the materialmust be held at this temperature for two hours,then cooled to 260°C (500°F) at the rate ofIO°C (50°F) per hour. After the temperatureof aluminum alloy has reached 260°C (500°F),the cooling rate in air is not critical. Itis to be noted that the temperature range specifiedbelow for 24ST alclad wi I part iai lyanneal the material. Full annealing of 24STalclad is not permitted, because of the rapiddiffusion of the base metal through the aluminumcoating occurring as a result of slow cool ihgfrom high temperatures. The following chartspecifies annealing temperature particularsapplicable to air furnace and salt bath furnace.ALLOY ANNEALING TEMPERATUREC366TL249I024ST AL. ALLOY24ST ALCLAD40. BENDING 24S ALCLAD.343-427338-349650-800640-660C373LTNONE(ROLLED ALCLAD)K77A K77A -75.125 -1Exercise caution in handling alclad materialsto avoid scratching through the pure aluminumcoating. Besides destroying the protectivealuminum coating, scratches in alclad materialare potential cracks which may develop underthe application of stress. Where the materialmust be flanged, a power or hand brake will befound particularly useful. Alclad sheet maybe formed and bent as required, but the allowablebend radius for the thickness of sheet,as specified herein, should not be exceeded.The following chart specifies particulars pertinentto the bending of alclad.ALCLADALLOWABLE BEND RADIUSFigure 29^Extrusion Equivalent Chart24ST2US04 TIMES THE THICKNESS OF THESHEET OR GREATER.2 TIMES THE THICKNESS OF THESHEET OR GREATER.RESTRICTED[20]

RESTRICTEDGENERALFIGURE 29A/V4/i STANDARD ALCOADIE NO.PRINCIPALDIMENSIONSNAA STANDARDALCOADIE NOPRINCIPALDIMENSIONSCI07LT-20(ROLLED ALCLAD)j020C234LT(ROLLEDALCLAD)NONECI23LT(ROLLEDALCLAD)NONE.040.75-»-| .562 [-*C245T K 16862ONE -HALF SIZE |.687CI48TCI80TC203TC204TK734-JJC250TK 14280 .75C265T*-| .531 [-«-K 14653KI4654.078 ^K 16869C266T\i C274T.625.125*-

I edJGENERALPAR. m TO 46RESTRICTEDlil. EXTRUSION EQUIVALENT DIE NUMBERS.In order to expedite the procurement of theextrusions required in the repair of this seriesairplane. North American Aviation standard partnumoers and equivalent die dimensions and num-Ders of the Aluminum Company of America (Alcoa)are incorporated in this Manual (see Figure2q) . It is to De noted that all extruded merrv-Ders are made of 24ST aluminum alloy material.n. EXTENT OF DAMAGE.When estimating the extent of damage, thesurrounding structure must De careful ly examinedto ensure that no secondary damage remains unnoticed.Secondary damage may De produced insome structure remote from the location of theprimary damage Dy the transmission of the damagingconcentrated load over the normal coursetoward the Dasic reacting structure. Damageof this nature usual ly occurs where the mostaorupt change in load travel is experienced.If this damage remains undetected, loads appliedin the normal course of operation maycause failure of the part. It is to oe notedthat all loads travel toward the structure comprisingthe wing attachment to the fuselage.Carefully Investigate all structure across whichthe loads must travel to reach these points.Particularly examine joints for security ofrivets. Dolts, Dent plates, etc.45. CORROSION PROTECTION.Figure 30— Stress Concent rat ionsVhen repairing parts, adequate measures mustDe taken to prevent the suDsequent occurrenceof corrosion. See Section II for particulars.^2, SCRATCHES, DENTS, AND CRACKS.Many parts 'oecome scratched and dented in thecourse of service, thos producing stress concentrationswhich may cause failure of the part.Sharp nicks in the smooth surface are especial lydangerous, and care must oe taken at all timesto smooth out such damage. When this is donehigh concentrations of stress disappear. (SeeFigure ^0.) The importance of smoothing outsurface Irregularities cannot be overemphasizedbecause the working stresses are usually sohigh that any defect of this kind may producestresses aoove the strength of the part. Careshould oe taken to ensure that no cracks remainundetected. Cracks in the skin generally occurat points of cold working and when the skinis dri I or dimpled for flush-type rivets orDolts. To arrest the growth of the crack, drilla No. 40 hole at each end of the crack and cleanout any rough material around the crack.i|3. SUPPORT OF STRUCTURE DURING REPAIR.It is essential that the structure to De repairedis adequately supported against distortionswhen any memoer, or part of a memDer, isremoved for assemoly. For Instance, it is veryimportant that the wing De adequately supportedwhen removing a panel of wing skin.Figure Sl-^Priming Aluminum SheetOne coat of zinc chromate primer should Deapplied to all surfaces of extruded repairmemDers and to the overlapping surfaces of al I24ST alclad sheet repa I r memDers. Always stirzinc chromate primer thoroughly Defore use andapply an even coat with a soft Dristled Drush.(See Figure 91.i|6.TOOLS - GENERAL.Wherever possiDle, power tools should Deused for repairs; Dut where facilities areRESTRICTED[22]

.ever-operatedler)RESTRICTEDGENERALPAR. 46 TO 19limited, hand tools sometimes may be employed.Riveting tools are needed most frequently andthese have been described in preceding paragraphs.Occasions arise when repairs are impracticaland a replacement part must be usedin order to restore the component to normaloperating condition. Other occasions arisewhen an assembly will be made more accessibleand hence more conveniently repaired if removedfrom the airplane. For all such occasions,a standard tool kit containing anassortment of tools will be necessary (seeFigure 32Jjaws. To decrease curvature and to increasecircumference, use stretcher jaws. (See FigureSHRINKINGJAWS SHOWNFigure 33— Metal Shrinker and Stretcher48. HAND ROLLER.Figure 32— Standard Tool Kit^7. METAL SHR INKER OR STRETCHER.Replacement sections for curved fairing,fillets, and cowling may be readily fabricatedby ut i I izing a metal shrinker and/orstretcher. The foot- Imachinehas an upper and a lower set of movable jaws,which move vertically and horizontally, shrinkingor stretching the metal. These two operationsare accomplished by the use of two interchangeablesets of jaws and side plates. Tooperate this machine, place the sheet betweenthe jaws, and move the sheet along in successiveuniform movements while operating thefoot lever. To increase the sheet curvatureand to decrease circumference, use shrinkerThe hand roller used for rolling and beadingsheet stock consists of a handle and a setof gears, which in turn are attached throughshafts to rollers, interchangeable as to size,shape, and depth of grooves. Several effectscan be obtained with this tool, the most commonof which is the beading of sheet stock.A guide is provided for lining up the metaland a crank is employed to raise or lower theupper ro I for a deep or shal low bead asrequired. To operate, adjust the depth tosuit, place edge of material on face of guide,push sheet snugly between rollers, and turnactuating handle. It is not necessary to forceor pull the material through the rollers andonly a slight pressure should be required tokeep the sheet drawing through the roller.(See Figure ^4 .49. HAND BRAKE.Inasmuch as bent-up sheet stock is employed[23]RESTRICTED

GENERALPARAGRAPH^9RESTRICTEDextensively in the repair of longerons, spars,frames, and rios, the use of a hand Drake willoe found particularly desirable. By meansof a hand Drake, accurate pending of sheetstock may oe accomplished readily. Most Drakesaccommodate wide sections, Dut the length ofthese sections is limited oy the capacity ofthe hand Drake. The hand brake consists essentially of three parts; the Ded, the jaw,and the taole. The Ded is stationary and hasa level drop surface. The jaw can De movedup and down Dy hand levers and can be adjustedin the horizontal plane Dy means of adjustingscrews. The taDle is hinged to the Dedin such a manner that it is possible to movethe taDle through an arc of aDout 135 degrees.The horizontal adjustment of the jaw dependsupon the thickness of the material Deing Dentand the radius of the Dend desired. It isimportant that this adjustment De made properly,as otherwise the material oeing bentmay De seriously damaged. If it is desiredto make a true radius inside the Dend, theuse of scrap material placed under the materialoetween the clamping jaws will providegreater radius as required. Scratches shouldDe avoided by protecting the sheets with wrappingFigure 35— Hand Brakepaper. When ut i I izing th is method of oending,adhere to the radii outlined in a precedingparagraph. (See Figure 35.>*Figure 34~-Hand Roller Figure 36— Bending Sheet Stock LocallyRESTRICTED[24]

)letRESTRICTEDGENERALPAR. 50 TO 5250. HAND METHODS OF BENDING.Hand methods of bending are seldom as satisfactoryas machine methods; however, if thework is done carefully, satisfactory resultswill be obtained. An improvised bending formmay be made by clamping a steel bar or a hardwoodblock on each side of the material andforcing the extended flange down with a smoothboard. (See Figure •^6.) Where the materialis too thick to bend in this manner, bend thematerial down with a steady pressure and, atthe same time, strike the back of the boardwith a lead mallet. Narrow flanges are bestbent by clamping the flange between the metalbars or wooden blocks and bending the mainbody of the material. If a sharper bend isdesired, a lead mallet may be used to forcethe sheet against the block.EXTRUDEDSECTIONFORMBLOCKFigure 38— Forming Blocking a C-type cast structure with a geared thrustarm, the arbor press, used in conjunction withvarious types of forming blocks, gives controlledthrust and a wide area of pressure.The arbor press has a variety of uses, butthe bending of extruded sections is of primaryconcern. To set up and operate the machineproperly, center the wood or lead formingblock on the base, place the extrudedsection on the forming block, place a protectivewood block over the extrusions, andapply a desired pressure by means of the handlever. (See Figure 57. J52. FORMING BLOCKS.51. ARBOR PRESS.Figure 37—Arbor PressFORMBLOCKWhere curved extruded sections are neededin repair, the use of an arbor press will greatlyfacilitate forming operations. Employ-Forming blocks of desired size and shapemay be fabricated from hardwood or lead tofacilitate hand forming of extruded sectionsor bar stock. Used in conjunction with a leadmallet, such blocks assist in making variousbends. Center the aluminum stock on formingblock and strike with lead ma i in consecutiveblows until correct curvature is obtained.Although this method is not as satisfactoryas the machine methods, satisfactorybends may be made if care is taken. (SeeFigure

istinearGENERALPAR. 53 TO 55RESTRICTEDFigure 39— Trimming MachineFigure 40— Use of Spiral Reamer53. TRIMMING MACHINE.S^. SPIRAL REAMER.Occasionally, straight shears will not sufficeA spiral reamer used with a drill motor isfor special trimming or cutting operation. quite satisfactory as a I cutter for cleaningholes in alclad prior to patching. MarkIn such instances, the use of a trimming machine,made up of two cutting blades and a the out I i ne to be cut with a soft penc i I andlarge C-shaped casting, is quite helpful. The drill a 1/4-inch diameter hole in one corner.Insert the reamer into the hole, tiltshears may accommodate large pieces of sheetstock. Sheet material between the shear pieces it toward the operator, and cut on the line.is trimmed by actuating the lever handle. Successivecuts may be obtained by raising the Sharp corners result in concentrations of stressCare should be taken to radius all corners.lever handle, si iding material to new position, and ultimate cracking of the material. (Seeand again actuating lever handle. (See Figure 99,) Figure 40.)55. STANDARD POWER AND HAND TOOLS.The fol lowing Iis a summation of standardpower and hand tools that may be required inrepair. The reference numbers refer to thetools illustrated on Figure 41.REF.NO.TOOLUSE1FILEHOLDERUSED TO HOLD FILES IN CURVED POSITION.2KEY-SEAT! NGUSED TO KEY-SEAT DRILLED HOLES PRIOR TO IN-STALLATION OF RIVNUTS.SQUEEZER,RIVNUTUSED TO SQUEEZE RIVNUTS IN DRILLED AND KEY-SEATED HOLES FOR DE-ICER BOOT SCREWS.RESTRICTED[26]

RESTRICTEDGENERALPARAGRAPH bbREF.NO.TOOLUSE4PINPUSHERUSED TO PLACE DE-ICER PINS IN POSITION WHENINSTALLING DE-ICER BOOT.5BURRINGTOOLUSED TO SMOOTH TRIMMED EDGES OF METAL.6BURNISHINGTOOLUSED TO SMOOTH MINOR SCRATCHES IN METAL.7PLIERS,CRIMPINGUSED TO CRIMP TERMINAL LUGS ON ENDS OF ELECTRICWIRES AND CABLES. TWO LARGER SIZES AVAILABLE.IRON,SOLDERINGUSED IN ELECTRICAL REPAIR WORK FOR SOLDERINGTERMINALS AND CONNECTIONS, AND FOR GENERAL TIN-NING. VARIOUS SIZES, TIPS, AND WATTAGE AVAILABLE.BRUSH,WIREUSED TO CLEAN FILES, WELDS, AND RUSTY METAL ANDFOR GENERAL ROUGHING.10CLAMP,"C"USED FOR HOLDING TWO OR MORE PIECES OF METALUNTIL DRILLING AND/OR RIVETING IS ACCOMPLISHED.11PLIER,CLAMPING12SKIN FASTENER,CLECOINSERTED IN DRILLED HOLES TO HOLD TWO OR MOREPIECES OF METAL UNTIL RIVETING IS ACCOMPLISHED.13GUN, CLECOSAFETYUSED TO INSERT SKIN FASTENERS.14WRENCH,TAPUSED IN CONJUNCTION WITH TAPS, EZY-OUTS, HANDAND TAPERED REAMERS.15PUNCH, DRIVEPIN "DRIFTS"USED FOR DRIVING OUT PINS, RIVETS, AND BOLTS.USE NEXT SIZE SMALLER THAN HOLE.16CHISELUSED FOR METAL CUTTING, AND REMOVAL OF RIVETHEADS WHERE DRILLING IS IMPRACTICAL.17PUNCHES,CENTERUSED FOR CENTERING HOLES AND AS A PUNCH BEFOREDRILLING TO ENSURE AN ACCURATE DRILL START.TYPES SHOWN ARE BLUNT AND SLIM TAPER.1819FASTENERS, SKINCLAMPS, SKINUSED TO SECURE METAL WHEN RIVETING, DRILLING, ETC.20SAWS,HOLEUSED WITH PILOT IN HAND AND POWER DRILLS, THESETOOLS ARE EXCELLENT FOR CUTTING CIRCULAR HOLES INDAMAGED SHEET METAL PRIOR TO APPLICATION OFSTANDARD BUTTON PATCH. SIZES RANGE FROM 5/8TO 2 INCHES.21PILOT, HOLE SAWUSED TO PI LOT HOLE SAWS.22WHEELS,EMERYSMOOTH AND ENLARGE INSIDE EDGES OF LARGE HOLES.SIZES SHOWN 1-1/2 AND 2 INCHES-[27]RESTRICTED

IGENERALPARAGRAPH 55RESTRICTEDREF.NO.TOOLUSE23DRILL,TWISTTHIS TYPE DRILL USED WITH HAND OR POWER DRILLSIS STANDARD. SIZES RANGE FROM NO. 80 TO 2NCHES.2425REAMER, HAND, SQUARESHANK, SPIRAL FLUTEDREAMER, SQUARE SHANK,TAPER PIN, SPIRALFLUTEDUSED FOR ENLARGING DRILLED HOLES TO DESIREDSIZE. VARIOUS SIZES AVAILABLE.USED FOR ENLARGING DRILLED HOLES INSHEETSTOCKAND FOR REAMING SHAFT HOLES FOR TAPER PINS.IN ELECTRIC AND AIR MOTORS, REAMERS ARE EXCEL-LENT FOR MAKING CUTOUTS IN ALUMINUM MATERIAL.26EXTRACTOR,EZY-OUTUSED FOR REMOVING THE REMAINING PORTIONS OFBROKEN SCREWS, BOLTS, AND PINS. DRILL HOLE INOBJECT APPROX IMATELY, ONE-HALF THE DIAMETER ,INSERT EZY-OUT, AND WITHDRAW OBJECT BY COUNTER-CLOCKWISE ROTATION.27TAPUSED FOR THREADING VARIOUS SIZES OF HOLES FORALL STANDARD THREADS. AVAILABLE IN DIFFERENTDIAMETERS, THREADS PER INCH, AND PITCH.28COUNTERSINKUSED FOR PREPARING HOLES IN HEAVY STOCK FOR USEOF COUNTERSUNK RIVETS AND SCREWS. AVAILABLEFROM 1/8-TO 2 INCHES. DEGREE OF TAPER FORRIVETS AND SCREWS, 78°, 82°, AND 100°.29COUNTERSINK,BACKUSED FOR COUNTERSINKING IN LOCATIONS WHEREMOTIVE POWER CANNOT BE APPLIED ON SIDE DESIRED.30SPOT-FACER,BACKUSED WITH VARIOUS SIZED PILOTS WHERE ACCURACYOF HOLE LOCATION IS DESIRED. ALSO USED FORFLUSHING BOLT HEADS IN CASTINGS. TYPE SHOWNIS USED IN LOCATIONS WHERE MOTIVE POWER CANNOTBE APPLIED ON SIDE DESIRED, BUT STRAIGHT SHANKTYPE ALSO IS AVAILABLE.31 FILES, HANDGENERAL: USED FOR FINISHING WORK ON ALL TYPESOF METAL, PLASTIC, AND FIBER, AND ENLARGINGHOLES, BURRING EDGES, SMOOTHING CUTS, REMOVINGMETAL, TOOL SHARPENING, ETC. MANY DIFFERENTLENGTHS, CUTS, AND SECTIONS ARE OBTAINABLE.VIXEN: A VERY COARSE FILE USED ON ROUGH WORKREQUIRING REMOVAL OF LARGE QUANTITIES OF MA-TERIAL.BASTARD: A COARSE FILE AVAILABLE IN VARIOUSLENGTHS AND SECTIONS.MILL: A FLAT FiNE-CUT FILE FOR FINAL FINISH-ING AND TOOL SHARPEN ING.SMOOTH: SIMILAR TO MILL, BUT OBTAINABLE INVAR lOUS SECT IONS.Swiss needle: a fine, slim, tapered fileUSED ON delicate WORK AND IN RESTRICTED AREAS.RESTRICTED[28]

RESTRICTEDGENERALFIGURE m2 3 4 5 6812131015 16 17189 ®20 21 22I23 24 25 26 27 28iin29 30Figure 41 ~-^ Standard Power and Hand Tools[29]RESTRICTED

GENERALPARAGRAPH 56EESTRICIED56. STANDARD METAL-WORKING HAND TOOLS. hand tools that may be required in repair.The reference numbers refer to the tools ilThe following list is a brief summation of lustrated on Figure 42.REF.NO.TOOLUSE1SNIPS,TINNERSUSED FOR STRAIGHT LINE CUTTING OF SHEET METALUP TO .064" THICK.SNIPS, COMBINA-TION CIRCLE"DUCK BILL"SNIPS, LEFT-HAND,DOUBLE ACTION"DUTCHMANS"SNIPS, RIGHT-HAND,DOUBLE ACTION"DUTCHMANS"HACK SAW, LARGE,ADJUSTABLEUSED FOR STRAIGHT LINE AND CURVED CUTTINGUSED FOR CUTTING TO LEFT AND/OR CUTTING ENDSOF TUBING.USED FOR CUTTING TO RIGHT AND/OR CUTTING ENDSOF TUB ING.USED FOR METAL, FIBER, AND PLASTIC CUTTING.BLADES VARY IN LENGTH AND IN NUMBER OF TEETHPER INCH.HACK SAW, SMALL USED FOR LIGHT CUTTING OF METAL, FIBER, ORPLAST I C.HACK SAW, KEYHOLEV ISE, DRILL PRESS,SPEEDUSED FOR CUTTING HOLES IN RESTRICTED PLACES.USED TO HOLD OBJECTS WHILE DRILLING.DRILL,HANDUSED FOR RESTRICTED OR SLOW DRILLING, AND ALSOIN THE PRESENCE OF GAS FUMES, IF AN AIR MOTORIS NOT AVAILABLE.10111213141516DIV IDERSSCRIBEMALLET, LEADMALLET, RUBBERMALLET, RAWHIDEMALLET, PARALYNMALLET, PARALYNUSED FOR SHEET METAL LAYOUT.USED TO FORM AND BEND METAL.17HAMMER,FINISHINGUSED TO FORM AND FINISH METAL.181920RULE, SQUARE,CENTER AND PRO-TRACTOR COMBINATIONRULE, 6" RIGIDPROTRACTOR, FLATUSED FOR SHEET METAL LAYOUT.21CALIPERS,SLIDEACCURATE INSIDE OR OUTSIDE DIAMETER MEASUREMENTIS OBTAI NED WITH TH IS TOOL.22RULE, 6* FLEXIBLEUSED FOR SHEET METAL LAYOUT.RESTRICTED[30]

RESTRICTEDGENERALFIGURE 1^2Figure 42— Standard Metal-working Hand Tools[31]RESTRICTED

GENERALPAR. 57 TO 58RESTRICTED57. ALUMINUM ALLOY MATERIAL SPECIFICATION EQUIVALENTS.

RESTRICTEDGENERALPAR. 59 TO 6059. COPPER ALLOY MATERIAL SPECIFICATION EQUIVALENTS.

heightto.RESTRICTEDFUSELAGEPAR. I TO2SECTION 2I. FUSELAGE GENERAL CONSTRUCTION.The complete fuselage is comprised of threeseparate detachable sections consisting of theengine mount, the front fuselage section, andthe rear fuselage section. The entire fuselagehas a maximum c ross-sect i ona I of 76- J /2inches and a maximum width of 46-13/16 inches.The detachable engine mount and the front fuselagestructure from the firewal I a positionaft of the rear cockpit seat are of chrome molybdenumsteel construction with welded steeljoints and fittings. On the earl ier AT-6C SeriesAirplanes and previous North AmericanAviation, Inc., trai ner ai rcraft, the fuselagesection from the rear cockpit to the tail isof riveted 245T ale lad aluminum semimonocoqueconstruction. On the later AT-6C Series Airplanes,the rear fuselage section is constructedof wood and the entire structure is assembledwith casein glue. The removable 24ST ale ladaluminum alloy side panels of the front fuselagesection serve only as fairings and do notFUSELAGEcarry loads. The tenn "fuselage station" usedin this Section and on NAA production drawingsspecifies the location of tubing cluster jointsand formers measured in inches aft of the firewall(see Figure i). Also, on production drawings,tubing cluster joints of the front fuselageare numbered from the firewall consecutivelyin reference with a U or L to designate anupper or a lower tubing cluster joint (seeFigure i). The complete fuselage is a-ttachedto the upper surface of the centersect ion bybolts passing through the welded longeronf i tt ings.2. CONSTRUCTION OF ENGINE MOUNT AND FRONTFUSELAGE TRUSS.The engine mount and the front fuselage trussextending from the firewall to a point justaft of the near cockpit are constructed of varioussizes of chrome molybdenum steel tubingconforming to SAE Spec. X4I30 (see Figure 2)The steel tube truss-type engine mount is®FIREWALL -I49.7578.375J^31.265 106.5 242.187514.51069687(JOINING SURFACE)Figure 1 — Fuselage Statior^s[34]RESTRICTED

FUSELAGEFIGURE 2RESTRICTED5 2UPPERTRUSSnote:all tubes arecm. steel exceptAS NOTEDSIDETRUSS—WALL THICKNESS056I^^M .065iI— .049I I .083.05824STLOWERTRUSSTUBE DIAMETERI.- 1-1/2 O.D. 5.- 1 O.D.2.-I-3/8 0.D.3.- I- 1/4 O.D.6-7/8 CD.4.- I- 1/8 CD. 7.-3/4 O.D.Figure 2 — Engine Mountand Front Fuselage Tube SizesRESTRICTED[35]

ectdaffectedRESTRICTEDFUSELAGEPAR. 2 TO «l3. WELDING STEEL PARTS - GENERAL.REF. DWG. 77-31901Figure 3—Engine Mount Truss Assemblybolted to the two upper and the two lower longeronsof the front fuselage tubing truss bysingle-bolt fittings. The engine is attachedto the engine mounting ring i n n i ne places byrubber-cushioned, dynamic mount fittings (seeFigure gj. Inasmuch as the cross bracing in thelower fuselage truss is discontinuous betweenthe centersect ion attachment points (see Figure4), this fuselage truss structure is rigidonly when attached to the centersect ion, andcare should be taken to support this tubularframework properly when it is removed fromthe centersect ion. This type of truss is inaccordance with conventional practice^ Thefirewall is not considered a structural memberalthough it actually gives some lateralstability to the front diagonal tube. It isto be noted that no welding is done on anytubes between their extremities. This providesample space on the tubes for clampingor otherwise attaching the equipment accessoriesand .other conponents. Reinforcing platesare welded on seme joint clusters (see Figure ^).Electric arc welding is employed throughoutthe entire structure in the original construction,but oxyacetylene welding may be substitutedin making repairs if arc welding fac -i Iities are not available. All tube sizes andwall thicknesses are noted on the truss assemblydiagram (see Figure 2).I damagedThe process of joining steel parts by weldingconsists in fusing the metal of tl'ie weldingwire or rod with the metal of the Joint endsuntil the joint is built up with new metal.This process can be accomplished by eitherelectric arc welding or by oxyacetylene welding.The fused metal of the joint is of acast structure and does not have the physicalproperties or strength of the metal parts orthe welding wire before being fused. The sectionsof the tube adjoining a gas weld willbe annealed by the welding heat for a distanceof from l/4-to 5/4-inch on each side of theweld. In cases of electric welding, this distancewi I I be greatly reduced, as the weld proceedsmuch faster than in the case of gas welding,and no preheating of the material is necessary.The melting point of steel is approxirmtelyI37I°C (2500°F) and the welder's torchflame is near this point at all times. A combinationeye and face shield and a leatherapron should be worn while welding, and properventilation should be provided. Inasmuch asthe presence of gas vapors will be a fire hazardall necessary fire precautions must be takenbefore any welding is attempted on a Iships that have been in service. Sandpaperor use a wire brush and clean al I areasprior to welding if a completely new weld isto be made. If a weld is to be made over afa i r i c we I bead , chip and fileI u re i n an e Ioff all of the existing bead before applyinga new weld to the area. After a weld is made,do not file or smooth the weld or apply anysolder or other filler to improve the appearanceof the we Id.^. ELECTRIC ARC WELDING OF STEEL.Before starting to weld, make certain thatthe surface of the parts to be welded is freeof loose scales, oxides, oil, and foreign natter.Jigs, clamping devices, and tack welding shallbe used wherever required to control warpingand ensure proper al ignment. Preheating, scarfing,and chamfering are not necessary on anytubing of the front fuselage truss, as none ofthe tubes are over .085 inch thick. Preheatingis required only on heavy fittings and forgings,and chamfering is required only on materialof .140- inch thickness and greater. The properelectric arc-welding rods to be used withchrome molybdenum steel shall be in accordancewith the following table (see Figure y). Ingeneral, a heavily coated, mild steel weldingrod should be used, and the diameter of[36]RESTRICTED

imileviatesFUSELAGEPAR. 1 TO 5the electrode should not be greater than thetube wall thickness unless the operator increasesthe travel speed sufficiently to preventoverheating, undercutting, and burningthrough the metal. The length of the arc shouldbe held to approximately 1/8-inch. Polarityshall be as recommended by the electrode manufactureror as found suitable for the specificwork being accomplished. All tack welding andall weld endings shall be accomplished withit he use of a c rate r e I i m i nat or i f one ava iable (see Figure 6). The small hole or craterat the weld end creates a stress concentrationand a fatigue point, and should therefore beeliminated if possible. The use of a cratere I nat or will a I this condi t ion anda smooth weld end will result (see Figure 8).Avoid rewelding as porosity in the weld mayresult. When a weld is built up by two ormore beads or passes, the preceding weld mustbe cleaned free of scale or flux by chippingor scraping and followed by brushing with a.RESTJilCrEDWING CENTER SECTIONATTACHMENT PROVISIONSGUNNER'S SEAT BRACE TUBEREF. DWG. 77-31105Figure 4— Front Fuselage Truss Assemblywire brush. Welding shall not be dressed byremoving metal from the joint unless furtherwelding is to be done on the dressed region.Unless otherwise specified, the maximum widthof welds for material thickness of not over.040 inch shall be 1/4-inch (see Figure g)I-.The depth of penetration shall be between 25and 40 percent of the thickness of the basemeta I5. OXYACETYLENE WELDING OF STEEL.I iThe oxyacetylene process is still consideredthe most flexible type of welding and general lybest suited for repair work on aircraft. Hewever,both electric arc welding and oxyacetylenewelding are acceptable, and repairs outnedwill be appi icable for either type ofwelding. Welding wire used for oxyacetylenewelding of chrome molybdenum tubing must conformto Spec. QQ-W-351, Grade E. The torchtips should be of proper size for theRESTRICTED[37]

sRESTJilCTEDFUSELAGEPAR. 5 TO 6FORWARDLEG OFNOSE-OVERSTRUCTURE //:Figure S-^Upper Tube Cluster Joint at Station 55thickness of the material to De welded. Thefollowing commonly used torch tip sizes aresat i factory :THICKNESS

FUSELAGEPAR. 6 TO 7RESmiCTEDBASEMATERIAL

.RESTRICTEDFUSELAGEPAR. 7 TO 10Figure 10—Oxyacetylene Yielding Equipmentinjured by contact with heat. Sheet asbestosor heavy wet rags will furnish sufficient insulationfor adjacent furnishings. All fireprecautions should be observed before any weldingis started. All tubing or sheet stockused for repairs must be X4I30 chrome molybdenumsteel. Tubing used for telescope reinforcementsor for spl icing must be of the samewall thickness as or greater than that of theoriginal member (see Figure 2J8. NEGLIGIBLE DAMAGE TO STEEL TUBES.Some forms of damage to tubular structuresmay be considered negligible. Such damage maytake the form of slight indentations, scratches,or minor bowing. Smooth dents not exceeding1/20 of the tube diameter in depth, withoutcracks, fractures, or sharp corners, and clearof the middle third of the length of the membermay be disregarded except to satisfy appearance.Tubular members should be carefullyscrutinized for the presence of sharp nicksand deep scratches which may have resultedfrom service. These nicks and scratches producestress concentrations that may cause failureof the part. Care must be taken to smoothout all sharp nicks and deep scratches witha fine file, fine emery paper, or steel wool.When this is accomplished, high concentrationsof stress disappear.Figure ll^Typical Welding Failures9. ESTIMATING EXTENT OF DAMAGE TO STEEL TUBES.When inspecting the forward fuselage trussand engine nount for possible danage, the structuresurrounding any visual damage must becarefully scrutinized to ensure that no secondarydamage remains undetected. Secondarydamage may be produced in some structure remotefrom the location of the primary damageby the transmission of the damaging concentratedload over the normal course toward thebasic reacting structure. Damage of this natureusually occurs where the most abrupt changein load travel is experienced. If this damageremains undetected, loads applied in thenormal course of operation may cause failureof the part. Visually examine all joints forcracks, welding flaws, or failures (see Figureii). If any doubt exists as to the presenceof cracks or other flaws, test the structureas outlined in an applicable subsequent paragraph.10. SMOOTH DENTS IN STEEL TUBES.A minor smooth dent in steel tubing may oftenbe removed by the following procedure: RenxDveone of the self-tapping screws provided at theextremities of the main steel tubes and apply[40]RESTRICTED

low-shapedFUSELAGEPAR. 10 TO 12RESTRICTEDAPPLY WRENCHAND TIGHTENSCREWTUBE WILL BEROUNDED WHENBLOCKS COMETOGETHERHEAVY DUTY"C" CLAMPGROOVED METALBLOCKSNOTE^ EGG-SHAPEDAREA MAY BEHEATED TO A DULLRED AS AN AIDTO RE-FORMINGFigure 12'-'Correcting Oval- shapedSteel Tube Distortionan air pressure of upwards of 75 Ibs./sq.in.to the inside of the steel truss. Heat thedented area evenly to a dull red with an acetylenetorch until the internal air pressureforces out the dent and restores the originaltube contour. If internal air. pressure andheat are not sufficient to remove the dent,tack weld a welding rod to the center of thedent and pull on the rod while heating the area.After the dent is removed, a I the area toair-cool and then release the internal air pressure.Replace the previously removed selftappingscrew.Important: Do not apply heat over a dul I redto the middle third of the lengthof any tube.li. STEEL TUBE CIRCUMFERENCE BENT TO AN OVALSHAPE.Where the circumference of a steel tube isbent to an. oval shape, the area may be restoredto normal in the cold condition by pressureexerted on the area through grooved steel formblocks (see Figure 12). Drill a steel blockto the diameter of the damaged tube; then sawthe block along the axis of the hole and separatethe two sections of the block. Applythe two form block sections to the oval-shapedarea on the affected tube. Slip a heavy clampover the blocks, tighten the clamp, and exertpressure on the area until the oval-shaped tubearea is restored to the normal circular shape(see Figure 12). If difficulty is encounteredin shaping the tube in the cold condition, heatthe area to a dull red; then apply the steelblocks and clamp. Remove the clamp and theblocks. If the oval-shaped area is longer thanthe length of the steel form blocks, reapplythe form blocks and the clamp to successiveaffected areas until the entire length of theova I area is restored to the normal circularshape.12. BOWED STEEL TUBES.Steel tubes which have been bowed withoutevidence of cracking may be straightened inthe cold condition as shown (see Figure 13).Cut three hardwood blocks grooved to fit thecontour of the tube, and line the grooves withleather or canvas. Obtain a length of channeliron equal to the length of the bow in the tube.Locate one of the grooved wood blocks at eitherextremity of the bow and apply the channel ironbeam so that the beam spans the bowed area andbacks up the two blocks. Apply the third blockon the opposite side of the tube at the pointof the maximum bend near the center of the bow.SI ip one end of a heavy "C" clamp over the channeliron beam and tighten the clamp down on theblock at the center of the bend. In order toallow for springback of the tube, continuetightening the clamp until the tube is bentslightly in the opposite direction (see Figure13). Remove the clamp and the blocks. Checkthe alignment of the tube by placing an accuratestraight edge on both the side and thetop of the tube. If the straight edge checkreveals a slight bcw in the tube, reapply theblocks and the clamp and check until the tubelines up with a straight edge in both referenceplanes. The maximum allowable tube bowis 1/16-inch measured from the straight edgeto the outer circumference of the tube. Ifthe tube is free from cracks and is withinthe allowable maximum bow of 1/16-inch afterstraightening, the tube is perfectly acceptablewithout additional reinforcement. However,if cracks appear at the point where themaximum bow was corrected, dri II a No. 40 holeat the ends of the crack and weld a spl it steelsleeve over the area as outlined in theRESTRICTED[41 J

RESTRICTEDFUSELAGEFIGURE 13PLACE BLOCK ATCENTER OF BENDFigure 13— Straight ening Bowed SteelTvbes[42]RESTRICTED

FUSELAGEPAR. 12 TO mRESTRICTEDPAINT MAINTUBULAR TRUSSYELLOW- GREENFigure 14^^ppl icat ionBEFORE REPAINTINGPLAGE MASKINGTAPE OVER TUBESAT CLIP LOCATIONSof Finish to Steel Tubesand weld an overlapping doubler over the areaas outlined for the repair of dents near acluster joint. No more than two cracks shouldbe repaired in the same general area. At theconclusion of the repair, apply one coat ofzinc chromate primer to the area from whichthe finish was previously removed. AddIy finishcoats to match the adjacent surface. Thesefinish coats consist of either two coats ofyel low-green lacquer or two coats of clearlacquer containing a minimum of 4 ounces ofaluminum paste per gal Ion of spraying material(see Figure 14}.m. SHARP DENTS AT A STEEL TUBE CLUSTER JOINT.Sharp dents at a steel tube cluster jointmay be repaired by welding an especially formedX4I30 chrome molybdenum steel patch plate overthe dented area and surrounding tubes (seeFigure 15)' To prepare the patch plate, cuta section of X4I30 chrome molybdenum steelplate of a thickness equal to or greater thanthat of the damaged tube. Trim the reinforcingplate as shown (see Figure 1^) so that theplate extends a minimum of two t imes theparagraph applicable to cracked tubes. In e\/erycase where a bent tube is restored, carefullytest all adjacent welded joints for cracks, andrepair the cracks as outlined in the followingparagraph.13. SMALL CRACKS AT STEEL "RJBING CLUSTER JOINTS.If it is necessary to check an individualtubing joint for cracks, apply a liberal coatof light oil to the affected area, thoroughlywipe the oil from the joint, and then applya c oa t of wh i t i ng . Ac rac k in t he j o i nt willbe made evident by the appearance of oil onthe whiting frcm the crack recess (see Figure 12).If a considerable number of welded joints must betested, another method of testing, using airpressure, is described in a subsequent paragraph.After locating the crack, remove the whiting andremove all finish from the area by rubbing withsteel wool or a wire brush. If the crack islocated in an original weld bead, carefullychip, file, or grind out the existing weldbead, and reweld over the crack along the originalweld line. When grinding off the originalweld bead, take particular care to avoidremoving any of the existing tube or gussetmaterial. If the crack is near a c luster jointbut away from the original weld bead, removethe finish from the area with steel wool, drilla No. 40 (.09P) hole at the ends of the crack.T S S = T OR T+MINIMUMcur AND FORM REINFORCING PLATEWELD ALL EDGES OFREINFORCING PLATECOMPLETE REPAIRFigure IS"—Reinforcing a Dent at aSteel Tube Cluster JointRESTRICTED[43]

RESTRICTEDFUSELAGEPAR. m TO 16diameter of the tube from the nearest edge ofthe dent, and over adjacent tubes 1-1/2 timesthe diameter of the tube. On the damaged clusterJoint area to be covered by the reinforcingplate, rub off all the existing finishwith steel wool. The reinforcing plate, ^aybe formed before any welding is attempted, orthe plate may be cut and tack welded to oneor more of the tubes forming the cluster jointthen heated and pounded around the joint contouras required to produce a smooth contour.Avoid unnecessary heating of the reinforcingplate while forming, but apply sufficient heatand pound the plate so that there is generallya gap of no more than 1/16-inch from the contourof the Joint to the reinforcing plate.While forming the plate, exercise care to preventdamage at the apex of the angle formedby any two adjacent "fingers" of the plate.After the reinforcing plate is formed and tackwelded to the cluster Joint, weld the plateedges to the cluster joint. At the conclusionof the repair, apply one coat of zinc chromateprimer to the plate area and to the accessiblearea from which the finish was previously removed.Apply finish coats to match the adjacentsurface. These finish coats consistof either two coats of yellow-green lacqueror two coats of clear lacquer containing aminimum of 4 ounces of aluminum paste pergallon of spraying material (see Figure 25J.15. SHARP DENTS OR CRACKS IN STEEL TUBE LENGTH.If a crack appears in a length of steel tube,usually as the result of previously straighteningthe tube, first drill a No. 40 (.098)hole at the ends of the crack, and then rubthe area with steel wool to remove the finisharound the tube for a distance of approximately3 inches on each side of the damage. If thedamage is in the form of a sharp dent whichcannot be removed by any of the methods previouslyoutlined, first rub the area with steelwool to remove the finish around the tube fora distance of approximately 3 inches on eachside of the damage. In order to reinforce thedent or the crack, select a length of a spareX4I30 chrome molybdenum steel tube sleeve havingan inside diameter approximately equal tothe outside diameter of the damaged tube andof the same wall thickness. Diagonally cutthe reinforcement steel sleeve at a 30-degreeangle on both ends so that the distance of thesleeve from the edge of the crack or dent isnot less than 1-1/4 times the diameter of thedamaged tube (see Figure 16). Cut through theentire length of the reinforcing sleeve and^DRILL NO 40 HOLE AT EACH END;CRACK-ARREST GROWTH OF CRACKpSNUG FITTING SLEEVE CUT IN HALf-jTT=B OR B +PREPARE REINFORCING SLEEVEh-liD-^( MINIMUM)WELD SLEEVE OVER CRACKCOMPLETED REPAIR ROTATED 90"Figure IS'-^Re in forcing a Dent or Crackin Steel Tube Lengthseparate the half sections of the sleeve. Clampthe two sleeve sections to the proper positionson the affected area of the originalsteel tube. Weld the reinforcing sleeve alongthe length of the two sides, and weld both endsof the sleeve to the damaged steel tube asshewn (see Figure 16). Apply "one coat of zincchromate primer to the reinforcing sleeve areaand to the accessible area from which the finishv\«s previously removed. Apply finish coatsto match the adjacent surface. These finishcoats consist of either two coats of ye I lowgreenlacquer or two coats of clear lacquercontaining a minimum of 4 ounces of aluminumpaste per gallon c»f spraying material.16. SPLICING STEEL TUBES - GENERAL.Tubular members of the forward fuselage trussmay be spliced by partial tube replacement usedwith internal or external reinforcing steelsleeves, or by the use of an externally telescopingtube replacement of the next larger diametertubing. Each type of splice has its particularadvantage or function, and the methodsinvolved are essent iai ly the same. All splicingis accomplished with electric arc welding oroxyacetylene welding. No bolted joints orsplices are permitted. If the original damaged[44]RESTRICTED

FUSELAGEFIGURE 17RESTRICTEDORIGINAL TUBEWELDSSPLICEMEMBPULLING WIRETRIMMED ANDWELDEDNTERNAL, SNUG-FITTINGSLEEVESiV-INCHSLEEVE-PULLINGWIRE**40 HOLE^ -*L= I" OR liD WHICHEVER IS GREATERFigure 17 —Steel TubeInner Sleeve SpliceRESTRICTED[45]

I determineI of/I6—i-nchI notI ingRESTRICTEDFUSELAGEPAR. 16 TO 17tube accommodates castings or fittings whichhave been fabricated to fit the tube contour,the spliced replacement tube must be of the samgdiameter. If no such fittings are applied tothe original damaged tube, the externally telescopingsplice replacement may be used. Twotypes of splice welds are permitted, the diagonaland the fishmouth. Being the strongerof the two, the fishmouth splice weld is preferredto the diagonal. However, the natureand location of the damage wi I whichtype must be used.Important: Splices must not be made in themiddle third of a tube bay, andonly one partial replacement tube can be insertedin any one bay of a structural member.If a member is damaged at a Joint so that itis impossible to retain a stub to which anothermember can be attached, replace the tube if itis a web member; and if the tube is a continuouslongeron, locate the splice in the adjacent bay.^'1ake no splice by butt welding any member betweenstation points. Nfeike no repairs to theengine mount members located within 3 inchesof the mounting ring, and make no repairs tothe mounting ring. Misalignment of a tubularstructure due to contraction and expansion ofthe metal during welding can be prevented byusing wood braces with notches in the ends tohold the tubes in position. When new tubes areused to replace bent or damaged tubes, the originalalignment of the structure must be maintainedand checked. This can be done by rreasuringthe distance between points of correspondingmembers on an undamaged aircraft.17. SPLICING STEEL TUBE BY INNER SLEEVEMETHOD.If the damage to a steel tube is such thatpartial replacement of the tube is necessary,the inner sleeve splice is recommended especiallywhere a smooth tube surface is desired(see Figure ij). Diagonally cut out the damagedportion of the tube with a hack saw, locatingthe cuts away from the middle third of the affectedtube bay. Burr and clean the edges ofthe cuts. Diagonally cut a replacement X4I30chrome molybdenum steel tube to match the diameterwall thickness and length of the removedportion of the damaged tube. At each end ofthe replacement tube, allcw a 1/8-inch gap fromthe diagonal cuts to the stubs of the originaltube. Select a length of X4I30 chrome molybdenumsteel tubing of the same wall thicknessand of an outside diameter approximately equalto the inside diameter of the damaged tube.This inner sleeve tube material should fit snuglyIIwithin the original tube, with a maximum toleranceof 1/64— inch. From this inner sleeve tubematerial, cut two sections of tubing, each ofsuch a length that the ends of the inner sleevewill be a minimum distance of \-\/A tube diametersfrom the nearest end of the diagonalcut (see Figure ly). Dip the replacement tubeand the inner sleeves into hot (74°C, I65°F)raw nseed oil; then w i pe the o i from theI iouter circumference of the tubes. Make a markon the outside of the diagonally cut originaltube stub midway along the diagonal cut (seeFigure i8). At a minimum distance of 2-1/4times the tube diameter measured from the nearestend of the diagonal cut, center punch thetube, and start drilling the No. 40 hole at a90-degree angle. After a shallow hole isstarted from which the dri II wi I jump out,slant the drill toward the cut and drill at a30-degree angle. Slanting the hole in this manneraligns the edges of the hole with the lineof pull of the sleeve-pulling wire, and preventsthe wire from scraping the hole edges. Burrthe edges of the hole with a round, needle-pointfile. Obtain a length of we Iding orbrazing wi re, insert one end into the dri I ledhole, and push the wire out the end of the tube(see Figure i8) . Weld the end of the wire tothe inner side of the reinforcing sleeve. Chamferthe end of the sleeve as an aid in slidingthe tube into position. With thin paint, metaldye, or emery paper, make a narrow markaround the center of the reinforcing sleeve.Slip the sleeve into the replacement tube sothat the welded wire is (80 degrees from thedrilled hole. If the inner sleeve fits yer^tightly in the replacement tube, chill thesleeve in cold water and heat the replacementtube and the stubs of existing tube.. The resultantexpansion of the tube stubs and the contractionof the inner sleeve temporarily allcxvmore clearance from the inner sleeve to the inside wa I the tube stubs. Al ign the originaltube stubs with the replacement tube. Pullon the exposed end of the s leeve-pu I I ing wireuntil the center mark on the sleeve is directlyin line with the center mark on the diagonalcut (see Figure i8). When this occurs, theinner sleeve is centered beneath the joint.Sharply bend the pulling wire over to hold thesleeve in position. At each side of the replacementtube, weld the inner sleeve to thetube stubs through the 1/8-inch gap between thestubs (see Figure iS). Completely fill the 1/8-inch gap and form a weld bead over gap. Afterthe joint is we Ided, sni p off the pu I wi reflush with the surface of the tube and weldover the hole. To the outside of the replacementtube, apply one coat of zinc chrcmate primer,L46]RESTRICTED

fromlewlowthicknessFUSELAGEPAR. 17 TO 19RESTRICTEDand match adjacent finish with two coats ofyellow-green lacquer or two coats of clearlacquer containing a minimum of 4 ounces ofaluminum paste per gal Ion of spraying material.18. SPLICING STEEL T'IBE BY OUTER SLEEVEMETHOD.If partial replacement of a tube is necessary,an outer sleeve splice may be used, whereconditions warrant, as an alternate splice tothe inner sleeve splice and the splice usinga larger diameter replacement tube. However,the outer sleeve splice requires the most amountof welding; and therefore, it should be usedonly where the other splicing methods are notsuitable. Squarely cut out the damaged sectionof the tube, locating the cuts away fromthe middle third of the tube bay. Cut a replacementX4I30 chrome molybdenum steel tubesection to match the outside diameter, wallthickness, and length of the removed tube. Thisreplacement tube must bear against the stubsof the original tube with a total tolerance notto exceed 1/32- inch- Dip the replacement tubeinto hot (74°C, I65°F) raw linseed oil; then wipethe oi I the outer circumference of thetube. Select a length of X4I30 chrome molybdenumsteel tubing of an inside diameter approximatelyequal to the outside diameter ofthe damaged tube and of the sarre wal I thickness.This outer sleeve tube material should fitsnuglyabout the original tube with a maximum toleranceof 1/64-inch. From this outer sleevetube material, diagonally or fishmouth cut twosections of tubing, each of such a length thatthe nearest ends of the outer sleeve are a minimumdistance of 1-1/4 tube diameters from theends of the cut on the original tube (see Figureig) . Use a f ishmouth-cut sleeve wherever possible.Burr all the edges of the sleeves, replacementtube, and briginal tube stubs- Slipthe two sleeves over the replacement tube, lineup the replacement tube with the original tubestubs, and slip the sleeves out over the centerof each joint (see Figure ig) . Adjust thesleeves to suit the area and to provide maximumreinforcement. Do not tack weld the twosleeves to the replacement tube before welding.Apply a uniform weld around both ends of oneof the reinforcing sleeves and a I the weldto cool. Then weld around both ends of- the remainingreinforcing tube (see Figure ig)^ Allowingone sleeve weld to cool before weldingthe remaining tube prevents undue warping. Tothe outside of the replacement tube, apply onecoat of zinc chromate primer and match adjacentfinish with two coats of ye I low-green lacqueror two coats of clear lacquer containing a minimumof 4 ounces of aluminum paste per gallon ofs pray i ng mate rial,19. SPLICING STEEL TUBE USING LARGER DIAMETERREPLACEMENT TUBE.This method of splicing steel tubes requiresthe least amount of cutting and welding. Hewever,this splicing method cannot be usedwherethe darmged tube is cut too near the adjacentcluster joints or where bracket mounting provisionsmake it necessary to maintain the samereplacement tube diameter as the original. Asan aid in installing the replacement tube,squarely cut the original damaged tube, leavinga minimum short stub equal to 2-1/2 tube diameterson one end and a minimum long stub egualto 4-1/2 tube diameters on the other end (seeFigure 2o). The cuts must be away from themiddle third of the affected tube. Select aspare length of X4I30 chrome molybdenum steeltubing having an inside diameter approximatelyequal to the outside diameter of the damagedtube and of the same wal I as or greaterthan the damaged tube. This replacement tubematerial should fit snugly about the originaltube with a maximum tolerance of 1/64- inch.From this replacement tube rmterial, diagonallyor fishmouth cut a section of tubing of sucha length that each end of the tube is a minimumdistance of 1-1/4 tube diameters from theend of the cut on the original tube. Use af ishmouth-cut replacement tube wherever possible(see Figure 20). Hcwever,a diagonally cuttube may also be used (see Figure 21)^ Burrthe edges of the replacement tube and the originaltube stubs. If a fishmouth cut is used,file out the sharp radius of the cut with asmall round file. Dip the replacement tube intohot (74°C, I65°F) raw linseed oil; then wipe theoil from the outer circumference of the tube.Spring the long stub of the original tube fromthe normal position; slip the replacement tubeover the long stub, then back over the shortstub. Center the replacement tube between thestubs of the original tube. In several placestack weld one end of the replacement tube; thenweld completely around the end. In order toprevent distortion, a I the weld to cool completely;then weld the remaining end of the replacementtube to the original tube (see Figure20). To the outside of the replacement tube,apply one coat of zinc chromate primer andmatch adjacent finish with two coats of ye I lavgreenlacquer or two coats of clear lacquercontaining a minimum of 4 ounces of aluminumpaste per gallon of spraying material.RESTRICTED[47]

-RESTRICTEDFUSELAGEFIGURE 18SLEEVE WILL ENDHEREMAKE MARK ATCENTER OF CUT^DRILL*40 HOLE,SLANTEDMINIMUMMINIMUMM/J/?/

inginseedFUSELAGEPARAGRAPH 20RESTRICTEDtMAXIMUM32 GAPVr/SPLICE MEMBER SQUARE-CUT ENDOF ORIGINAL TUBE-SL=l"0R|'/4D ^ C B1S = DWHICHEVER IS GREATER^^^ ^^ ^"^ ^ = ^ OR B +CC/r OUT DAMAGED PORTION OF TUBE, PREPARE SPLICE MEMBER AND CUT SLEEVESSLIP SLEEVES OVER JOINTS AND WELD IN PLACECOMPLETED SPLICE USING DIAGONALLY- CUT SLEEVESSLEEVES MAY BE ROTATEDTO SUIT CONDITIONS ANDTO PROVIDE MAXIMUMREINFORCEMENT\iHMlii||iiMiC A BC=B OR B-»- A=B OR B +COMPLETED SPLICE USING FISHMOUTH-CUT SLEEVESFigure 19 — SteelTube Outer Sleeve Splice20. REPLACING STEEL TUBES.Tubes damaged beyond the limitations specifiedin the preceding paragraphs on tube splicingshould be replaced. Tube replacement is necessarywhere an original tube stub is too shortto attach a replacement and where splice weldswill be made in the middle third of a member.When it is necessary to remove a member at ajoint or from a cluster of tubes, use a sharpfine-toothed hack saw and remove the tube carefully and completely from the structure. Whilecutting out the tube, exercise caution to preventany damage to adjacent tubes or welds.Where new welds are to be made over the locationof existing welds upon the insertion ofthe new member, completely chip or file off theold welds. Dip the replacement tube into hot(74°C, I65°F) raw I oi I; then wipe the oi Ifrom the outer circumference of the tube. Wheninstal I a new tube member, al lav a clearanceof 1/32- inch at either end for expansion. Unlessa welding j ig is avai lable, the actualprocess of welding should be accomplished inas systematic a manner as possible pertinentto the application of heat and the resultantpossible distortion. After the new tube hasbeen welded in place, clean the welded joints witha wire brush. To the outside surface of thereplacement tube, apply one coat of zinc chromateprimer, and match adjacent finish with twocoats of yel low-green lacquer or two coats ofclear lacquer containing a minimum of 4 ouncesof aluminum paste per gallon of spraying material(see Figure 14). Check the alignment ofthe truss as outlined in a subsequent paragraph.RESTRICTED[49]

RESTRICTEDFUSELAGEFIGURE 20LONG STUB ON ONEEND TO FACILITATESLIPPING SPLICEMEMBER OVERORIGINAL TUBESNUG FITTING,TELESCOPINGSPLICE MEMBER-ORIGINAL TUBEL= I" OR Iy D WHICHEVER IS GREATERl^/EiV A. ROTATED 90"Figure 20 — Steel Tube Fishmouth Splice Using Larger Diameter Replacement Tube[50]RESTRICTED

welded.I ightI redFUSELAGEPAR. 21 TO 21* RESTRICTEDpOIMUb -SNUG FITTING,r-| I I INt),ST TELESCOPINGJ SPLICE MEMBERpORlGINALpTUBE^S-T OR T+ [-« L ^1 DL=l"OR liD, WHICHEVER IS GREATERFigure 21-^SteeI Tube Diagonal Splice UsingLarger Diameter Replacement Tube21. TESTING WELDED STEEL JOINTS.In the event of an accident not visual ly indicatingdamage, the fuselage frame should betested for the security of the welded jointsas outlined in the previous paragraph on smallcracks at steel tubing cluster joints. Anothermethod of testing is as follows: Remove oneof the self-tapping screws provided at the extremitiesof the main members, and apply an airpressure of 30 Ibs./sq.in. to the interior ofthe tubular framework. Thoroughly mix one cupof liquid castile soap with one pint of waterand apply this solution with a soft-bristledbrush to al Ijoints or to the joints inquestion. A cracked weld will become evidentby the formation of large soap bubbles. Markthe weld for subsequent repair. The entire forwardfuselage truss should be airtight to preventacceleration of corrosion. Upon completionof testing, reinsert any self-tappingscrews that were removed. Typical welding flawswhich can be easily located by this method areillustrated (see Figure ii)tat ing major repair occurs to the forward fuselagetruss, care must be exercised to maintainproper alignment. In cases of major misalignmentconcluding repair, replace the structuralmembers affected. Minor misalignments whichcannot be detected by a visual inspection canusually be detected by means of simple measuringdevices and straight edges. One practicalmethod is to measure a corresponding memberon undamaged aircraft and compare measurements.If necessary, the application of heatmay be utilized as. a means of relieving minormisal ignment of a tube. Apply heat to a shortdistance on either side of the outside pointof maximum curvature. Heat to a du I andremove the torch. Allow the tube to air-cool.DO NOT QUENCH. Contraction of the metal willforce the tube into proper alignment. Theapplication of heat may be accomplished withoutdisassembling the aircraft provided allmetal parts or items liable to damage byproximity to heat are removed or otherwise protectedby asbestos sheeting or damp rags.21. EXHAUST MANIFOLD ASSEMBLY CONSTRUCTION.The exhaust manifold and collector ring assemblyconstitutes one part of the engine installation,and col lects and carries away the exhaustgases from the engine. The entire assembly isconstructed of stainless steel and is of theslip joint type (see Figures 22 and 23J . Thecollector is made from sheet stock, and the smallcomponent parts are made from bar stock and22. PROTECTING STEEL TUBE AGAINST CORROSION.Prior to welding, dip all replacement tubesand inner reinforcing sleeves into hot (74°C,I65°F) raw linseed oil; then wipe the oil fromthe outer circumference of the tubes. Thistreatment protects the interior of the tubesfrom corrosion. Outer reinforcing sleeves donot require this treatment. After weld ing, applyone coat of zinc chromate primer to all outsideunfinished surfaces, and match adjacentfinish with two coats of yellow-green lacqueror two coats of clear lacquer containing a minimumof 4 ounces of aluminum paste per gal Ionof spraying material (see Figure 14).23. ALIGNMENT OF ENGINE MOUNT AND FRONT FUSE-LAGE TRUSS.In the event that structural damage necessi-Figure22— Exhaust Mani fold AssemblyRESTRICTED[51]

icateI buttlowsIRESTRICTEDFUSELAGEPAR. 24 TO 28forgings. The collector is assembled by meansof arc and gas welding. Arc welding is used onall fittings such as exhaust port bolt flanges,collar bolt tubes, doublers, and lap joints.Gas welding is employed on a I joints ofthe main collector ring assembly.BEADED END PLAIN FNO25. EXHAUST MANIFOLD REPAIR.All repairs to the exhaust manifold assemblyother than part replacement are accomplishedby welding. The type of welding used is notcritical, as both gas and arc welding are satisfactory.However, welded repairs should follovthe design of the original construction asfar as possible. Permissible welded repairsconsist of welding cracks around exhaust portnipples, welding reinforcement ordoublers os/erfatigue points, welding patches o\/er bullet orother small holes, and repairing cracks in themain body of the col lector ring. For gas weldingthe manifolds, use a mixture of Cromaloy(Linde Air Products) or equivalent, and Inconelflux with sodium si I and water or withshellac as a binder. Use a neutral flame to preventoxidation and carbon pickup. Point the flamein the direction of the weld progress to preheatthe metal. The weld should be made asrapidly as possible to prevent an excessive puddlingand carbon pickup. For arc welding theexhaust manifold, use a coated, stainless steelelectrode of 1/16- or 3/32-inch diameter. Afterwelding has been accomplished, the parts mustbe passivated to improve their corrosion resistance.Irmierse the parts in a 17 to 20 percentaqueous solution of nitric acid at 49° to66°C (120° to I50°F) for 21 minutes, and thoroughlyrinse in water. The use of hot water forrinsing gives better results and promotes rapiddrying.MAIN EXHAUSTSTACKFigure 24 — Exhaust Manifold Slip Joint26. EXHAUST MANIFOLD COLLAR REPLACEMENTS.The col lar assembi ies which join the varioussections of the collector ring are of the singlebeadedtype which a I for si i ppage at aljoints (see Figures 23 and 24). Worn collarsshould be replaced and new bolts used to securethem. The collar bolt tubes may be replacedif they break. Grind the original weld from thecollar before welding the new bolt tube intoplace. The collar may be used as a jig byinserting the bolts and by tacking the newbolt tube into position. The bolts may thenbe removed and the weld completed.27. EXHAUST MANIFOLD REASSEMBLY.Before reassembly of repaired members, allwelded repairs should be lightly sandblastedto clean them. After sandblasting, the partsmust be passivated as outlined in a precedingparagraph. Upon reinstallation, care shouldbe taken to align all parts properly to preventexcessive wear. The brass nuts securingthe exhaust port flanges should be tightenedto 165-175 inch-pounds. Brass nuts should beused on all collar attaching bolts and shouldbe tightened one castellation beyond fingert ightness.28. ENGINE FIREWALL CONSTRUCTION.Figure 23— Exhaust Mani fold Assembly ExplodedA corrosion- resistant, .019 inch thick sheetsteel firewall is located at the forv/ard endof the main fuselage truss assembly at fuselagestation 0. This firewall forms a fireproofbulkhead between the engine and the pilot'scompartment. Five stiffeners are locatedon the aft side of the firewall to add rigidityand strength and to keep cans from forming inthe bulkhead between attachment points (seeFigure 2^). Four round holes in the firewallpermit the firewall to be slipped over the four[52]RESTRICTED

loyI by.I ing1 aFUSELAGEPAR. 28 TO 29RESTRICTED.038FIREWALLSTIFFENERSiS^.-,y-^^^^.019" CORROSIONRESISTANT STEEL-"sisiiSi:/ JlCOWLINGATTACHINGANGLE1COTTON WEBBINGAIR FOIL STRIPAFT SIDE FORWARD SIDEFigure 25— Engine Firewall Assemblyengine mount attachment bolt fittings. FourU-bolt assemblies attach the firewall to theforemost diagonal tube member of the forwardfuselage truss. A break in the horizontal firewallstiffening angles on the aft side of thefirewall permits the firewall to be clamped flatagainst the surface of this tube. A corrosionresistantattaching angle is spot-welded aroundthe periphery of the firewall. Dzus fastenerattachment springs are located at intervalsaround this angle to provide for the attachmentof the engine accessory compartment cowling,the landing wheel fairings, and the fairinglocated around the periphery of the fuselagejust aft of the firewall (see Figure 25J29. FIREWALL WEB REPAIR.Removal of the firewall cannot readi ly be accomplishedto permit repair members to be spotweldedto the firewall. Repair members must,therefore, be secured to the f i rewa I meansof AN44I-4 iron rivets, 1/8- inch diameter. Ifthese rivets are not available, Type AN442-AD4(AI7S aluminum a I ) rivets may be substitutedif they are dipped in zinc chromate primer priorto driving. This rivet corrosion protectionis necessary because of the contact of dissimiI iar metals. Before dri I holes in the f i rewaI I , check the opposite side for clearance onall engine accessories, electrical junctionboxes, and other equipment. Holes through thefirewall may be repaired by riveting on a flatpatch of iOIQ inch thick or thicker corrosionresistantsteel (see Figure 26). Cut awaythe damage to form a circular opening. A slewspeedhole saw or a round file may be used toaccomplish this. Cut a patch of .019 inch thickor "thicker corrosion-resistant steel sheet.Al low a 3/4-inch overlap around the edge ofthe hole; and with a No. 50 dri I rcwI, dri Iof holes around the patch in the lap area. Spacethe holes approximately 5/B-inch apart and 3/8-inch in from the edge of the patch to providep" roper edge distance for the rivets (see Figure26). Upon completion of drilling, removethe patch and clean all surfaces of chips.Realign the patch and insert two or more skinfasteners to hold the patch in place until twoor more rivets have been driven. Insert andupset an AN44I-4 iron rivet, 1/8-inch diameter,in each hole. The rivets may be driven fromRESTRICTED[53]

I toI firewall-LDRESTRICTEDPAR,FUSELAGE29 TO 323/4 MINIMUMOVERLAP\AN44I-4 IRONRIVETS-SPACED5/8" ON CENTER5/8Repairs to the corrosion- resistant steel cowlr-SPOTWELDS\ (REF.) ,/O O Oii^.038 CORROSION-RESISTANTSTEEL SPLICE MEMBERAN44I-4 IRON RIVETS•r(1 2 REQJo o o.019 OR .019 +CORROSION - RESISTANTSTEEL PATCH PLATE.019 THICKCORROSIONRESISTANT STEELFIREWALL CREF.)-FIREWALLSTIFFENER-Hi/^h-.019 CORROSION RESISTANTSTEEL FIREWALL (REF)Figure 26— Firewall Web RepairFigure 27 — FirewallVide- flanged Sti ffener Spliceeither side of the firewall, depending uponworking area available.30. FIREWALL STIFFENER REPAIR.Inasmuch as a I stiffening anglesare spot -welded to the f i rewa I I, damaged anglestiffeners cannot be removed without subsequentdamage to the firewall. To splice these stiffeners,proceed as follows (see Figure 21):Smooth out the damage with a f i le and remove a I Isharp edges. Fabricate a simi lar stiffener ofcorrosion-resistant steel of the same gage andcross-sectional area as that of the damagedmember. Center this splice member over thedamage. The length of the splice member shouldbe such that it wi II extend at least 5 incheson each side of the damage and should be securedto the damaged stiffener by three equal \y spacedrivets on the upstanding flange on each sideof the damage and attached to the firewal Iflangeby three equal ly spaced rivets on each side ofthe damage (see figure 27J . If the upstandingflange of the stiffener is not at least b/f^inchin height, the rivets in the upstandingflange must necessarily be omitted. Where thiscondition exists, the splice member should bethicker than the damaged angle section, oneadditional rivet must be used on each side ofthe damage, and the length of the splice memberincreased accordingly (see Figure 28). Whendrilling corrosion-resistant steel, use a slowspeeddri I avoid burning the tip.31. COWLING AHACHING FIREWALL ANGLE REPAIR.ing attaching angle which is spot-welded aroundthe periphery of the firewall are made in thesame manner as set forth in the preceding paragraphon firewall stiffeners. The splice membermust be curved slightly to conform to the curveof the attaching angle. This can be accomplishedwith the use of a metal shrinker.32. REPAIR MATERIALS FOR STEEL STRUCTURES.The followi-ng list is a summation of thematerials required for the repair of the enginemount, the front fuselage truss, the exhaustmanifold assembly, and the engine fi newel I.MATERIALSSHEET, STAINLESS STEELTUBE, CHROME MOLYBDENUMSTEEL X4130RIVETS, ANUUl-U, 1/8-INCH DIAMETERPRI MER, Zl NC CHROMATEWHITINGROD, WELD INOSHEET, CHROME MOLYBDENUMSTEEL XUI30FLUX, CROMALOY (LINDEAl R PRODUCTS CO. ,NEW YORK)OIL, LI NSEEDLACQUER, CLEARP ASTE , ALUM I NUMLACQUER, YELLOW-GREENREMARKSN AA SPEC. SS 2118DIAMETER AND WALL THICK-NESS TO SUIT (spec57-I8O-2)MI steel, LENGTH ASREQU I REDUSE AS DIRECTED (SPEC.1U080)STAINLESS STEEL-COATED,AND CHROME ALLOYSPEC AN-QQ-S-685INCONEL WITH SODIUM SILI-CATE AND WATER, OR WITHSHELLAC AS A B INDERSPEC. JJJ-0-336SPEC. 3-158SPEC. TT-A-tt66; TYPE BSPEC. 3-100-H[54]RESTRICTED

RED.FUSELAGEPAR. 33 TO 3»*RESTRICTED.038% CORROSION- RESISTANTSTEEL SPLICE MEMBER1AN 44 -4 IRON RIVETS (8 REQ.)019 CORROSIONRESISTANT STEELFIREWALL (REF)stack or from the fixed fuselage gun blast aremade of .025 and .038 inch thick corrosionresistantsteel. Cowling arrangement, material,material thickness, and major assembly partnumbers are illustrated (see Figures 29 and c^o)Padded support brackets on the engine supportthe engine cowling assembly, which is held inposition by aligning pins and is securely clampedaround the periphery of the engine by one exteriorand two interior clamping bolts (seeFigure 2g) • The flush-type dzus fasteners which.secure the engine accessory compartment cowl ing•-SPOT WELDSFigure 28-'Firewall Narrow- flanged Stiffener Splice33. REPAIR TOOLS FOR STEEL STRUCTURES.To repair the various members of the enginemount properly, the forward fuselage truss,the exhaust manifold assembly, and the enginef i rewal I, all or part of the fol lowing toolsare requ i red.TOOLREMARKSEQUIPMENT, GAS AND ARC WELD-INGSEE FIGURES 6 AND 10TOOLS, SHEET METAL CUTTING,FORMING, AND BENDING ASREQU I SEE SECTION 1TOOLS, RIVETING, STANDARD SEE SECTION 1BRUSH, WIREBRUSH, PAINTSTEEL WOOL34. ENGINE COWLING CONSTRUCTION.A streamlined, spot-welded, and flush-riveted,removable cowling assembly encloses and strearrvlinesthe engine, the exhaust collector ring,and the engine accessory compartment (see Figure'?o)', All removable parts of the cowlingare secured by means of flush-type dzus fastenersand interior and exterior clamping bolts.Kost of the cowl ing sections are made from .032and .040 inch thick 24ST alclad. All partswhich are subjected to heat from the exhaustPADDED SUPPORTING STRIPGUN TROUGHFigure 29—'Engine Cowling Cons true t ionRESTRICTED[55]

RESTRICTEDFUSELAGEFIGURE 3077-31002-1977-3107577-3103277-3102777-3103377-3103477-3106677-3103077-3106577-3107777-31002-1877-3103577-3102877-3104255-3106977-31002-1777-31066Figure 30— Engine Cowling Assembly[56]RESTRICTED

FUSELAGEPAR. 34 TO 37RESTRICTEDare secured to four stainless steel channelswhich extend from the firewall to the baffleplate ring just aft of the exhaust collectorring.35. ENGINE COWLING REPAIR.Permissible repairs consist of re-formingdented areas, replacements of damaged or brokendzus fasteners, and the patching of holes. Inasmuchas all parts of the cowling assemblyare in the high-pressure area of the air streamaround the engine, all patches to the cowlingmust be of the flush type, using a filler pieceto replace the removed material and an overlappingpiece on the inside of the cowling.Refer to Section III for flush-type skin repairsand to Section IX for removal and replacementof damaged or broken cowl ing dzus fasteners.Hole patches to corrosion- resistant steel sectionsof the cowling may be secured by spotwelding if such equipment is available, or asingle row of Type AN450D8 countersunk steelrivets may be used around the damage to securethe steel patch plate. Some of the corrosion--.040 X 1-1/2 X 2-3/4 SHEET24STALCLAD (2 REQ.)_DRILL NO. 30 (.1285)COUNTERSINK IOO°X7/32"AN426-AD4 RIVET(8REQ.)2|." PLUS LENGTHOF DAMAGEFigure 32-^ Front Fuselage Side PanelForward Former SpliceLEFT SIDE PANEL RIGHT SIDE PANELREF DWG. 77-31050 REF DWG. ATGAaB 77-31051AT6C 88-310511. 6e-3l073 - FORMER2. 52-31200-U - FORMER (3 REO-)3. CSeULT - STRINGER (12 REO-)8. S2-3l200-'i - FORMER - AT-CA- AT-ep66-31200U. ZEE SECTION STIFFENER - 2U ST5. S2-3I200-8 - FORMERe. f5IS-3l038-U - DOOR ASSEMBLY7. ee-31009 - VENTILATOR ASSEMBLY

RESTRICTEDFUSELAGEFIGURE 33LEFT SIDE1. ELEVATOR HORN RIGGING ACCESS 7. FLEXIBLE GUN STOWAGE SLOT2. VERTICAL STABILIZER- FRONT SUPPORT 8. UPPER LONGERONS3. TAIL HOIST TUBE 9. BAGGAGE COMPARTMENT DOOR4. ELEVATOR HORN OPENING lO RIGGING ACCESS5. HORIZONTAL STABILIZER SUPPORT II. SHOCK STRUT TRUNNION ACCESS6. REAR SECTION RIGGING ACCESSFigure 33-^ Aluminum Rear Fuselage Covered Assembly[68]RESTRICTED

FUSELAGEFIGURE 34RESTRICTED14 15 16 17 18 19 20 21 22 23 24 25 26ATAMERICANA(0/?r//CONTRACT NOS. DRAWING6A AT-6B AT-6C NOS.AVIATIONPART NUMBERSTITLEI .2.3.4.5.6.7.8.9.10.t I.12.13.11.15.16,17.18.19.20.•21,22.2324,25,2627,77 77 88 31 I 13 BULKHEAD - STATION 106.5 R.H.52 3lli|2 FITTING58, 31 122 LONGERON5577* 78 31 I 17 BULKHEAD - STAT ION 137 UPPER55 3 1391-3 BULKHEAD - STAT ION 15855 3 1416 FORMER - STATIO N 168-177* 31 I 19 BULKHEAD-STAT ION 17969 3 II 20 BULKHEAD - STAT ION 20052 31 109 FITTING -HORIZ ONTAL S66 31 I 14-12. ..BULKHEAD - STAT ION 21556 311 15-6 BULKHEAD - STAT ION 22755 31 I 16 BULKHEAD - STAT ION 2427 2 31 146 FORMER - STATIO N 106-15 5 3 1 169-6 FORMER - FUSELA GE REAR36 31289 MOUNT, C AMERA77 77* 31 170 FORMER -L .H.FUSELA GE REAR88* 31497 FORMER - FUSELA GE REAR6 4 6 4* 31245 BULKHEAD - STAT ION 13788* 31496 BULKHEAD - STAT ION 13777 31 I 12 BULKHEAD - STAT ION 13777 , 3 1 121 LONGERON -LOWE77 77* 3 139 5 FORMER -R L.H.STATIO 156-188* 31495 FORMER - STATIOI 56- I50 31 121 FORMER - STATIO I58-I55* 31396 FORMER -170-355 31397 FORMER -50.55.56.55.31 122 FORMER -31398 FORMER -31 123 FORMER -3 1402 CHANNEL* INDICATES CODED- FUSELAGE ATTACHING L.H.-UPPER L.H.STATIOSTATIOSTATIOSTATIOSTATIO-ASSEMPARTBLY185-9179-1200-3200-3-1/4 UPPER/2 UPPER LEFT-1/2 UPPER-3/4 UPPERTABILI ZER ATTACHING-5/8 REAR-3/8 REAR-3/16 UPPER11 LOWERSECTION IB LOWERSECTION IB LOWERSECTION IB LOWERLOWERLOWER1/16 LOWER1/ 16 LOWER/4 LEFT SIDE/8 LOWER/ 16 LOWER/2 LEFT SIDE/4 LOWER/4 LEFT SIDEFigure 34'^Aluminum Rear Fuselage Part NumbersRESTRICTED[59]

RESTRICTEDFUSELAGEPAR. 37 TO 39040X4-|/8"X2-3/8"SHEET24 ST ALCLADAN442-AD4RIVET-(8REQ.)TRIM REINFORCEMENTAROUND EXISTINGSTRINGER CUTOUTSwidth of 1-5/8 inches. Along the length of thesplice member, bend up a 1/4- inch, QO-degreeflange; and then bend up a 1/2- inch, 90-degreeflange measured from the first bend. Observea minimum bend radius of 1/8- inch. Trim thesplice member to match any existing stringercutouts in the affected area. With a No. 40(.098) drill, drill out the affected skin rivetsat each side of the damage. Clamp thesplice member to the proper position on thedamaged former. Center punch the rivet locationsin the noted pattern. Drill the center- punchedrivet locations with a No. 30 (.1285) drill.Remove the spl ice member, and burr all the rivetholes. Apply one coat of zinc chromate primerto all overlapping surfaces. Again clamp thesplice member to the damaged former; and intothe eight rivet holes, insert and drive AN442-AD4 rivets.39. ALUMINUM REAR FUSELAGE CONSTRUCTION.Figure 35-^Splice of Rear Four Formers ofFront Fuselage Side Panelsmaterial with a hack saw. To serve as the splicemembers, cut two sheets of .040 inch thick 24STale lad having a length equal to the damage plus2-3/4 inches and a width of approximately I-1/2 inches. Along the length of one of thesplice sheets bend a 90-degree, 1/2-inch flange.Along the length of the other splice sheet,bend a 90-degree, 5/8- inch flange. Trim thesplice sheets to the proper shapes. If thereare any skin rivets in the affected area, drillthem out. Clamp the splice members to the damagedformer. Center punch the required eightrivet locations in the noted pattern, and drillthe center-punched rivet locations with a No. 30(.I2B5) drill. Cut countersink the rivet holes100 degrees by 7/32- inch. Remove the splicemembers, and burr all the rivet holes. Applyone coat of zinc chromate primer to all overlappingsurfaces. Again clamp the members tothe damaged former. Into the rivet holes, insertand drive eight AN426-AD4 rivets.The semimonocoque rear fuselage section isI I feet 3.22 inches in length and is constructedwholly of 24ST a Ic lad, consisting essentiallyof four longeron formers and thin skin stiffenedwith stringers (see Figures 33 and 34J.The longerons extend the entire length of thestructure and form the upper and lower pointsof attachment to the forward fuselage trussassembly (see Figure 36). The upper and lowerATTACHMENT POINTSFORUPPER LONGERONS38. REAR FOUR FORMERS OF FRONT FUSELAGE SIDEPANELS.Any of the rear four formers of the frontfuselage side panels may be repaired as shown(see Figure -^f^) . If the damage is in the formof a crack, drill a No. 40 (.098) hole at theend of the crack. For the splice nember, cuta sheet of .040 inch thick 24ST a lei ad havingan approximate length of 4-1/8 inches and aATTACHMENT POINTSFORLOWER LONGERONSFigure 36— Attachment of Front Trussto Aluminum Rear Fuse lageL60]RESTFICTEE

1/16-inchII theoverlappingFUSELAGEPAR. 39 TO mRESTRICTED1. UPPER LONGERONS2. FUSELAGE FRAME3' SKIN STRINGERS4. BULKHEAD5. LIFT TUBEFigure 37— Aluminum Rear Fuselage6. LOWER LONGERONSInteriorlongerons are designed to resist the main bendingloads. A torsional ly rigid structure isforrred by the skin structure and the main bulkheads(see Figure 34J. The light, rolled-sheetstringers serve only to stiffen the skin andto prevent undue distortion and buckling. Inasmuchas the rear fuselage section is semimonocoque,the outer skin resists al I shear.40. ALUMINUM REAR FUSELAGE FORMERS.A typical repair which nay be used for a I I therear fuselage formers is illustrated (see Figure59 J . If the damage is in the form of acrack, perhaps at a stringer cutout, dri I I aNo. 40 (.09P) hole at the end of the crack.For the spl ice members, cut two sheets of .040inch thick 24ST alclad, each having a lengthof b-3/4 inches and a width of 2-3/4 inches.Along the length of one of the splice rtembers,bend anI flange. Along the lengthof the other splice member, bend a 13/ 16- inchflange. Observe a minimum bend radius of 1/8-inch. Trim the splice members around any existingcutouts in the damaged former, Dri Iout the affected skin rivets at each side ofFigure JS'—Aluminum Rear Fuselage InteriorExplodedthe damage, using a Mo. 40 (.098) drill. Clampthe splice members to the damaged former. Centerpunch the required sixteen rivet locations inthe noted pattern. With a No. 30 (.1285) drill,dri I center- punched rivet locations and drillthe splice members through the existing rivetholes in the skin. Observe a minimum edgedistance of 3/8- inch measured from the centerof the rivet hole to the edge of the materialinvolved. Remove the splice members, and burrall the rivet holes. Apply one coat of zincchromate primer to al I surfaces.Again clamp the splice members to the damagedformer. Through the holes in the skin, insertand drive BI227-AD4 rivets in the quantity required.Into the remaining rivet holes, insertand drive sixteen AN442-AD4 rivets (see Figure39^-41. REPLACING ALUMINUM FORMERS.If replacement of a damaged former is warranted,a duplicate spare part should be utilizedfor the replacement. The replacementpart numbers of the front fuselage fairingpanels are shown (see Figure 31). The partRESTRICTED[bi]

adI theIoutRESTRICTEDFUSELAGEPAR. 41 TO 42numbers of the aluminum rear fuselage formersare also shown (see Figure q^f). Where Immediatereplacement is necessary and a spare partis not available^ replacement formers may befabricated locally, using the damaged formeras a template. The procedure may be summed upas fol lows: Lay out the shape of the affectedformer upon a panel of 1/2- inch wood. In severalplaces, temporarily nail this panel toa second panel of 1/2- inch wood. On a bandsaw or with a keyhole saw, cut around the markedcontour. Pull out the nails and separate thetwo equivalent panels cut to the contour ofthe affected former. Cut a sheet of 24ST alcI of the same thickness and developed widthand length as the damaged fonner. If any stringercutouts are present in the damaged fonner, nakecuts in the replacement sheet to match. Clamp thesheet of ale I ad between two wood form blocks ina vise; and with a rawhide mal let, pound theflanges of the replacement former over the edgeof the wood form blocks (see Figure 40)- Removethe former from the vise and make any furthermodifications that may be necessary, suchas fairlead holes, etc. Smear al I overlappingsurfaces of the replacement former with zincchromate primer. Install the replacement formerin exactly the same manner as the original.REPLACE EXISTINGBI227-AD4 RIVETSTHROUGH SKIN..040 X 5-3/4" X2-3/4" SHEETS. 24STALCLAD (2 REQ.)42. ALUMINUM STRINGER TYPE CI07LT-40.The Type CI(j7LT-40 stringer is used in severallocations of the rear fuselage lower coveredframe assembly on the AT-6C Airplanesonly. The stringer is formed of rolled 24STale lad sheet .040 inch thick having no equivalentAlcoa die nuniber. If the damage isslight in nature, a CI07LT-40 splice member6-1/2 inches long will suffice (see Fi^re 41).If the da/nage is extensive, cut out the damagedmaterial with a hack saw, locating the cut midwaybetween two skin rivets at each side ofthe damage. For the spl ice member, cut a lengthof CI07LT-20 stringer equal to 6-1/2 inchesplus the length of the damage. Skin rivetsat the end of the splice must have at least3/8-inch edge distance and this may necessitatemaking the splice member slightly longerthan 6-1/2 inches plus the damage length. Ifthe prepared CI07LT-20 stringer material isnot available, bend up an angle locally from.(J40 inch thick 24ST alclad sheet material toserve as the splice member. Dri I the affectedskin rivets at each side of the damage,and clamp the stringer splice member to thedamaged stringer. With a Mo. 40 (.098) drill,dri I spl ice member through the existingskin rivet holes. If the existing skin rivetsare spaced at 1-1/4 inches on centers, or greater,double the number of existing skin rivet holesat the splice. With a No. 30 (.1285) drill,drill the five holes at 5>/8-inch on centersthrough the stringer upstanding flange and theAN442-AD4 RIVETS(16 REQ.)SKIN(REF.)RAWHIDEMALLETFRAMEFLANGESECTION A A.TRIM SHEETSAROUND EXIST-ING STRINGERCUTOUTSFi^re 39 -^Aluminum Rear Fuselage Former SpliceFigure 40 — Fabr ica t ing Repl acement Formers[62]RESTRICTED

-.I thesurfacesadFUSELAGEPAR. 1^2 TO 45RESTRICTEDspl ice member at each side of the damage. Removethe splice member, and burr the rivet holes.Apply one coat of zinc ch romate primer to a IIoverlapping surfaces of the splice member.Clamp the spl ice member to the damaged stringer.At each side of the damage, drive the fiveAN442-AD4 rivets in the upstanding flange, andrivet the spl ice member through the skin withthe same size and type rivets previously removedfrom the skin (see Figure 41) • :-LOCATE CUT BETWEEN2 SKIN RIVETS ON EACHSIDE OF DAMAGECI07LT40 SPLICEMEMBER-6^ PLUS LENGTH OF DAMAGE-SKIN(REF)AN442-AD4RIVETS(5 REQ. ONEACH SIDE)-EXISTING SKIN RIVETST 5"8JT' .040kfj^Figure 41-'Splice for Stringer Type C107LT-4013. ALUMINUM STRINGER TYPE CI80T,The Type CI SOT stringer is used in the constructionof the fuselage rear section upperdeck assembly. The stringer is a bulb angleof extruded 24ST aluminum alloy material, havingthe Alcoa type designation of KI428(J. Thelongest length of the stringer employed in anylocation is 21 inches; therefore, if damageoccurs to this type stringer, replace the entiredamaged stringer section.44. ALUMINUM STRINGER TYPE C234LT.The stringer Type C234LT is used extensivelyin the construction of the rear aluminum fuselage.The stringer is a Z section of .025inch thick 24ST ale lad, having no equivalentAlcoa die number. If the damage is extensive,cut out the damaged material with a hack saw,locating the cut midway between two skin rivetsat each side of the damage (see Figure 42JSkin rivets at the end of the spl ice must haveat least 3/8- inch edge distance, and this maynecessitate making the splice member slightlylonger than that noted. Clamp the stringersplice member to the damaged stringer. Witha No. 21 (.159) drill, drill five rivet holesat each side of the damage at 5/8- inch on centersafter center punching the hole locations.Dri I spl ice member through the skin withthe same size drill and spacing as the existingskin rivets through the damaged stringer.Remove the splice member, and burr all the rivetholes. Apply one coat of zinc ch romate primerto al I of the splice member. Againclamp the splice member to the stringer. Ateach side of the damage through the stringerupstanding flange, insert and drive five AN442-AD5 rivets. Through the skin and the splicemember, insert and drive BI227-AD rivets ofthe size and quantity required (see Figure 42) •45. ALUMINUM STRINGER TYPE C364LT.The stringer Type C364LT is used in the constructionof the front fuselage fairing sidepanels. The stringer is formed of .020 inchthick 24ST aid ad, having no equivalent Alcoadie number. If the damage is greater than acrack, cut out the damaged material, locatingthe cut between two skin rivets at each sideof the damage. For the splice member, cut asheet of .032 inch thick 24ST ale I 2 incheswide and having z length of 7-1/2 inches plusthe length of the damage (see Figure 44)- Alongthe length of the splice sheet, bend a 1/4-inch, 90-degree flange; and then make a 90-degree, 7/ 16- inch bend measured from the firstflange. Along the opposite edge of the lengthof the spl ice sheet, bend a 90-degree, 1/2- inchflange. Use a hand brake if available andobserve a minimum tend radius of |/8-inch.With a No. 40 (.098) drill, drill out the affectedskin rivets at each side of the damage.Clamp the splice sheet to the damaged stringer,and drill the splice member through the existingskin rivet holes, using a No. 40 (.098)drill. Double the number of skin rivet locationsat the splice location if the existingrivet holes in the skin are spaced greater than1-1/4 inches on centers. In the upstandingflange of the splice member, center punch fiverivet locations at each side of the damage atan average spacing of 3/4-inch on centers.Drill the center-punched rivet locations witha No. 30 (.1285) drill. Remove the splice members,and burr all the rivet holes. Apply onecoat of zinc chromate primer to all overlappingsurfaces of the splice sheet. Again clampthe splice sheet to the damaged stringer.RESTRICTED[63]

I theRESWICTEDReplace the skin rivets with rivets of the samesize and type as previously removed. In theremaining rivet holes, insert and drive tenAN442-AD4 rivets (see Figure ^i/j.^6. ALUMINUM STRINGER TYPE C366T.The stringer Type C366T is used in severallocations in the lower section of the rearfuselage. The stringer is a 24ST aluminumright-angled bulb section having the equivalentAlcoa die number of L249I0. The longestlength of the extruded section employed in theconstruction of the rear fuselage is 23 inches;therefore, if damage occurs to this type stringer,replace the entire damaged section.^7. ALUMINUM REAR FUSELAGE BAGGAGE COMPART-MENT INTERCOSTAL ANGLE.The NA 55-31401 intercostal angle locatedon the left side of the rear fuselage belowthe baggage compartment may either be replacedor spliced if damaged. The intercostal angleextends 51 inches aft from the fuselage monocoouejoining surface and is formed of J-shaped,.081 inch thick, subsequently heat-treated 24S0alclad. The spHice of the member is shown(see Figure ^gj. If the damage is extensive,cut out the damaged material with a hack saw,locating the cut between two skin rivets ateach side of the damage. For the splice merrvber,cut a sheet of .081 inch thick 24S0 alcladhaving a width of 2-5/ If^ inches and havinga length of 9 inches plus the length of thedamage. Along the length of the splice member,bend up a 90-degree, 5/l^inch flange; and thenbend up a 90-degree, 1-3/32 inch flange measuredfrom the first bend. Use a hand brake to bendthe sheet, and observe a minimum bend radius of5/32- inch. After the 24S0 sheet is fonned tothe proper shape, heat treat the member to thefull hardness of 24ST alclad as outlined inSection I. If any rework on the splice sheetis necessary, accomplish the work within thefirst hour after the material is heat treated,as the material may be worked with comparativeease during this period. Within a distanceof 4-1/2 inches on each side of the damage,drill out the existing rivets in the damagedintercostal angle. Clamp the splice memberto the damaged intercostal angle; and with aNo. 30 (. 1285) dri I I, dri spl ice member1through the existing rivet holes in the skin.At each side of the damage along the top flangeof the intercostal angle, center punch sevenrivet locations at 5/e-inch on centers, andobserve a minimum edge, d istance of 3/8-inchFUSELAGEPAR. 45 TO 48measured from the center-punch fnark to edge ofthe material involved. If the damage is locatedin the baggage compartment where the top flangeof the intercostal angle is riveted to theapron assembly of the baggage compartment,double the number of the rivet locations throughthe apron assembly and the intercostal angle atthe splice location. With a No. 21 (.159)drill, drill the fourteen center-punched rivetlocations along the top flange of the inter-K-AArJMMLOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGE3iEXISTING SKIN RIVETSAN442-AD5 RIVETS5 REQ. ON EACH SIDE)C203T SPLICEMEMBEREXISTING SKIN RIVETS7^^SAME SIZE AND SPACINGAS EXISTING SKIN RIVETS-7-^ PLUS LENGTH OF DAMAGE-.025-JrvFigure 42 — Splice for Stringer Type C234LTcostal angle. Rwiove the splice member, andburr all the rivet holes. Apply one coat ofzinc chromate primer to all overlapping surfaces.Again clamp the splice member to thedamaged intercostal angle; and through the skinrivet holes, insert and drive BI227-AD4 rivetsin the quantity required. Through the rivetholes in the top flange of the intercostalangle, insert and drive fourteen AN442-AD5rivets (see Figure 43J . If the damage is locatedin the baggage compartment, rivet thetop flange of the intercostal angle throughthe baggage compartment apron assembly withBI227-AD5 rivets.l^8. ALUMINUM REAR FUSELAGE UPPER LONGERONS.The rear fuselage upper longerons which extend120 inches aft of the fuselage joiningsurface on the left and the right side of thefuselage is formed of a bent-up U-shaped sheetof .064 inch thick 24ST alclad. A damagedlongeron may be replaced, but repair is quite[64]RESTRICTED

angeIIFUSELAGEPARAGRAPH^SRESTRICTEDsimple and is preferable to replacement. Ifthe damage is extensive, cut out the damagedmaterial with a hack saw, locating the cutbetween two skin rivets at each side of thedaniage. For the splice members, cut two sheetsof .064 inch thick 24ST ale lad, each 3 incheswide and having a length of 9 inches plus thelength of the damage (see Figure 46). Alongthe length of one of the splice members, benda 90-degree, I- 1/16 inch flange. Along thelength of the other spl ice member, bend a oneinchflange to match the open angle of thedamaged longeron. Use a hand brake to bendthe splice members, and observe a minimum bendradius of 3/16-inch. With a No. 40 (.098)drill, drill out the affected skin rivets inthe damaged longeron for a distance of 4-1/2inches on each side of the damage. Locate thedrill in the depression of the skin rivet, andapply the drill power in short bursts untilthe rivet head twists from the shank; thenpunch out the rivet. If the damage is locatedon the right side of the fuselage between thefirst and second formers aft of the fuselagejoining surface, drill out the affected rivetssecuring the gunner's hood cowling attachingnut plates to the skin side flange of the longeron.If the damage is located on the leftI I, drill the spl ice member f Iside of the fuselage between the first andsecond formers aft of the fuselage joining surface,drill out the cowling nut plates; and inaddition, drill out the rivets securing thebaggage compartment hinge to the skin sideflange of the longeron. Clamp the splice membersto the proper positions on the damagedlongeron (see Figure 46). With a No. 30 (.1285)d ri throughthe existing skin rivet, holes. If the damageis located between the first and second formersaft of the fuselage joining surface, drillthe one splice member through the existingskin rivet holes with the same size drill asthe existing holes. Center punch the 36 splicerivet locations at an average spacing of 3/4- inchon centers observing a minimum edge distanceof 3/8- inch measured from the center-punch markto the edge of the material involved (see Figure.46) With a No. 21 ( . 159) dri I ,dri Ithe 36 center-punched rivet locations. Removethe splice members, and burr all rivetholes. Apply one coat of zinc chromate to alloverlapping surfaces of the splice members.Again clamp the splice members to the properpositions on the damaged longeron. Throughthe skin rivet holes, insert and drive BI227-AD4 rivets in the quantity required. If theREPLACE EXISTING SKIN RIVETSAT SPLICE WITH BI227-AD4 RIVETSO 00000FILLERSKIN (REF.)^AAN442-AD5 RIVETS(14 REQ.)rOP VIEWTRIM DAMAGE9" PLUS LENGTH OF DAMAGE08IX2-5/l6"X9'+SPLICESHEET, 2430 ALCLADHEAT TREAT AFTERFORMING5/32R ^SECTIONA AFigure 43— Aluminum Rear Fuselage Baggage Compartment Intercostal Angle SpliceRESTRICTED[65]

ongeron,IIImarkI intoI powerRESTRICTEDFUSELAGEPAR. 48 TO 49SKIN(REF)REPLACE EXISTINGLOCATE CUT BETWEEN SKIN RIVETS THROUGH2 SKIN RIVETS ON EACH SPLICE SHEETSIDE OF DAMAGEAN442-AD4H i' \* RIVETS (5 REGON EACH SIDE)L HCAMERA DOOR032 X 2"X 7-1/2+ SPLICE SHEET24ST ALCLADC364LTFigure 44— Splice for Stringer Type C364LTdamage is located on the right side betweenthe first and second frames aft of the fuselagejoining surface, rerivet the existingnut plates to the skin side flange of the longeronwith AN426-AD3 rivets. If the damageis located on the left side of the fuselagebetween the first and second frames aft of thejoining surface, in addition to rerivetingthe nut plates, rerivet the baggage compartmenthinge to the longeron with AN426-AD4 rivetsin the quantity required. Into the remaining36 rivet holes through the splice members andthe damaged longeron, insert and drive AN442-AD5 rivets (see Figure 46).49. ALUMINUM REAR FUSELAGE LOWER LONGERONS.IIThe lower longerons extending inches aftof the fuselage joining surface on the leftand right side of the rear fuselage may bespliced as shown (see Figure 47J . if the damageis extensive, cut out the damaged materialwith a hack saw, locating the cut between theskin rivets at each side of the damage. Onthe upper flange of the damaged longeron, dri Iout the skin rivets for a distance of 3-3/4inches on each side of the damage and removeany existing former attaching angles. On thelower flange of the damaged d r i Mout the skin rivets along the entire lengthof affected skin panel and bend the skin panelupwards to gain access to the inside of thelongeron at the affected area. Use a No. 40Figure 45— Aluminum Rear Fuselage Doors( .09P ) drill and place the dri I depressionin the rivet head; apply the dri I inshort bursts until the rivet head twists fromthe shank; then punch out the remaining portionof the rivet. Use this rrethod to removeall skin rivets, and take particularcare toprevent the elongation of any of the existingrivet holes. For the splice member, cut a sheetof .051 inch thick 24ST alclad having a widthof 7 inches and having a length of 7-1/2 inchesplus the length of the damage. Along the lengthof the splice sheet, bend a 5/P-inch flangeto an angle matching the upper longeron angle.Apply the splice member to the top surface ofthe damaged longeron so that the bent-up spl icemember flange fits the inside radius of thelongeron flange. At the intersection of thetop and side surfaces of the longeron, makea grease penci I along the length of thesplice member. The distance between the bendof the first 5/P-inch flange and the markedline should be between 2-3/16 and 2-3/4 inches.Along this line brake a QO-degree flange. Measurethe distance on the damaged longeron betweenthe intersection of the top and sideof the longeron to the beginning of the bendin the lower flange of the longeron, Markthis distance on the splice member and on ahand brake; bend the splice member flangealong this line to match the bottom flangeon the longeron. Trim the excess materialfrom the bottom flange of the splice member,so that the bottom flange has a constant[66]RESTRICTED

6-twoI , RemoveIFUSELAGEPARAGRAPH 49RESTRICTEDFigure 46— Aluminum Rear Fuselage Upper Longeron Splicewidth of 5/8-inch. In all of the bending operationson the splice member, observe a minimumbend radius of 3/ 1inch. Apply the splicemember to the damaged longeron so that thesplice member overlaps the damage on each sideby 5-3/4 inches. At each side of the damage,on the top and side surface of the spl ice member,lay out with a penci I rows of rivetlocati.ons in the noted pattern (see Figure.4J) Center punch the 40 rivet locations; andwith a No. 21 (.159) drill, drill the splicemember through at opposite ends and fasten themember in place with skin fasteners. Dri Ithe remaining center-punched rivet locationswith the No. 2! ( . 159) dri I the spl icemember, and burr all the rivet holes. Applyone coat of zinc chromate primer to a M overlappingsurfaces of the splice member. Againsecure the spl ice member to the longeron withskin fasteners. Through the 40 rivet holesin the splice member, insert and drive AN442-AD5 rivets. To provide access for bucking therivets on the inside of the longeron, bend theskin panel upward at the lower flange of thelongeron, the securing skin rivets having beenpreviously drilled out. Rerivet to the longeronany previously removed former attachingangles. On the entire length of the skinedge at the lower flange of the longeron, replacethe previously drilled-out skin rivetswith BI227-AD4 rivets except in the immediatelongeron repair area. With a No. 30 (.1285)drill, drill the upper and lower spl ice mem -ber flanges through the existing skin rivethole locations- Using a thin hook-shaped hacksaw blade with the teeth removed, scrape outthe drill chips between the sheets at the skinrivet locations. At the splice location, replaceRESTRICTED[67]

.lyreI letsshapes.RESTRICTEDFUSELAGEPAR. il9 TO 52BI227-AD4 skin rivets through the splice member,bucking the rivets from the interior ofthe fuselage (see Figure 47J.50. ALUMINUM SKIN AND BULKHEAD WEBS.The skin arrangement of the fuselage consistspri nci pa I of 24ST ale lad aluminum a Moy (seeFigure 48). The bulkhead webs of the rearfuselage are also formed of thin 24ST ale lad.For the repair of the skin and the bulkheadwebs, see the pertinent paragraphs applicableto skin in Section III. Inasmuch as accessto the most of the fuselage interior may bereadily gained, all fuselage skin patches shouldbe applied to the interior. On the inner orouter skin of the baggage compartment door andcamera door (see Figure 45J , any patching maybe accomplished with LSI 127-4-2 Cherry blindrivets. The double skin construction of thesedoors provides no interior access for rivetbuck i ng51. FAIRINGS AND FILLETS.On the earlier AT-6C Series Airplanes andprevious North American Aviation, Inc., trainera i raft, the fa i ri ngs and f i a re madefrom 3S aluminum alloy, 52S aluminum alloy, andsubsequently heat-treated 24S0 ale lad (seeFigure 48). For the repair of the a I clad fairingsand fillets, see the pertinent paragraphson skin repa i rs in Sect ion III. Patch i ng 5Sand 52S aluminum alloy may be accomplished bywelding equivalent rmterial over the hole, usinga torch and a 2S aluminum ai loy welding wire.On later AT-6C Series Airplanes, some of thealuminum fairings and fillets are replaced withplastic fillets of phenolic sheet and Chemold.These plastic fairings and fillets are consideredexpendables; and as such, they shouldbe replaced rather than repaired (see Figure ^g)52. EXTRUDED ALUMINUM REPAIR MATERIAL.The extruded c ross-sect iona I that arerequired in the repair of the structural membersof the fuselage are listed below. Thematerial used in forming these extrusions is24ST aluminum alloy conforming to Federal Specificat ion -QQ-A-354. Numbers with a C prefixare North American Aviation, Inc., standardparts. Following the North American Aviation,Inc., part number, the A luminum Company-3-3/

FUSELAGEFIGURE 1+8RESTRICTED.032 .025W4JUZCJDDHL-^051-.032 .025TOP OF FUSELAGE".040SIDE OF FUSELAGEBOTTOM OF FUSELAGEi^--CODE. 040 3S AL ALLOY. 051 52S AL. ALLOY{I - . 036 CORR. RESIST. STEELIZa - . 032 52SAL. ALLOYIZZ - 025 CORR. RESIST STEELNOTE.ALL SKIN IS 24 ST ALCLADEXCEPT AS NOTED ANDDECIMAL FRACTIONS INDICATESKIN THICKNESS IN INCHESFigure 48— Fuselage Skin ArrangementRESTRICTED[69]

RESTRICTEDFUSELAGEPAR. 52 TO 54PHENOLIC SHEET88-3 1 03UCHEMOLDPLASTICFigure 49— Plastic Fairings and Filletsof America (Alcoa) equivalent die number islisted in parentheses.EXTRUSION DIE NUMBERSC180TC366TK1U280)L24910)53. ALCLAD SHEET REPAIR MATERIAL.The aluminum sheet material used in the repairof the fuselage should be 24ST ale lad conformingto Federal Specification QQ-A-362 .The various thicknesses of sheets that may berequired are listed below. The lengths andwidths of the sheets must be determined locallyfor the extent of repairs to be undertaken,with reference to the repair procedures out-lined in the appropriate paragraphs of this Section.THICKNESS OF 2i+STALCLAD (INCHES) REMARKSI NTERCOSTAL.020 SKIN. BULKHEADS (C36ULT).025 SKIN, BULKHEADS (C23nLT).032 SKIN, BULKHEADS, FORMERS.040 SKIN, FORMERS, (C107LT-U0).05 1 SK IN , LONGERONS.064 LONGERONS.08154. RIVETS REQUIRED FOR ALUMINUM REPAIRS.The fo Mowing types of rivets may be requiredfor the repair of the fuselage. Typesof rivets are described in Section I.[70]RESTRICTED

FUSELAGEPAR. 5^; JO 37iCRESTRICTEDSPRUCESTRINGERSMAHOGANY3-PLYSKIN AND WEBSJKLXmmmFigure50 — Interior View of Wooden Fusel agePART NO.

RESTRICTEDFUSELAGEFIGURE 5252-31109 LH52-3II09-I RH88-31532 LH88-31532-1 RHSTA.149^STA. I06|^NOTES^ALL WEB MEMBERS AND SKIN PANELS AREMADE FROM 3-PLY MAHOGANY PLYWOOD CON-FORMING TO AAF SPEC. AN-NN-P5IIA TYPE IIALL STRINGERS AND FRAME FLANGES MADEFROM SPRUCE CONFORMING TO AAF SPEC. AN-S-6AFOR ASSEMBLY CASEIN GLUE (AAF SPEC. C-G-456)ISUSED.ONLY FITTINGSABLE ITEMS)ARE NOTED (THE ONLY REPLACE-DETAIL AREF. DWG. 88-31106Figure 52 — Wooden Rear Fuselage Frame Assembly[72]RESTRICTED

I loseFUSELAGEPAR. 57 TO 58RESTRICTEDstringer cutouts in the frames are reinforcedwith wooden gussets. The rectangular sprucestringers pass through the frames and are rigidlyattached to the frames by the use of smal I,thin, tapered spruce wedges secured with caseinglue. An intercostal located at the baggagecompartment cutout is used to take directbending loads, because of the shear transferaround the cutout. The two upper and two lowerlongerons support the entire fuselage bendingloads, neglecting the effect of the continuousportions of the fuselage skin. The upper longeronis formed of rectangular spruce tapering towardthe rear bulkhead. The lower longeron is madefrom rectangular spruce, several tricleats,and a plywood web, forming a closed triangularbox sect ion.REARFUSELAGEPLYWOODFRAME58. REPAIR WOOD MOISTURE CONTROL.For obvious reasons, as much care and thoroughnessmust be taken with the repair of woodenaircraft parts as is taken with the originalconstruction of the parts. For that reason,care should be taken to make certain that themoisture content of new or spliced parts is approximatelythe same as that in the originalpart. Wood gives off or takes on moisture fromthe surrounding air. Like many other hygroscopicmaterials, wood shrinks as it loses moistureand swells as it absorbs moisture. It isof the utmost importance that particular atten-Figure 53~-Interior View of Fuselage ShowingSkin Grain Direct ionLOWERLONGERONOF REARFUSELAGEf-PLYWOOD SKINFRONT FUSELAGETRUSS JOINTFigure 54^Attachwent of Lower SteelLongeron to Rear Wooden Fuselaget i on be paid to the moisture content of thepart or parts to be repaired and the new woodto be used in the repair. If the part to berepaired has a moisture content of 10 percentand a piece of wood is spliced to it with amoisture content of 15 percent, when the finishedassembly is out in service, both partswi I or absorb moisture to the extent offinally having the same moisture content, dependingupon the relative humidity and temperature.The sealer coats applied to the finalwood assembly serve only to retard any increaseor decrease in moisture content. Therefore,if the part having 15 percent moisturecontent goes down to 12 percent, it wi I I shrink;and if the part having 10 percent goes up to12 percent, it wi M swe I I, causing a strainin the glue Joint. In general, if the woodis selected, prepared, and packaged in accordancewith the applicable material specificationfor that particular type of wood, noconsideration need be given to the moisturecontent of the wood, for it is necessari ly thesame as that used in original construction.If, however, there is any reason to doubt thecorrectness of the specified moisture content,RESTRICTED[73]

PARTimitsRESTRICTEDFUSELAGEPAR. 58 TO 59because of storing the exposed wood for a considerableperiod, the moisture content of thewood should be checked and corrected beforeusing the wood. The correct method of determiningmoisture content is to base the determinationon oven-dry weight. The equipmentnecessary to determine moisture content consistsof a drying oven and balance scale. Thedrying oven must be capable of maintaining aconstant temperature of from 100° to I05°C(212° to 221 °F). The balance scale must be accurateto within one half of one percent. Thetest specimen should be about one inch widein the direction of the grain, about two incheslong across the grain, and of the full depthof the material. The specimen should be takenfar enough from the end to ensure that the rroisturecontent has not been affected by end drying.All loose splinters must be removed andthe test specimen weighed immediately aftersawing, to avoid any moisture change resultingfrom exposure to air. The specimen should beplaced in a drying oven and dried at 100° toI05°C (212° to 22I°F) to constant weight. Thiswill occur after about 24 hours. Immediatelyafter removal from the oven, the specimen shouldbe weighed. The formula for determining moisturecontent is as follows:W = Original weight (before drying)D = Oven-dry weightMIX 2 PARTSWATER BY WEIGHTTO I CASEIN »k-GLUE POWDER(SPEC. C-G-456)NOTE:USE THE GLUEIMMEDIATELYAFTER MIXING''^Figure 55— Mixing Casein GluePORCELAINOR ENAMEL:i CONTAINERThe percentage of moisture content = 'JjLy' y |qqIf the moisture content of the wood is abovethe al lowable I specified in the appi i-cable material specification for the particulartype of wood, kiln-dry the wood-stock tothe proper moisture content as outlined inSpec. AN-W-2 before the stock is used for repair.A variation of 2 percent in moisturecontent is the maximum permissible between theoriginal structure and the patch or spliced replacement.59. SUBSTITUTES FOR SPECIFIED WOODS.In the repairs outlined in this Section, threeplymahogany poplar core, plywood (Spec. AN-NN-P5IIA , and spruce (Spec. AN-S-6) are the on 'ywoods specified. In the event that these woodsare not available, the following substitutesmay be used. Any other substitutes are notacceptable and cannot be recommended. The followingplywoods may be substituted for threeplymahogany poplar core plywood and shall conformto AC Spec. AN-NN-P5IIA: (I) three-plyDouglas fir plywood, poplar core; (2) threeplyDouglas fir plywood, Douglas fir core;(3) three-ply red gum plywood, red gum core.The following woods may be substituted forspruce: (I) yellow poplar. Spec. AN-P-17; (2)western hemlock, Spec. AN-H-4; (3) noble fir.WOODENFRAME ATSTA 106-3'Figure 56 —Attachment of Upper SteelLongeron to Rear Wooden FuselageLv4]RESTRICTED

containersI containerstimesFUSELAGEPAR. 59 TO 62RESTRICTEDWITH A BRUSH, APPLY CASEINGLUE (SPEC. C-G-456) TO BOTfSURFACES OF THE JOINT.ASSEMBLE JOINT S CLAMPSECURELYFigure 57-^Applying Casein Glue With BrushSpec. AN-F-6. In the case of a replacementonly, no substitute for compregwood Is permissible.Compregwood is defined as impregnatedand compressed laminated maple having adensity of from 80 to 85 pounds per cubic foot.Compregwood may be purchased from Formica InsulatingCorp., Cincinnati, Ohio, or equiva-I ent.60. CASEIN GLUE FOR WOOD REPAIRS.Casein glue is the only type glue used in theoriginal construction of the wooden structures,and hence the only glue recoirmiended for repairs.Any commercial casein glue conforming to FederalSpec. C-G-456 may be used. Casein, whichis the chief protein constituent of milk, isthe chief ingredient. The other ingredientsare hyd rated I ime, sodium hydroxide, and water.JERGENSONPRESSUREBLOCKS {2 RlCLAMP(3REQTO GLUED SCARF JOINT,APPLY PRESSURE OF APPROX150 LBS. PER SQ. IN. FOR 4 HRS.Figure SS-'Correct Equalized Application ofGluing PressureThis glue can be obtained in a powder form andmixed with water immediately prior to use.Casein glue is water-resistant but not waterproof,and is subject to mold growth.61. MIXING CASEIN GLUE FOR WOOD REPAIRS.IAll casein glue is obtained in the powderedform and mixed immediately prior to use. Allproportions of glue and water used in mixingmust be determined by weight and not by volume.The proportion of water and glue powder formixing must conform to the directions whichwill be found on each container of certifiedglue, or which will be furnished by the manufacturerof the glue. In lieu of directions,a proportion of two parts of water by weightto one part of glue by weight is recommendedfor softwoods, such as spruce. For hardwoods,such as mahogany, use 1.8 parts of water byweight to one part glue powder by weight. Thewater should be at room temperature (21 °C, 70°F).The use of a mechanical power-driven mixer isrecommended to ensure full strength, the mixingto continue for approximately 5 minutesat a high speed of approximately 140 RPM. Stopthe mixer when the glue thickens, and do notadd water. Allow the mixture to stand forone-half hour for chemical digestion. Againoperate the mixer for approximately 10 minutesat a slow speed of from 60 to 90 RPM, orfor such additional time as required to dissolveall casein particles. The mixture mustbe of smooth consistency and free of air bubbles,foam, and lumps. If power-driven equipmentis not available, hand methods may besubstituted if the above procedure is followedas closely as possible. Ordinary roomtemperature and humidity are satisfactoryfor the mixing room. The glue must not bemixed in containers made of aluminum, brass,copper, or zinc. Smal I made of porcelainor enamelware are preferred (see Figure55^'/ larger containers or mixing bowls shallbe of iron or steel. A I and brushesmust be kept thoroughly clean at al andmust contain none of the remaining glue of thepreceding batch. The glue must be used within4 hours after mixing, for the glue jellsafter a longer period and cannot be used. Donot heat the glue or the wood. The mixing andapplication of casein glue is strictly a coldprocess.62. ASSEMBLING GLUED JOINTS OF WOOD.Apply a sufficient quantity of glue to alljoints to penetrate into the pores of the wood, andRESTRICTEDL75]

areas,I.RESTRICTEDcoat all parts of the joint. The glue and thev\ooci shall be at room temperature, and the surfacesof the wood for gluing shall be smoothand cleaned of all greases, varnish, or othercoating materials. Tooth-planing or scratchingis not recommended. However, when gluingcompregwood, the surface should be roughenedbefore the application of glue. The glue shallnot be spread thin or brushed out in thin layers.A spread of approximately 1-1/2 ounces of wetglue per square foot of single glue line issatisfactory. Where practicable and where repai r cond it ions will permit, the g lue shou Idbe applied to both surfaces. A mechanicalspreader is preferred; but as most repairs wi Icover only smal I a brush may be used(see Figure 57J . Brushes must be of good qualityto prevent loose or broken bristles becomingmixed with the film of glue. In most cases,clamps should be applied to all joints immediatelyafter the glue is applied, and should beadjusted to give a uniform pressure on al I partsof the joint. Any type of clamp may be usedthat will give satisfactory results and thatwill exert a pressure of approximately 150Ibs./sq.in. for softwood and 200 Ibs./sq.In.for hardwood (see Figure ^8). The clamps mustnot be removed from the joints within 4 hoursafter clamping, and a longer period is preferredand required where joints cover an extendedarea or where it is necessary to use large quantitiesof glue in proportion to the volume ofwood in the member. Handwork on glued partsmay be done as soon as they are removed fromthe clamps, but machining should not be attempteduntil after 12 hours. When gluing compregwoodor wood of extremely fine-grained nature,it Is often desirable to coat both surfaceswith an even coat of glue and allow it to setuntil It reaches a state of tackiness. Applya final coat of glue and clamp the joint immedlate ly63. APPLICATION OF WOOD-GLUING PRESSURE WITHCLAMPS.FUSELAGEPAR. 62 TO 64In most cases, pressure for gluing may be applledwith clamps, but care should be taken toprovide uniform pressure (see Figure ^8). TheImportance of properly applying gluing pressurecannot be stressed too strongly. Uniform pressuremay be provided by first applying pressureblocks to both sides of the joint and thenclamping the area (see Figure ^8). Use a sufficientnumber of clamps to provide a uniformpressure; otherwise a poor joint will result(see Figure 50 j. Various types of clamps areavailable, and each has its- own part Icular advantage.The most cprrmonly used wood clamp Is theJorgensen clamp (see Figure 61). The Jorgensenclamp has two wooden jaws equipped with two handscrews for exerting pressure. Another c lamp, thebar clamp, is made up of a bar about 2 incheswide and from 18 inches long to 6 feet long, dependingupon the type (see Figure 61). The "C"clamp is aval lable in various sizes and is probablythe most useful type of clamp for repairwork (see Figure 61). Any type of clamp rmy beused that will give satisfactory results andthat will exert a pressure of approximately 150Figure 59^— Incorrect Concentrated Applicationof Gluing Pressuretbs./sq.in. for softwood and 200 Ibs./sq.in. forhardwood. The clamps must not be remcved from thejoints within 4 hours after clamping, and a longerperiod is preferred and required where jointscover an extended area.64. APPLICATION OF WOOD-GLUING PRESSURE WITHNAILING STRIPS.In assembling the skin to the interior structure,and in other instances where the applicationof clamps is impossible, pressure[76]RESTRICTED

ingingkeyhole.FUSELAGEPAR. 64 TO 65required for gluing can be obtained by temporarilynai I st ri ps of vjoo6 over the gl ued area.These nai I strips are general ly made of basswoodor soft pine; however, spruce or any otheravailable wood can be used. The nailing stripsare cut 1/4-inch thick and 3/8-inch to 1/2-inchwide depending upon the use. The size of nai Isused varies according to the thickness andwidth of the material to be glued. For securing1/16-inch to 3/32-inch thick plywood skinto stringers while the glue is hardening, aspacing of 1-1/2 inches apart, using a 5/P-inchNo. 20 gage nail, is recomnended (see Figure 6o)Sufficient nai ling strips must be used to coverentirely the scarf joint area. For example,if the width of the scarf joint is 1-1/2 inches,three 1/2-inch wide nailing strips must beused to provide pressure. Place the nailingstrips side by side and alternate the nailpositions in the adjacent strips. The nailingstrips should be secured first to the centerof the scarf joint length and then thenailing should progress outward from the center.Drive the nai Is down as tightly as possible,in order to provide adequate gluing pressure.After the nai ling strips are secured, the glued/DRIVE NO 20 GAGE NAILSTHROUGH NAILING STRIP,SKIN AND INTERIOR FRAMESTRUCTURE^^OUTSIDE SKINSURFACE (REFSOFT PINE NAILING STRIP,'/4 X 1/2 X REQUIRED LENGTHREMOVE AFTER GLUE BETWEENSKIN AND INTERIOR STRUCTUREHARDENSFigure 60— Using Nailing Strips to ApplyGluing Pressure to SkinBARRESTRICTEDCLAMPJERGENSONBAR CLAMP '-C CLAMP CLAMPFigure 61 —-Types of Clamps for Woodskin portion must be free from wrinkles, bulges,and joint gaps. To prevent glue from seepingfrom the joint and sticking to the nailing strip,waxed paper may be placed under the nailingstrips. After the glue has hardened, removethe nails and the nailing strips.65. WOODEN REAR FUSELAGE STRINGERS.The wooden rear fuselage stringers are formedof rectangular spruce having a thickness of1/2-inch and a width of 5/8-inch. Damagedstringers may be repaired by splicing in a newstringer section as shown (see Figure 62). Ifthe damage is extensive, at each side of thedamage, diagonally cut out the stringer sectionwith a smal I saw (see Figure 76JMake these cuts at a ten-to-one taper whichamounts to a 5-inch taper on this particulartype of stringer (see Figure 62). Plane thestringer material down to the skin. With achisel, pare off any small remaining portionsof the damaged stringer adjacent to the scarfwhich may be inaccessible for the manipulationof the wood plane (see Figure 63 j. Duringthese cutting operations, take particularcare to prevent damage to the skin. Cut a replacementstringer section of spruce (Spec.AN-S-6 ) slightly longer than the removed portionof the original stringer and having crosssectionaldimensions of 1/2- by 5/8-inch. Scarfcutthe ends of the replacement stringer tomatch the cuts on the original stringer. Carefully plane down the saw-cut ends of the scarfcuts on the original stringer and the replacement(see Figure 6^) so that they are perfectlysmooth when measured with a straightedge(see Figure 67 j, and so that they form a perfectfit when assembled. Do not use sandpaper.RESTRICTED[77]

8-inchIRESTRICTEDFUSELAGEPARAGRAPH 65SPRUCE STRINGER -REPLACEMENT 1/2 x 5/8 xLENGTH OF REPLACEMENTTAPER CUT DAMAGED STRINGERAND SECURE REPLACEMENTSTRINGER WITH CASEIN GLUETOP VIEW7T3 PLY MAHOGANYPLYWOOD SKIN REF.SPRUCE OR 3 PLY MAHOGANY PLYWOOD REINFORCE-MENT l/8"x 5/8" X 6" (2REQ.) SECURE TO STRINGERWITH CASEIN GLUE. CLAMP JOINT WITH UNIFORMPRESSURE OF 150 LBS. PER SQ. IN. FOR AT LEAST 4HOURS.BEVEL 45 (4 PLACES)DIRECTION OF SPRUCE ORSECURE STRINGER REPLACEMENTOUTER FACE PLYWOOD GRAIN— TO SKIN WITH CASEIN GLUE.SIDE VIEWFigure 62^Wooden Rear Fuselage Stringer SpliceI theIon surfaces which are to be glued, since thewood dust particles prevent proper gluing. Toserve as the reinforcements for each of thescarf joints, cut two spruce sheets (Spec. AN-S-6)or two sheets of three-ply mahogany poplarcore plywood (Spec. AN-NN-P5I lA), each 1/8-inch thick 5/8-inch wide and 6 inches in length(see Figure 62). Clean all wood chips fromthe surfaces to be glued. With a brush, applyone coat of casein glue (Spec. C-G-456)to all joint surfaces and to the skin and replacementstringer contact surfaces. Assemblethe scarf Joint as shown (see Figure 62) andimmediately apply a series of clamps to thejoint area to provide a uniform pressure of approximately150 Ibs./sq.in. To provide gluingpressure along the skin for the stringer replacement,cut a strip of soft pine l/4--inchthick and 1/2— inch wide and having a lengthequal to the length of the stringer replacement.On the outside skin surface, na isoft-pine strip to the stringer through theskin, using 5/8-inch long No. 20 gage nails at1-1/2 inch spacing (see Figure 60). Allcw theglue to remain under pressure for at least 4hours; then remove the clamps and nailing strip.Carefully scrape off all excess glue aroundthe joints. To the entire added wood material,apply one coat of sealer (AAF Spec. 141 13) andallow to dry for at least one hour. Apply asecond coat of sealer (AAF Spec. 14113) to a Iadded material except mahogany if used, and tothe end-grain surfaces of the f /8-inch reinforcement.To mahogany and the end-gra in surfacesof the reinforcement and to thenail holes in the outside skin, rub on one coatof unthinned filler ( NAA Formula No. 1398 orler to remain onequivalent ), a II owing the f i Ithe surface from 2 to 5 minutes; then removethe excess f i I ler with exce Isior. Allow fillerto dry for at least 2 hours; then to all addedmaterial except the end-grain surfaces, applyone finish coat of either cellulose nitratelacquer (Spec. AN-TT-L-51 ) or aircraft enamel(Spec. 3-98) as required to match adjacentsurfaces. Apply finish tape to the end-grainSCARF CUTON ORIGINALSTRINGER 4CAREFULLY PARE OFFDAMAGED STRINGER WOODADJACENT TO SCARF CUTFigure 63— Removing Portion of DawagedWood Stringer Adjacent to Scarf CutTL78]RESTRICTED

.FUSELAGEPAR . 65 TO 66RESTRICTED6-1/4 SPRUCE INTERCOSTAL REPLACE-MENT 5/8" X 3/4"X LENGTH OFREPLACEMENT —TAPER CUT DAMAGED INTERCOSTAL ANDSECURE REPLACEMENT WITH CASEIN GLUE(REF.)5/32•Vl^•|JU-5/32" ^^^ ^^^1^8(REF) 3/4£/VDVIEWSPRUCE OR 3 PLY MAHOGANY PLYWOOD REINFORCEMENT5/32" X 3/4" X 7-1/2" (2 REQ.) SECURE TO INTERCOSTALWITH CASEIN GLUE CLAMP JOINT WITH UNIFORM PRES-SURE OF 150*/ SO IN. FOR AT LEAST 4 HOURS.BEVEL 45° (4 PLACES)^DIRECTION OF SPRUCE GRAIN OROUTER FACE PLYWOOD GRAINFigure 64— Baggage Compartment Wooden Intercostal Splicesurface as outlined in the paragraph applicableto finishes for wood.66. WOODEN REAR FUSELAGE BAGGAGE COMPARTMENTINTERCOSTAL.The NA 88-3(108-29 intercostal located on theleft side of the rear fuselage belcw the baggagecompartment may be repaired by splicing in anew intercostal angle (see Figure ^4). Theintercostal extends 51 inches aft from the fuselagemonocoque joining surface and is formedof rectangular spruce having a thickness of5/8-inch and a width of 3/4-inch. If the damageSCARF GUT MUST BESMOOTH AND A PERFECTFIT WHEN ASSEMBLEDFigure 65- -Smoothing Scarf Cut With a Wood Planeis extensive, make a diagonal cut with a keyholesaw on one side of the damage so that damagedsection is cut free from the longest lengthof undamaged intercostal. N4ake this cut ata ten-to-one taper, which amounts to a 6-1/4inch taper on this particular type of intercostal(see Figure 64). Carefully plane offthe shorter damaged length of the intercostaluntil the skin is bared. With a chisel, pareoff any small rermining portions of the damagedintercostal adjacent to the scarf which maybe inaccessible for the manipulation of thewood plane. During these cutting operations,take particular care to prevent damage to theskin. Cut a replacement intercostal sectionof spruce (Spec. AN-S-6) slightly longer thanthe removed portion of the original intercostaland having cross-sect iona I dimensions of 1/2-by 5/8-inch. Scarf-cut the end of the replacementintercostal to match the cut end of theoriginal intercostal. Carefully plane downthe ends of the scarf cuts (see Figure 6^)so that they are perfectly smooth when measuredwith a straightedge, and so that they forma perfect fit when assembled (see Figure 67JDo not use sandpaper on surfaces which are tobe glued, since the wood dust particles preventproper gluing. Cut two spruce sheets(Spec. AN-S-6) or two sheets of three-ply mahoganypoplar core plywood (Spec. AN-NN-P5IIA)each 5/32-inch thick, 3/4-inch wide, and 7-1/2inches in length (see Figure 64). Clean allwood chips from the surfaces to be glued. Witha brush, apply one coat of casein glue (Spec.C-G-456 ) to all joint surfaces and to the skinand replacement intercostal contact surfaces.Assemble the scarf joint as shown (see Figure64) and immediately apply a series of clampsto the joint area to provide a uniform pressureRESTRICTED[79]

lowRESTRICTEDFUSELAGEPAR. 66 TO 67SKIN(REF)SPRUCE OR 3 PLY MAHOGANY PLYWOOD REINFORCEMENT3y{6 X 9 " X WIDTH OF LONGERON ( 2 REQ.). SECURE TOLONGERON WITH CASEIN GLUE. CLAMP JOINT WITHUNIFORM PRESSURE OF 150 LBS. PER SO.IN. FORATLEAST 4 HOURS.SPRUCE LONGERONREPLACEMENT SECTIONFigure 66— Wooden Rear Fuselage Upper Longeron Spliceof approxi irately 150 Ibs./sq.in. (see Figure ^8).To provide gluing pressure along the skin forthe intercostal replacement, cut a strip ofsoft pine 1/4-inch thick and 1/2-inch wideandhaving a length equal to the length of the intercostalreplacement. On the outside skinsurface, nail the soft-pine strip to the strin^rthrough the skin, using 5/8- inch long No. 20gage nails at I- 1/2 inch spacing (see Figure6o) . Allow the glue to remain under pressurefor at least 4 hours; then remove the clampsand nailing strip. Carefu My scrape of f a IIexcess glue around the joints. To the entireadded wood material, apply one coat of sealer(AAF Spec. 141 13 ) and a I to dry for at leastone hour. Apply a second coat of sealer (AAFSpec. 14II3) to all added material except tomahogany if used, and to the end-grain surfacesof the 5/32-inch reinforcements. Tomahogany and the end-grain surfaces of the 5/32inch reinforcements and to the nail holes inthe outside skin, rub on one coat of unthinnedf i I ler ( NAA Formula No. 1398 or equivalent ),al lowing the f i I ler to remain on the surfacefrom 2 to 5 minutes; then remove the excessf i I ler with excels ior. Allow the filler todry for at least 2 hours; then to all addedmaterial except the end-grain surfaces, applyone finish coat of either cellulose nitratelacquer (Spec. AN-TT-L-51) or aircraft enamel(Spec. 3-98) as required to match adjacent surfaces.Apply finish tape to the exposed endgrainsurfaces as outlined in the paragraphapplicable to finishes for wood.67. WOODEN REAR FUSELAGE UPPER LONGERONS.The wooden rear fuselage upper longeron ( NA88-31552) which extends 120 inches aft of theAPPLY STRAIGHTEDGE TO SCARFCUT TAPERSCARF CUT TAPERMUST BE SMOOTH ftPERFECTLY STRAIGHTFigure 67—Checking Scarf Cut With Straightedge[80]RESTRICTED

emainingFUSELAGEPARAGRAPH 67EESTRICTEDFigure 68 — Cutting Wood Scarf Jointfuselage Joining surface on the left and theright side of the fuselage is formed of rectangularspruce having a thickness of 3/4-inchand a width which tapers from 2-l/P inches atthe forward position to a 3/4-inch width atthe aft location. At the forward fuselagejoining surface, the longeron is spliced to acompregwood section. The complete replacementof a damaged longeron is not recommended,inasmuch as splicing and partial replacementmay be accomplished with far less work. No repairsto the upper longeron are permissiblewithin the first 40 inches aft of the fuselagejoining surfaces. If damage occurs in thisarea, replace the forward longeron sectionand locate the splice aft of the fuselagejoining surface. If the damage is extensive,cut out the damaged material with a keyholesaw. Make the cut at a ten-to-one taper, whichamounts to a 7- 1/2 inch taper on this particulartype of longeron (see Figure 66). Carefully plane off the damaged length of the longeronuntil the skin is bared. With a chisel,pare off any smal I portions of thedamaged longeron adjacent to the scarf whichmay be inaccessible for the manipulation ofthe wood plane. During these cutting operations,take particular care to prevent damageto the skin. Cut a replacement longeron sectionof spruce (Spec. AN-S-6) slightly longer thanthe removed portion of the original longeronand having maximum cross-sectional dimensionsof 5/4- X 2-1/8 inches (see Figure 66). Scarf-BOTTOMFigure 69— Wooden Rear Fuselage Skin Arrangementcut the ends of the replacement longeron tomatch the cut end of the original longeron.Carefully plane down the ends of the scarfcuts so that they are perfectly smooth whenmeasured with a straightedge (see Figure 61),and so that they form a perfect fit when assembled.Do not use sandpaper on surfaceswhich are to be glued, since the wood dustparticles prevent proper gluing. Cut two sprucesheets (Spec. AN-S-6) or two sheets of threeplymahogany poplar core plywood (Spec. AN-NN-P5IIA), each 3/16-inch thick, 9 inches long,and having a width equal to the width of thelongeron at the affected area (see Figure66). Clean all wood chips from the surfacesto be glued. With a brush, apply one coat ofcasein glue (Spec. C-G-456 ) to all joint surfacesand to the skin and replacement longeroncontact surfaces. Assemble the scarf joint asshown (see Figure 66), and immediately applya series of clamps to the joint area to providea uniform pressure of approximately 150 Ibs./sq.in. /'see Figure ^8). To provide gluing pressurealong the skin for the longeron replacementsection, cut a strip of soft pine 1/4-inch thick and 1/2-inch wide and having alength equal to the length of the longeronreplacement. On the outside skin surface, nai IRESTRICTED[81]

icabi10I lerRESTRICTEDFUSELAGEPAR. 67 TO 68BEVEL 45° (6 PLACES)SPRUCE LONGERON REPLACEMENT5/8X3XLENGTH OF REPLACEMENT (I REQ.)TAPER CUT DAMAGED LONGERON AND TRICLEATSSECURE REPLACEMENTS WITH CASEIN GLUESPRUCE TRICLEATREPLACEMENT 1/2 XI5/I6"X LENGTH OF REPLACEMENT (2 REQ)-SPRUCE OR 3 PLY MAHOGANY PLYWOOD REINFORCEMENT, 3/16X3X7-1/2 (REQ.),3/16X2X7-1/2 (I REQ.), 1/8X1X6 (I REQ.), 1/16 X3XI-5/8 (IREQ.) SECURE TO LONGERONWITH CASEIN GLUE. CLAMP JOINTS WITH UNIFORM PRESSURE OF I50#/SQ. IN.FOR AT LEAST 4 HOURS.-DIRECTION OF SPRUCE GRAINOR OUTER FACE PLYWOOD GRAIN.ORIGINAL WEBTOP VIEWLONGERON WEB REPLACEMENT 3 PLY MAHOGANYPLYWOOD 1/16 X 3-1/4 X LENGTH OF REPLACEMENTEA/D VIEWFigure70 — \Vooden Fear Fuselage Lower Longeron Splicethe soft-pine strip to the longeron throughthe skin, using ^/8-inch long No. 20 gage nailsat one-inch spacing. Al low the glue to renainunder pressure for at least 4 hours; thenremove the clamps and the nailing strip. Carefullyscrape off all excess glue around theJoints. To the entire added wood material,apply one coat of sealer (AAF Spec. 141 13),and allow to dry for at least one hour. Applya second coat of sealer (AAF Spec. 141 15)to all added material except mahogany if used,and to the end-grain surfaces of the 3/16-inch reinforcements. To mahogany and the endgrainsurfaces of the 3/16-inch reinforcementsand to the nail holes in the outsideskin, apply one coat of unthinned filler ( NAAFormula No. 1398 or equivalent), allowing thefiller to remai n on the surface from 2 to 5ler withminutes; then remove the excess f i Iexcelsior. Al low the f i to dry for atleast 2 hours; then to all added material exceptthe end-grain surfaces, apply one finishcoat of either eel lu lose nitrate lacquer (Spec.AN-TT-L-'^M or aircraft enamel (Spec. 3-98)as required to match adjacent surfaces. Tapethe exposed end-grain surfaces of the reinforcementsas outlined in the paragraph ap-e to f i nishes.pI68. WOODEN REAR FUSELAGE LOWER LONGERONS.The wooden rear fuselage lower longerons ( NA88-31 121 ), which extend I inches aft of thefuselage joining surface on the left and theright side of the fuselage, are each formed ofa main member of rectangular spruce having athickness of 5/8-inch and a width of approximately3 inches. This main spruce member is[62]RESTRICTED

lowingFUSELAGEPARAGRAPH 68RESTRICTEDsupported along the entire length by two sprucetrie I eat sections and a plyvwDod web (see Figurejo). At both the front and rear ends of thelongeron, the main spruce member of the longeronis spliced to a short compregwood sectionfor greater strength. The complete replacementof a damaged longeron is not recommended, butsplicing and partial replacement may be accomplished.No repairs to the lower longeronare permissible at the coinpregwood ends ofthe longerons. If damage occurs in this area,replace the affected longeron section, andlocate the spl ice beyond the compregwood area.At the location where the splice is to belocated, vertically cut the longeron web and3 PLY MAHOGANY PLYWOOD PATCH OF SAMETHICKNESS AS SKIN. SECURE TO SKIN WITHCASEIN GLUE.SECTION A-AFigure 71— Flush Patch for Plywood Skin HolesLess Than 2 Inches Wideremove the affected web section as required(see Figure yo) . If the damage is extensive,cut out the damaged longeron material with akeyhole saw. On the main spruce member of thelongeron, make the cut at a ten-to-one taper,which amounts to a 7-1/2 inch taper on thisparticular section (see Figure jo). If the twospruce tricleats are damaged, taper cut outthe damaged material as noted (see Figure no).,Carefully plane off the length of the damagedlongeron and tricleats until the skin is bared.With a chisel, pare off any small remainingportions of the damaged material adjacent tothe scarf cuts which may be inaccessible forthe manipulation of the wood plane. Duringthese cutting operations, take particular careto prevent damage to the skin. Cut replacementtricleats and a new longeron section ofspruce (Spec. AN-S-6) to match the removed sections.Scarf-cut the ends of the replacementmembers to match the cut ends of the originallongeron sections. Carefully plane down theends of the scarf cuts (see Figure 6^) so thatthey are perfectly smooth when measured witha straightedge (see Figure 67), and so thatthey form a perfect fit when assembled. Toserve as reinforcements, cut the four piecesof spruce or four pieces of three-ply mahoganypoplar core plywDod to the sizes noted (seeFigure yo) . Do not use sandpaper on surfaces,which are to be glued, since the wood dust particlesprevent proper gluing. Clean all woodchips from the surfaces to be glued. With abrush, apply one coat of casein glue (Spec.C-G-456) to all longeron and tricleat jointsurfaces and to the skin and replacement longeroncontact surfaces. Assemble the scarf jointsas shown (see Figure no), and immediately applya series of clamps to the joint area to p rovidea uniform pressure of approximately 150Ibs./sq. in. To provide gluing pressure alongthe skin for the replacement sections, cutstrips of soft pine 1/4- inch thick and 1/2- inchwide and having a length equal to the lengthof the replacement. On the outside skin surface,nail the pine strips to the replacementsections through the skin, using 5/8-inch longNo. 20 gage nails at 1-1/2 inch spacing. Usea sufficient number of nailing strips to covercompletely the glued area. Allow the glueto remain under pressure for at least 4 hours;then remove the clamps and nailing strips.Carefully scrape off all excess glue aroundthe joints. To the entire added wood materialexcept the area to be glued to the longeron web,apply two coats of sealer (AAF Spec. 141 13);a I each coat to dry for at least one hour.To all end-grain surfaces, apply one coat ofunthinned filler (NAA Formula No. 1398 or equ i v-a lent J, allowing the filler to remain on thesurface from 2 to 5 minutes; then remove theexcess filler with excelsior. To serve asthe longeron web replacement, cut a sheet ofthree-ply mahogany plywood 1/16-inch thick,3-1/4 inches wide, and having a length equal tothe length of the replacement. With a spokeshave,scarf the ends of the longeron, plywoodweb (see Figure n^) , original web, and replacementweb at a 15/16-inch taper so that theyform a perfect fit when assembled. Apply twocoats of sealer (AAF Spec. 141 13) to the insideof the longeron web except the areas to beglued. Assemble the longeron web splice andreinforcement with casein glue (Spec. C-G-456),RESTRICTED[83]

holesRESTRICTEDFUSELAGEPAR. 68 TO 70and apply gluing pressure for at least 4 hours.To the web replacement, apply two coats ofsealer (AAF Spec. 141 15) allowing each coatto dry for at least one hour. To all outsidesurfaces except the end-grain surfaces of theweb reinforcement, apply one finish coat ofeither cellulose nitrate lacquer (Spec. AN-TT-L-51 ) or aircraft enamel (Spec. 3-98) as requiredto match adjacent surfaces. Tape the endgrainsurfaces of the web reinforcement as outlinedin the paragraphs applicable to finishes.69. FLUSH PATCH FOR PLYWOOD SKIN HOLES LESSTHAN 2 INCHES WIDE.Wherever possible, the flush patch should beused for smalI in the skin. However,where the application of this flush patch isnot feasible because of stringers irmiediatelyadjacent to the skin hole, the external patchor skin section replacement may be used. Theprocedure for the application of an inner patchto the fuselage skin is as fol lows: With a smallhole saw, cut out the damage to a circular shapeso that the removed skin includes all cracksand broken edges. The trimmed skin hole shouldnot exceed 2 inches. To serve as the patch,cut a circular piece of three-ply mahogany poplarcore plyv^ood of the same thickness as the skinand large enough to lap the skin hole by 3/4-inch all around (see Figure Ji)' To serve asthe filler, cut another circular piece of threeplymahogany poplar core plywood of the samethickness as the skin and of the same shape asthe removed material. Apply one coat of caseinglue to the contact surfaces of the patch andfiller and assemble as shown (see Figure yi)-:To provide gluing pressure, apply a short stripof soft pine over the f i I ler and temporari lynail the filler to the patch with No. 20 gagenails passed through the pine strip. With asmall wood plane, carefully plane off a ^^erythin layer of wood on the inside of the skinaround the hole where the patch is to be applied.To the end-grain surfaces of the skinhole, filler, and patch, apply one coat of sealer(AAF Spec. 141 13). Al low to dry for one hour;then apply one coat of unthinned filler ( NAAFormula No. 1398 or equivalent , allowing thefil ler to remain on the surface from 2 to 5minutes. Then remove the excess filter withexcelsior. Apply one coat of casein ^lue tothe contact surfaces of the patch and the skin,and assemble as shown (see Figure yi)- To providegluing pressure, apply a short strip ofsoft pine over the patch and temporarily nailthe patch to the skin in several places withNo. 20 gage nails passed through the pine strip.Allow the glue to dry for at least 4 hours;then pull out the nails and remove the nailingstrips. lo the accessible patch area onthe inside of the fuselage, apply one coat ofsealer (AAF Spec. 141 13), al low to dry for onehour, and then apply one coat of either cellulosenitrate lacquer (Spec. AN-TT-L-51 ) or aircraftenamel (Spec. 3-98) as required to matchadjacent surfaces. To the outside surface ofthe filler, apply two coats of sealer, allowinga drying time of one hour Detween coats. Trenapply one coat of surfacer (AAF Spec. 141 15),allow to dry, and sand off. Lastly, applyone spray coat of aircraft enamel (Spec. 5-98)pigmented with 16 ounces of aluminum pasteSECTION A-ATRIM DAMAGETO ROUND SHAPE'3/4" PATCHOVERLAP3 PLY MAHOGANY PLYWOODPATCH OF SAME THICKNESSAS SKIN. SECURE TO SKINWITH CASEIN GLUE.Figure 72— External Patch for Plywood Skin HolesLess Than 2 Inches Wide(Spec. AN-TT-A-461) per gallon of i>nthinnedenan)e I .70. EXTERNAL PATCH FOR PLYWOOD SKIN HOLES LESSTHAN 2 INCHES WIDE.^wr>^-The external patch should be used only in anemergency where the appi ication of a smal I innerflushpatch or a section replacement is impossible.With a small hole saw, cut out thedamage to a circular shape so that the removedskin includes all cracks and broken edges. Thetrimmed hole should not exceed 2 inches. Toserve as the patch, cut a circular piece ofthree-ply mahogany poplar core plywood (Spec.AN-NN-P-51 lA) of the same thickness as the skin[84]RESTRICTED

inglowonglowdFUSELAGEPAR. 70 TO 71RESmiCIW71. REPLACING DAMAGED PLYWOOD SKIN SECTION.LSCARF CUT MUST BESMOOTH a STRAIGHTFigure 73— Scarfing Plywood Skin With Spokeshaveand large enough to lap the skin hole by 3/4-inch all around (see Figure 72 J. Bevel the edgesof the patch. With small wood plane, carefullyplane off the finish on the outside of the skinaround the hole where the patch is to be applied.Apply one coat of casein glue (Spec.C-G-456) to the contact surfaces of the patchand the skin, and apply the patch. To providegluing pressure, apply a short strip of softpine over the patch and tempo rari ly na i I thepatch to the skin in several places with No.20 gage nails passed through the pine strip.Allow the glue to remain under pressure for atleast 4 hours; then remove the nai Is and nai I-strip. Carefully scrape off all excessglue around the joint. To the inside and theoutside of the patch, apply one coat of sealer(AAF Spec. 141 13 ) and a I to dry for atlease one hour. Apply a second coat of sealerto the inside and the outside of the patch exceptto the end-grain surfaces of the skin holeand patch. To the end-grain surfaces of theskin hole and patch,, apply onecoat of unthinnedfiller (NAA Formula No. 1398 or equivalent),allowing the filler to remain on the surfacefrom 2 to 5 minutes; then remove the excessf I I ler with excels ior. Allow the filler todry for at least 2 hours; then to the outsideof the patch, apply one coat of surface r (AAFSpec. 141 15) and allow to dry- To the insideof the patch, apply one coat of either cellulosenitrate lacquer (Spec. AN-TT-L-51 ) oraircraft enamel (Spec. 3-98) as required tomatch adjacent surfaces. To the outside of thepatch, apply one coat of clear aircraft enamel(Spec. 3-98) pigmented with 16 ounces of aluminumpaste (Spec. AN-TT-A-461 ) per gallonof unthinned enamel. This completes the repai r procedure.If a hole in plywood skin exceeds 2 inchesin width, cut out the entire affected plywooda rea i n a I i tud i na I i rect i on at the neareststringer on each side of the da.nage and as faras required in a transverse direction betweenframes (see Figure 74 j. Use a fine keyhole sawto remove the main portion of the skin, and achisel for the removal of small areas inaccessiblefor the manipulation of the saw. Usegreat care to cut the edges of the plywoodstraight and square. After the skin is thustrinnmed, scarf the skin edges at a 15-to-l tapervith a spokeshave or a drawknife (see Figure73 J. This scarf taper amounts to a 15/16-inchscarf on 1/16-inch skin, a 1-13/32 inch taperon 3/32-inch skin, and a 1-7/8 inch taper on1/8- inch skin. (See Figure J4. ) The scarfcuts must be smooth and perfectly straight whenmeasured with a straightedge. To serve as theskin section replacement, cut a sheet of 3- plyrrahogany poplar core plywood (Spec. AN-NN-P5IIA)of the same thickness, length, and width as theremoved skin (see Figure 6g) . The replacementskin section must be cut so that the wood graind i rect ion wi 1 I be laid 45 degrees to the verticalframes. If any appreciable bending in thereplacement skin section is necessary in orderto match the original skin contour, soak theplywood sheet in hot water, bend over a form,and clamp in place until dry. The form overwhich the plywood is to be bent should be madeso as to aI for some springback after theplywood has dried. Placing the plywood in anatmosphere of saturated steam is the most desirablerrethod for softening the plywood, butthe method is a more difficult process to manage.With a spokeshave, scarf the edges ofthe skin replacement section to match the preparededges of existing skin- This replacementskin section must fit the scarf edges ofthe original skin smoothly and snugly withoutwrinkles, bulges, or gaps. Apply one coatof casein glue (Spec. C-G-456) to both surfacesof the scarf-cut joint and to any interiorstructure contact surface of the replacementskin. Press the replacement skin into positionand make certain the placement is correct.After the replacement skin is pressedinto place, avoid any unnecessary adjusting andsliding of the skin section, as this actiontends to spread the glue, thus producing a poorjoint. To prevent the skin from slipping undersubsequent operations, drive three No. 20 gagenails through the center of the skin splice,leaving the nail heads protruding sufficientlyto allow subsequent removal of the nails. ToRESTRICTED[85]

.lewRESTRICTEDFUSELAGEPAR. 71 TO 73provide gluing pressure for the skin scarfjoint, secure the scarf joint with nailingstrips of I /4 X 1/2 X 12 soft pine. Na istripsI theover the scarf joint with 5/8-inch No.20 gage nails spaced at f-l/2 inches on centers(see Figure do). To prevent bulges, nailfirst at the center of the scarf joint length,and then progress outward from the center.Use sufficient nailing strips to cover entirelythe scarf joint area. For example, ifthe width of the scarf joint is 7/8-inch, usetwo sections of f/2- inch wide nailing strips.After several of the strips are secured, removethe three nails previously driven to securethe skin panel. Al low the scarf jointarea to remain under pressure for about 4 hours;then remove the nails and the nailing strips.The replacement skin section must be free ofwrinkles, bulges, and joint gaps. Scrape offthe excess glue from around the scarf Joints.To the inside and outside of the skin panelreplacement, apply two coats of sealer ( AAFSpec. 141 13), allowing a drying time of onehour for each coat. To the outside of the skinpanel only, apply one cOat of unthinned filler( NAA Formula No. 1398 or equivalent) by rubbingler on the wood with a c loth, the directionthe f i I of the strokes being across the grain.Allow the filler to remain on the outside surfacefrom 2 to 5 minutes, then remove the excessby wiping across the grain with excelsioror with a burlap rag. To the outside ofthe skin section, brush on one coat of surfacer(AAF Spec. 14115) and a I to dry for hours.*=-Lightly sand the surface with fine sandpaper.To the inside of the skin section replacement,match adjacent surfaces with either one coatof cellulose nitrate lacquer (Spec. AN-TT-L-51) or aircraft enamel (Spec. 3-98) as required.To the outside of the skin sectionreplacement, apply one coat of aircraft enamel(Spec. 3-98) pigmented with 16 ounces of aluninumpaste (Spec. AN-TT-A-46f) per gallon of unthi nned ename I72. BULKHEAD PLYWOOD WEBS.The bulkhead webs consist of thin 3-ply mahoganystiffened with spruce members (see Figure^o). The webs may be repaired as outlinedin the paragraphs pertaining to the repairof the plywood skin.73. WOODEN REAR FUSELAGE FORMERS.The wooden rear fuselage formers are formedFRAME(REF.)FRAME(REF.)STRINGERS(REF.)15/16" ON 1/16" SKINI- 13/32" ON 3/32" SKIN1-7/8" ON 1/8" SKIN15 TO ITAPERTRIM OUT DAMAGED SKIN BETWEEN NEARESTSTRINGERS AND SCARF SKIN EDGES.ORIGINAL SKINSTRINGER (REF.)-3 PLY MAHOGANY PLYWOOD REPLACEMENT SKIN OF SAME THICKNESS ASDAMAGED SKIN. SECURE WITH CASEIN GLUE AND APPLY PRESSURE WITHNAILING STRIPS FOR 4 HOURS. REMOVE NAILING STRIPS AFTER GLUE HARDENS.SECTION A-AFigure 74—Plywood Skin Section Replacement[86]RESTRICTED

eateateatseatsulerFUSELAGEPARAGRAPH 73RESTRICTEDof a narrow 5-p\y mahogany poplar core web withsmall spruce trie I glued along the edges.If the formers become damaged, they may be repairedby splicing in a new former section asshown (see Figure 75J . At each side of thedamage, cut the web section with a smal I finetoothedsaw and remove the damaged web. Removeany of the damaged spruce trie I adjacentto the skin. Scarf-cut the damagedspruce trie I section located along the innerside of the former, making the cut at a 3-inchtaper. Cut and form new spruce (Spec. AN-S-6)trie I sections to match the spruce trie I eatspreviously removed. Carefully plane down thescarf cuts of the inner tricleat joint so thatthe cuts are perfectly smooth when measuredwith a straightedge, and so that they form aperfect fit when assembled. With a spokeshave,scarf-cut the edges of the original web (seeFigure 73 J at a 15-to-l taper, which amountsto a 15/16-inch taper on this 1/16-inch thickplywood (see Figure 75J . Apply one coat ofcasein glue (Spec. C-G-456) to the skin andthe outer replacement trie! eats, and secure thetricleats to the skin. To provide gluing pressurefor the spruce tricleats, apply nai I ingstrips over the joint area on the outside skin,SKIN (REF.)REPLACE DAMAGED SPRUCE TRICLEATADJACENT TO SKINNEW SPRUCE TRICLEAT SECTIONl/2"X 1/2" X LENGTH OFREPLACEMENTREPLACE B-1349 PLYWOODGUSSETS OVER WEBREPLACEMENTWEB REPLACEMENT 3-PLYMAHOGANY POPLAR COREPLYWOOD I/I6"X 5"X LENGTHOF REPLACEMENT 3^niSCARF CUT DAMAGED SPRUCETRICLEAT AND SPLICE INNEW SECTION WITH CASEINGLUE \1EXISTINGSPRUCE TRICLEATEXISTING WEB >-SECTIONS/D£AA^1VIEWFigure TS-^Wooden Rear Euselage Former SpliceSPECIALKEY HOLENARROW DOVE TAILSPECIAL GROSS GUTFigure 76— Typesof Sawsback up the tricleats, and then nail the skinto the tricleats. Apply one coat of caseinglue (Spec. C-G-456) to both surfaces of theinner spruce tricleat scarf joint, assemblethe joint, and clamp in place. Allow the glueto dry for at least 4 hours, and then removethe clamps and the nailing strips. Apply onecoat of casein glue to the replacement web sectionand to scarf joint contact surfaces of thetricleats and replacement web. Press the replacementweb into the proper position on theformer, and clamp the glued surfaces. Allowthe glue to dry for at least 4 hours; then removethe clamps. To the entire added material,apply one coat of sealer (AAF Spec. 141 13) andallow to dry for one hour. To a I I the addedmaterial except the mahogany web and the endgrainsurfaces of the outer tricleats, applyanother coat of sealer (AAF Spec. 141 13). Tothe end-grain surfaces of the outer tricleatsand mahogany web, apply one coat of unthinnedf i I ler (NAA Formula No. 1398 or equivalent) byrubbing on with a cloth. Al low the fi I toremain on the surface from 2 to 5 minutes, andthen remove the excess by wiping with excelsioror with a burlap rag. Match adjacent finishwith either one coat of eel I lose nitrateRESTRICTED[87]

•RESTRICTED*'t1--'v4»'«4-rir..areasare as follows (see Figure 79 JfffO.'.-iA'.r' -.*-?.>,HALF ROUND RASPHALFCOARSE FLATHALF ROUND MILLFLAT MILLFigure 77^Types of FilesPrior to gluingApply doped sealing tape to a 1 1ROUND BASTARDfor Woodlacquer (Spec. AN-TT-L-51) or one coat of aircraftenamel (Spec. 5-98) as required.74. GENERAL FINISH REQUIREMENTS FOR WOOD.or finishing any wood surface,thoroughly clean the surface to removesawdust, grease, and any other foreign matter,from the wood. Wherever possible, accomplishall gluing prior to finishing, and remove a I Iexcess glue from visible interior or exteriorparts before applying any finish. Whenever thepeculiarity of the part is such that it is moreconveniently glued subsequent to finishing,mask all areas to be glued with Scotch Tape(or suitable equivalent). No adhesive must beleft on the wood when the tape is removed .previouslytaped. Before attaching metal fittings,supports, etc., protect them from direct contactwith the wood by applying two coats ofsealer, or one coat of sealer and one coat off I I ler, to the wood under the fitting. For convenience,the wood surfaces may be divided intothree groups, in order to faci litate thespecification of finishes. The wood surfacegroups and the corresponding finish for each75. WOOD FINISH MATERIALS.:FUSELAGEPAR. 73 TO 76The following materials are required in theappi ication of a protective finish to the woodrepai r materialMATERIALENAMEL. CAMOUFLAGESEALER, WOOD, LIQUIDSURFACER, WOOD,L l(JU I DENAMEL, AIRCRAFTPASTE, ALUMINUMDOPE,CELLULOSENITRATE, CLEARLACQUER, CELLULOSENITRATEFILLER,WOOD76. WOOD REPAIR MATERIALS.SPeClFICATlOH1U109, AAF1U113, AAF1U115, AAF3-98, U.S. ARMYAN-TT-A-461, ARMY-NAVYAN-TT-D-514, ARMY-NAVYAN-TT-L-51 ,ARMY-NAVYNAA FORMULA NO. 1398OREQUIVALENTThe following is a summation of the typesof wood and glue that may be required in therepair of the wooden rear fuselage section.The quantities of wood required for a particularFigure 78— Types of Planes for Wood[68]RESTRICTED

II Inclsurfaces,FUSELAGEFIGURE 79RESTRICTEDWOOD SURFACEDESIGNATIONWHERELOCATEDREQUIREDFINISHTypeThis group iudesthe interior ofclosed members orsurfaces normal 1y remainingunseen, suchas stab i 1 i zer i nteriors,hollow longerons,the undersideof floorboards, etc.Apply two brush coats of sealer (AAF Spec. 14113), allowing eachcoat to dry one hour. The second coat of sealer may be aluminizedwith 8 ounces of aluminum paste per gallon of unthinnedsealer. Do not thin the sealer to less than 25? nonvolatilecontent. To assure complete coverage, and as a guide in subsequentsanding, it is permissible to add a suitable dye to theseal e r.TypeThis group includesthe interior opensurfaces, parts, andmembers vi si bl e tooccupants of theai rcraft, such as thef usel age compartmentsand attachingparts.As outlined for TypeIapply one coat of sealer (AAFSpec. 14113) arid allow to dry for one hour. Apply a second coatof unaluminized sealer to all surfaces except the exposed openend grains and mahogany.To mahogany and to exposed open end grains, apply one coat offiller (nAA Formula no. 1398 or equivalent) by rubbing with acloth. The direction of the strokes shall be across the grain.Allow the filler to remain on the surface from 2 to 5 minutes,and then remove the excess with excelsior or with a burlap rag.Tape all end-grain surfaces with regtilation glider fabric orequivalent, predoping the tape with one heavy coat of clear dope(Spec. AN-TT-D-514) and allowing to dry. Tape shall be pinked onall edges and shall g'rve a minimum lap of 1/2-inch onto thestraight-grain surface. Apply a second coat of dope to the tapeand immediately press the tape down firmly on exposed end grain,eliminating all air pockets. Apply an additional coat of cleardope if the nap of the tape is not satisfactorily filled andsealed down. Doped tapes will not adhere satisfactorily to finishedareas. End-grain surfaces not suitable for taping shouldhave an extra coat of sealer in addition to the regular finish.When finishing at drain holes and other inaccessible places,use a stiff wire covered with cloth or use an ordinary pipecleaner. Allow tape to dry before applying any finish.Match adjacent surfaces with either one coat of cellulosenitrate lacquer (Spec. An-TT-l-51) or aircraft enamel (Spec.3-98). The finish coat shall be of sufficient thickness togive complete hiding of the finishing materials.TypeIi ncl udesThis groupthe exterior surfaces,parts, andm embe rs, such assurfaces exposed tothe weather and toview from the outside.Apply two coats of unaluminized sealer (AAF Spec. IUII3) asoutlined for Type I surfaces.Apply one coat of filler (nAA Formula no. 1398 or equivalent)as outlined for Type II surfaces.Apply one coat of surfacer (AAF Spec. 14115) and allow to air-dryfor 6 hours. With dry sandpaper, sand the surfacer after it hasbeen thoroughly dried. Sand so as to produce the minimum possiblethickness of the surfacer coat but do not sand to thefiller coat beneath.Match adjacent finish with one coat of clear enamel (Spec. 3-98)pigmented with 16 ounces of aluminum paste (Spec. AN-TT-A-461)per gallon of unthinned enamel.Figure 79'^Finish Requirements for WoodRESTRICTED[89]

RESTRICTEDFUSELAGEPAR. 76 TO 77repair must be determined local \y with referenceto the applicable repair procedure outlinedin this Section. ^MATERIALPLYWOOD, 3-PLY MAHOG-ANY, POPLAR COREPLYWOOD, 3-PLY DOUG-LAS FIR, POPLAR COREPLYWOOD, 3-PLY DOUG-LAS FIR, DOUGLASFIR COREPLYWOOD, 3-PLY REDGUM, RED GUM ORPOPLAR CORESPRUCEPOPLAR, YELLOWHEMLOC K , WESTER NFIR, NOBLECOMPREGWOODSPECIFICATIONAN-NN-P-511A

izerizer.RESTJilCITJ)FIXED SURFACESPAR, I TO 3SECTION 3FJXED SURFACES1. WING CONSTRUCTION.area of 28.34 square feet, including 5.53square feet of fuselage area. The stabilizerconsists of two interchangeable sectionsThe low wing is of full-cantilever, stressedskin design and consists of a detachable centersectionand detachable outer wing sections. of the stabilizer consists essentially of twobolted together at the center. Each sectionThe entire wing has a span of 42 feet 1/4-inch spars, regularly spaced ribs, and externaland an area of 253.72 square feet. The structureis formed from 24ST ale lad sheet and 24ST ure ^). The rear spar of each of the stabi-skin stiffened by spanwise stringers (see Figaluminumalloy extrusions and sheet, consistingI izer sections carries all shear loads as aof spars, power pressed ribs, and skin stiffenedin a spanwise direction by stringers. The spar Is carried as shear in the skin to thecantilever beam. The torque about the rearouter wing sections and the centersect ion are inboard rib where the shear is reacted by ajoined by external bolting angles along the vertical couple at the spar attaching points.upper and lower skin surfaces. A plate rib On the later AT-6C Series Airplanes, the aluminumhorizontal stabilizer is substituted byat each joint distributes shear from the singleouter panel main spar to the two centersectionspars. The centersect ion is of con-Figure 5J. The basic construction of thea stabi I constructed entirely of wood (seestant chord design and has an incidence of wooden horizontal stabi I d i ffers littleplus 2 degrees, and no dihedral angle. Each from that of the aluminum structure. The onlyouter wing incorporates an angle of incidence noticeable difference is the four additionalof plus 2 degrees at the root which gradually nose ribs which are attached to the front sparchanges to an angle of incidence of degrees to give additional support to the plywood leadingedge. The spar caps are made fromcom-at the tip. Each outer wing is bolted to thecentersect ion at a dihedral angle of 5 degrees pregwood with the webs of mahogany-poplar plywood(see Figure 6). The rib webs are also41 minutes. The skin panels vary from .020 to.064 inch in thickness and are lap jointed. Type of mahogany-poplar plywood with spruce tricleatcaps. The stringers are of spruce andAD rivets (AITS aluminum) are used for assembly,all skin riveting being flush to approximately are held securely in position in the rib cutoutsby means of small wedges and glue. Ven-33% chord on the top skin and to 10% chord onthe bottom skin. The notation "wing station" tilating holes are provided in the web of allindicates the distance in inches measured from ribs and these holes are reinforced by a plywoodwasher which is glued around the hole (seethe center line of the airplane to any positionoutboard of the centerline. The term "station" Figure 5J . The 3-ply, mahogany-poplar plywoodskin is glued to the stringers, ribs, andas used in reference to each outer wing maysignify the distance in inches outboard of the spar caps by means of casein glue. The twowing joint, or outboard of the centerline of sections of the wood stabilizer are bolted atthe airplane as noted (see Figure g) . The the rear spar through aluminum alloy fittingscentersect ion is attached to the underside of (see Figure 6). The attaching fitting on thethe welded steel fuselage tubing truss by bolt front spar consists of a light aluminum sheetfittings at the two spars of the centersection.For quick reference, the part numbersbolted to the spar (see Figure j)of the rib and spar structure of the centersectionare illustrated (see Figure i). The3. VERTICAL STABILIZER CONSTRUCTION.part numbers of the rib and spar structure of The vertical stabilizer is a full-cantileverthe outer wing are also shown (see Figure 2). aluminum alloy structure having an area of 5.33The part numbers of the centersect i on fuel square feet and a normal setting of 1-3/4 degreesto the left. The stabilizer is construct-tank compartment doors are shown (see. Figure3J.ed of 24ST ale lad sheet consisting of two spars,several ribs, .and skin stiffened in a spanwisedirection with stringers (see Figure 8).2. HORIZONTAL STABILIZER CONSTRUCTION.The rear spar is the main load-bearing memberof the structure and is formed of a flatThe horizontal stabilizer is a full-cantileveraluminum alloy structure having a symmetricalairfoil. The stabilizer has a total length. The leading edge portion of24ST sheet, with U flanges formed along thethe[91]RESTRICTED

FIXED SURFACESFIGURE IRESTRICTEDHORTH AHERICAIt AVIATIOH PART NUMBERSCONTRACT NOS.AT-6A AT-68 AT-6C77 8"* 88DRAWINGNOS.CONTRACT NOS.

RESTRICTED FIXED SURFACESFIGURE 2HORTH SHERICiH tVltJIOH PiRJ HUHBIRS

FIXED SURFACESFIGURE 3RESTRICTEDEXTERIORVIEW1.

RESTRICTED FIXED SURFACESFIGURE ^55-21010(84-2l010(AT-6B)55-21024( 84-21026 (AT-6B)55-2102655-2100955-2100455-21013(84- 21007 (AT55-2100755-2101777-21003REF. DWG.AT6A-77-2I00IAT6B-84-2I00IFigure 4-^Alumintim Horizontal Stabilizer Structure[95]RESTRICTED

FIXED SURFACESFIGURE 5RESTRICTED77-21010(2 REQ)52-2102688-2103288-2103352-21018A/OTES:1. ALL WEB MEMBERS AND SKIN PANELSARE MADE FROM 3 PLY MAHOGANYPLYWOOD OR EQUIVALENT CONFORMINGTO AAF SPEC. AN-NN-P5IIA.2. ALL STRINGERS, RIB CAPS ANDSPAR CAPS ARE MADE FROMSPRUCE, CONFORMING TO AAF SPEC. AN-S-63. FOR ASSEMBLY, CASEIN GLUE IS USEDCONFORMING TO AAF SPEC. C-G-4564. PART NUMBERS NOTED.REFER TO FITTINGS(THE ONLY REPLACEABLE ITEMS)f?EF. DWG. 88-21001 (AT- 60Figure 5— Wooden Horizontal Stabilizer StructureRESTRICTED[96 1

.loyRESTRICTEDFIXED SURFACESPAR. 3 TO 5stabilizer is flush riveted, the remaining portionbeing riveted with brazier head rivets.^. STRINGERS - GENERAL.Most of the stringers used in the constructionof the fixed surfaces are extruded shapesof 24ST a luminum a Imaterial. However, afew rolled shapes of 24ST ale lad sheet materialare also employed. In the AT-6A wing, thestringer types and locations in the right wingare symmetrical with those in the left wing(see Figure g) . In the AT-6B and AT-6C wing.Figure 6 — Rear Spar Joint ofWooden Hor izonta 1 Stabi lizerthe stringer types and locations in the leftwing (see Figure lo) are symmetrical with thosein the right wing except for a small portionof the upper and lower surfaces of the outerwing leading edge at the wing joint (see Figures11 and 12). On the AT-6A stabilizers,C I07LT rolled ale lad stringers are employed(see Figure 23 j . On the AT-6B and AT-6C stabilizers.Type 0373 LT rolled ale lad stringersare employed (see Figure 1^) . From these referencedrawings, determine the type number ofthe damaged stringer and then refer to the followingapplicable paragraph for the splice procedure.Those stringer types having a C prefixare North American Aviation, Inc., standardparts. Those stringer types having a Kprefix are Aluminum Oompany of America (Alcoa)die number stringers. For quick reference,the fol lowing chart sums up the types of stringersand the approximate quantities involvedin the original construction of the fixed surfaces(see Figure 1^) . For the Alcoa equivalents ofNorth American Aviation, Inc., standard extrusions,see the complete summation of extrusions inSect i on I5. STRINGER REPAIRS - GENERAL.In repairing damage to stringers, do not makeany splices in the bay at the end of the stringer;replace a portion of the damaged stringer andlocate the splice outboard of the first bay.If skin rivets through the stringer are spacedat 1-1/4 inches on centers or greater, doublethe number of existing skin rivets through thestringer at the splice. Skin rivets and splicerivets at the end of the splice must have atleast a 3/8-inch edge distance, and all splicemembers must be in the fully heat-treated condition.Where ribs interfere with the installationof a stringer spl ice, mod ify the rib cutoutas required. It is to be noted that wheredamaged stringers are short enough to permitentire removal, the stringers should be replaced.At the root of the outer wing, a CI23LT stringeris reinforced with a short length of Type C203Tstringer in one location and a CI40T in anotherposition. Repairs to these stringer combinationsshould be made by replacing the shortreinforcing stringer and by splicing the CI25LTstringer outboard of the doubled location. Ifany stringers are replaced near the wing joint,file the stringers so that they bear directlyupon a straight edge supported by the upperand lower attaching angles. During the reassemblyof the wing, where some of the stringersnear the wing joint were replaced, it maybe necessary to taper the end of the new stringersa \/ery small amount if interference occursin the wing instal lation.FRONTSPARFigure 7 — Wooden Horizontal Stabi I izerFront Spar Fitting[97]RESTRICTED

1FIXED SURFACESFIGURE 8RESmiCTED19-2301052-2300919-2300756-2300355-2300544-2302084-23020(AT-6BaG)52-23004REFDWG.AT6A-55-2300I-I6AT6B-84-2300IAT6C- 84-2300Figure 8— Vertical Stabilizer Rib and Spar Part NunhersRESTRICTED[98]

RESTRICTED FIXED SURFACESPARAGRAPH 6notes:spanwise linesindicate stringers;chordwise linesindicate stationlocations.* 1/8-inch trimmedfrom bulbITEM NO

.FIXED SURFACESPAR. 6 TO 7RESTRICTEDthrough the splice member, using the existingrivet holes as a guide. If the skin rivetsare spaced at 1-1/4 inches on centers or greater,double the number of existing skin rivets atthe spl ice. Remove the spl ice member and burrthe rivet holes. Apply one coat of zinc chromateprimer to all overlapping surfaces of thespl ice member. Clamp the spl ice member to thedamaged stringer. At each side of the damage,drive the five AN442-AD3 rivets in the upstandingflange and rivet the splice member throughthe sKin with the same size and type rivetspreviously removed from the skin (see Figure i6)7. STRINGER TYPE CI23LT.The Type CI23LT stringer is used extensivelyNOTES:ITEM NO.SPANWISE LINES 2INDICATE STRINGERS; 3CHOROWISE LINES 4INDICATE STATIONS 5* 1/8-INCH MILLED ^FROM STRINGER BULB 78STRINGERin the construction of the wing, both in theupper and lower surface. The stringer is formedof rolled 24ST alclad sheet, .040 inch thick.The stringer has no equivalent Alcoa number.For damaged stringers in the lower surface ofthe wing, if the damage is slight in nature, aType C203T extruded 24ST aluminum al loy spl icemember 7-1/2 inches in length will suffice(see Figure ij) . It Is to be noted that, ifthe stringer Is in the upper surface of thewing, the splice member may be shortened to6 inches with four splice rivets and two skinrivets on each side of the damage. If thedamage is extensive, cut out the damaged materialwith a hack saw, locating the cut midwaybetween two skin rivets at each side ofdamage. Skin rivets at the end of the splice

I equal.48TI lerthe.RESTRICTED FIXED SURFACESPAR. 7 TO 8must have at least 3/8- inch edge distance andthis may necessitate making the splice memberslightly longer than that noted. If the extrudedC203T splice member is not available,an .081 inch thick 24ST ale lad sheet bent upto a similar shape will sat isfactor ily serveas the spl ice member. Clamp the stringer spl icemember to the damaged stringer. With a No.21 ( . I 59 ) drill, at each s ide of the damagedrill the required holes at 3/4-inch on centersthrough the upstanding flange of the stringerand spl ice member. With a dri I to thesize of the skin rivet holes in the location,drill holes through the skin and splice memberat the same spacing as existing skin rivets.Remove the splice member and burr the rivetholes. Apply one coat of zinc chromate primerto all surfaces of the splice member and reclampthe member to the proper position on thedamaged stringer. At each side of the damage,drive the AN442-AD5 rivets in the upstandingflange and rivet the splice member skin flangewith the same size and type rivets employed tosecure the stringer to the skin (see Figure ly)slight, the length of the CI 48T splice memberneed not exceed 10-1/4 inches (see Figure i8)If the damage is extensive, cut out the damagedmaterial, locating the cut midway between twoskin rivets on each side of the damage. Cuta CI48T splice member equal to 10-1/4 inchesplus the length of the damaged area. Cut alength of CI48T as a f i I ler to match the removedmaterial. Clamp the CI48T splice memberand the C I f i to the proper positionson the damaged stringer; and with a No. II (.191)drill, drill eight rivet holes in the upstandingflanges at each side of the damage at 5/8-inch on centers. Using a drill ofthe samesize as the existing skin rivet holes in thestringer, drill rivet locations through theNOTE SPANWISE LINESINDICATE STRINGERS.CHORDWISE LINESINDICATE STATIONLOCATIONS. DOTTEDLINES INDICATESKIN PANELS.ITEM NOS. STRINGER-TWO C250TC250T-C366T- C265T- C366T-TWO C366T-MAIN SPARWINGJOINTSP4NWISE LINES INDIC»TESTRINGERS. CHORDWISELINES INDICATE STATIONLOCATIONS. DOTTED LINESINDICATE SKIN PANELS.ITEM NOS. STRINGERSCUJLT2/.'..'.C266T^'_|'K77A5". I' MAIN SPARFigure 11-^AT-6B and AT-6C Right Outer WingUpper Surface Stringer Details8. STRINGER TYPE C|il8T.The Type CI48T stringer is employed only inthe upper surface of the wing centersect ion.The stringer is formed of extruded 24ST aluminumalloy material, and it has the equivalentAlcoa die number of K734JJ. If the damage isFigure 12—AT-6B and AT-6C Right Outer ViingLower Surface Stringer Detailss-kin and splice member at the same spacing asexisting skin rivets. Remove the splice memberand the f i I ler and burr al rivet holes.IApply one coat of zinc chromate primer to allsurfaces of the splice member and the filler,and again clamp the members to the proper positionson the damaged stringer. At each sideL loi] RESTRICTED

FIXED SURFACESPAR. 8 TO 9RESTRICTEDVERTICALSTABILIZER2- -- CI07LT3- -- FRONT SPAR4- - REAR SPARnote:NOTE:SPANWISE LINESINDICATE STRINGERS.CHORDWISE LINESINDICATE RIB LOCATIONSITEM NOS. STRINGERSPANWISE LINES INDICATESTRINGERS. CHORDWISE LINESINDICATE RIB LOCATIONS.ITEM NOS.

RESTRICTED FIXED SURFACESFIGURE 15STRINGERTYPEUSED ONAIRCRAFTLOCATIONAPPROXIMATE TOTALLENGTH USED PER AIRPUNEC107LTAT-6ASTABILIZERS35 FEETCI23LTALLWING CENTERSECTIONAND OUTER PANEL350 FEETC148TallWING CENTERSECTIONUPPER SURFACE55 FEETC148T(doubled)ALLWING CENTERSECTIONUPPER SURFACE20 FEETC180TAT-6BAT-6CWING OUTER PANELLOWER SURFACE20 FEETC204TALLWING CENTERSECTIONLOWER SURFACE7 FEETC204T(doubled)ALLWING CENTERSECTIONLOWER SURFACE40 FEETC250TAT-6AOUTER WINGSURFACELOWER10 FEETC250T(doubled)ALLWING CENTERSECTIONLOWER SURFACE (WiNGOUTER PANEL LOWERSURFACE AT-6C ONLY)25 FEETC250T - C366T(combination)AT-6BAT-6CRIGHT OUTER WINGUPPER SURFACE10 FEETC265TALLOUTER WINGSURFACEUPPER35 FEETC266TALLOUTER WINGSURFACELOWER35 FEETC274TALLWING CENTERSECTIONUPPER SURFACE10 FEETC366TALLWING UPPER SURFACE275 FEETC366T(doubled)AT-6BAT-6CRIGHT OUTER WINGUPPER SURFACE (ONLYREPAIRABLE -LOCATION)10 FEETC373I-TAT-6BAT-6CSTABILI ZERS50 FEETK77AALLOUTER WINGSURFACELOWER50 FEETK77A(doubled)ALLWING CENTERSECTIONUPPER SURFACE10 FEETFigure 15— Summation of Stringer Types Used in Fixed Surfaces[ 103]RESTRICTED

4280.4654.1 equalice1 theFIXED SURFACESPAR. 10 TO IIRESTRICTED10. STRINGER TYPE CI80T.The Type CI80T stringer is used in one locationin lower surface of the AT-6B and AT-6Couter wing. The stringer is a bulb angle ofextruded 24ST aluminum al loy material, havingthe Alcoa type designation of K I If thestringer is only slightly damaged, a splicemember 9 inches long will suffice (see Figure2o). If the damage is extensive, cut out thedamaged material, locating the cut betweentwo skin rivets at each side of the damage.Prepare a filler of CI8UT extrusion to matchthe removed material. Cut a section of CI80Tequal to the length of the damage plus 9 inches.Clamp the splice member and the filler to theproper position on the damaged stringer. Ateach side of the damage, center punch six rivetlocations in the stringer upstanding flange at3/4-inch intervals. With a No. 21 (.159) drill,drill through the center-punched holes. Centerpunch new rivet locations on the skin, using thesame spacing as the existing rivets. With adrill equal to the size of the existing rivetholes in the skin, drill the center-punchedholes in the skin. Remove the splice memberand burr the rivet holes. Apply one coat ofzinc chromate primer to all surfaces of thesplice member and again clamp the member to thedamaged stringer. At each side of the damage,insert and drive six AN442-AU?i rivets in thestringer upstanding flange. Rivet the splicemember flange and filler to the skin with thesame type and size rivets employed in the- damagedstringer (see Fii/ure 20).I I. STRINGER TYPE C20i^T.The Type C204T stringer is employed only inthe lower surface of the centersect ion. Thestringer is formed of extruded 24ST aluminumalloy and has the equivalent Alcoa die numberK 1If the damage is of a minor nature,a 10-1/4 inch length of Type C204T extrusionwill be sufficient for the splice member (seeFigure 21) • If the damage is extensive innature, cut out the damaged material, locatingthe cut between two skin rivets at each sideof the damage. Prepare a f i I ler section ofC204T to match the removed material. For thesplice member, cut a C204T extruded section equalto the length of the damage plus 10-1/4 inches.Clamp the sp 1 member to the damaged stringerand center punch eight rivet locations at eachside of the damage at 5/8-inch spacing. Witha No. 21 ( . 159) drill, dri 1 eight rivetlocations at each side of the damage. Witha dri 1 to the size of the existing rivet7/2 PLUS LENGTHOF DAMAGE6>2" PLUS LENGTHOF DAMAGELOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGEAN442-AD3 RIVETS(5 REG. ON EACH SIDE)EXISTING SKIN RIVETSLOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGEAN442-AD5RIVETS-(5 REQ.ON EACH SIDE)C203TSPLICEMEMBERCI07LT-20SPLICEMEMBEREXISTINGSKIN RIVETS-^ (PT.h%HFigure 16— Splice for Stringer Type C107LT-20 Figure IT— Splice for Stringer Type C123LTRESTRICTED[ 104]

iceI TORESTRICTED FIXED SURFACESPAR. I13holes in the stringer, drill new rivet locationsin the sp I member and filler through theskin. Remove the splice member, burr the rivetholes and apply one coat of zinc chromate primerto all surfaces of the splice member. Againclamp the spl ice member to the damaged stringerand insert and drive eight AN442-AD5 rivetson each side of the damage. Rivet the fillerto the splice member through the existing skinholes. Using the same size and type rivetsas used to secure the damaged stringer, rivetthe splice [Tiember to the skin (see Figure 21).10 74" PLUS LENGTH OF DAMAGE5 '^"Q O O O O O O O~i^

FIXED SURFACESPAR. 13 TO mRESTRICTEDQ9" PLUS LENGTH OF DAMAGE^>^^V«^Q© OLOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGE.BUTT FITCI80T FILLERAN442-AD5 RIVETS(6 REQ. ON EACH SIDE)EXISTING SKIN RIVETSCI80TSPUCE MEMBEREXISTINGSKIN RIVETSSAME SIZE ANDSPACING AS EXISTINGSKIN RIVETS

1 surfaces.I equalI outRESTRICTED FIXED SURFACESPAR. m TO 15— -^\^/^ 101/4" PLUS LENGTH OF DAMAGEL LOCATE CUTBETWEENSKIN RIVETSON EACH SIDEOF DAMAGE.BUTT FITC204T FILLERS~TAAN442-AD5 RIVETS(8 REQ. ON EACH SIDE)EXISTING SKIN RIVETSC204T SPLICEFigure 22—Splice for Doubled C204TType Str ingersMEMBERSThe skin rivets at the end of the splice musthave at least 3/8-inch edge distance and thismay necessitate making the splice slightly longerthan that noted. Drill out the affectedskin rivets at each side of the damage. Chamferthe outside cornerB of the spl ice members to fitthe stringer radius. Clamp the splice membersand the fil lens to the damaged stringers. Centerpunch eight rivet locations in the upstandingflanges at each side of the darmge at an averagespacing of 5/8- inch on centers. With a No. fl(.191) drill, drill the center-punched rivetlocations. With a dri I to the size ofthe existing skin rivet holes, drill through thesplice members and fillers, using the existingrivet holes in the skin as a guide. Removethe spl ice members and f i I lers, and burr andclean the rivet holes. Apply one coat of zincchromate primer to a 1 Again clamp thesplice members and fillers to the damagedstringers. At each side of the damage in theupstanding flanges, insert and drive eight AN442-AD6 rivets. Using the same size and type rivetpreviously removed, rivet the splice memberflanges and fillers to the skin (see Figure 24).15. COMBINATION C250T - C366T STRINGERS.On the AT-6B and AT-6C Series Airplanes, thecombination of stringer Types C250T and 566Tis used in one location on the upper surfaceof the right outer wing just aft of the winggun compartment (see Figure 2^). If the damageis extensive, cut out the damaged material,locating the cut between two skin rivets ateach side of the damage. Cut filler extrusionsto match the removed material.. For thesplice members, cut a section of Type C204Textrusion equal to the length of the damageplus 11-1/2 inches; and cut a length of C250Textrusion equal to the length of the damageplus 14 inches (see Figure 25 J. It is to benoted that the skin rivets at the ends of thesplice must have at least 3/8-inch edge distanceand this may necessitate making the splicemembers slightly longer than that noted. Fora length of 6 inches on each side of the damage,trim the bulb of the damaged C366T stringerto permit fitting the C204T splice member.Dri I the affected skin rivets at each sideof the damage. Chamfer the corners of thesplice members to fit the inside radius of eachstringer. Clamp the spl ice members to the properpositions on the darraged stringers. At eachside of the damage in the upstanding flanges,center punch eleven rivet locations at an averagespacing of 5/8-inch on centers. With a No. 21(.159) drill, drill the center-punched rivetlocations in the stringers. With a drill equalC250T SPLICEMEMBER (ONELEG REMOVED)-LOCATE CUTBETWEEN 2SKIN RIVETSON ^ACH SIDEOF DAMAGEBUTT FIT C250FILLEREXISTING SKIN RIVETSVe±10)4' PLUS LENGTH OF DAMAGE5'^AN442- AD6RIVETS (8REQUIRED'ON EACH SIDE)EXISTING SKIN RIVETS^VrC250T SPLICE^ ^6 /MEMBERJ x; ^Tk^TiSAME SIZE AND SPACINGAS EXISTING SKIN RIVETSFigure 23—Splice for Stringer Type C250TLi07]RESTRICTED

I theAFIXED SURFACESPAR. 15 TO 16RESTRICTEDf-^VW-h--lOjj^" PLUS LENGTH OF DAMAGE•5'/8I 1 IJ >\i —i I\CZ3 CTD i >—4\1—"^3/^kLOCATE CAN442-AD6 RIVETSBETWEEN(8 REG. ON EACH SIDE)SKIN RIVETSON EACH SIDE AEXISTING SKIN RIVETSOF DAMAGE.BUTT FIT CI48T C250T SPLICE MEMBERSFILLER16. STRINGER TYPE C265T.The Type C265T stringer is employed as theleading edge stringer on the upper surface ofthe outer wing. The stringer is formed to a24ST aluminum alloy shape having the equivalentk^ ID1/4." PLUS LENGTH OF DAMAGE^8Figure 24-'Splice for DoubledC250T Type StringersLOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGEAN442-AD5 RIVETS(8 REQ. ON EACH SIDE)EXISTINGSKIN RIVETSBEND UP STRINGERFLANGE TO FITSPLICE MEMBERC265TSPLICEMEMBERto the size of the existing skin rivet holes, drillthe spl ice members and f i I lers through the skin.Remove the spl ice members and burr a I rivetholes. Apply one coat of zinc chromate primerto all surfaces of the splice members. A-gain clamp the members to the damaged stringh*^14" PLUS LENGTH OF DAMAGE7-ers. At each side of the damage in the upstandingflanges, insert and drive eleven AN442-AD5 rivets. Replace the skin rivets through thespl ice members and f i I lers with the same size andtype rivets previously removed (see Figure 25J.-\i\-K'XbhT;064-^ LOCATE CUTBETWEEN 2SKIN RIVETSAT EACH SIDEOF DAMAGEC250T SPLICEMEMBERAN442-AD5 fJIVETS(II REQ. ON EACH SIDE)EXISTING SKIN RIVETSTRIM BULBTO FIT C204TSPLICE MEMBERC204TSPLICE MEMBERFigure IS—Splice for CombinationC250T - C366T StringersFigure 26'-Spiice for Stringer Type C265TAlcoa die number L2355P. If the damage isslight, a C265T splice member, 10-1/4 inches long,will suffice (see Figure 26). If the damageis extensive, cut out the damaged material,locating the cut between two skin rivets at eachside of the damage. For the splice member, cuta section of C265T equal to 10-1/4 inches plusthe length of the damage. It is to be notedthat the skin rivets at the ends of the splicemember must have an edge distance of at least3/8- inch, and this may necessitate making thesplice slightly longer than that noted. AtRESTRICTEDL 108]

ampI asIII theI outIRESTRICTED FIXED SURFACESPAR. 16 TO 18each side of the damage, drill out the affectedskin rivets. In order to fit the spl ice member,bend up the flange of the damaged stringer (seeFigure 2^). Clamp the splice member to theproper position on the damaged stringer. Ateach side of the damage in the upstanding flangeof the splice member, center punch eight rivetlocations at an average spacing of 5/8-inchon centers. With a No. 21 ( . 159) dri II, dri Ithe center-punched rivet locations. Redri Ithe existing skin rivet locations through thespl ice member with the same size dri I originallyused to drill the holes. Remove thesplice member and burr all the rivet holes.Apply one coat of zinc chromate to all surfacesof the splice member and rec I the memberto the damaged stringer. At each side ofthe damage in the upstanding flange, insertand drive eight AN442-AD5 rivets. Rivet thesplice member to the skin with the same typeand size rivets originally used (see Figure 26).—^^^ 10'// PLUS LENGTH OF DAMAGEL LOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGE5'/e"K'lfH .064AN442-AD5 RIVETS(8 REQ. ON EACH SIDE)EXISTING SKIN RIVETSBEND UP STRINGERFLANGE TO FITSPLICE MEMBERC266TSPLICEMEMBERFigure 27— Splice for Stringer Type C266T17. STRINGER TYPE C266T.The Type C266T stringer is employed as theI asleading edge stringer on the lower surface ofthe outer wing. The stringer is formed of a24ST aluminum alloy shape having the equivalentAlcoa die number L23558. If the damage is extensive,cut out the damaged material with ahack saw, locating the cut between two skinrivets at each side of the damage. For thesplice member, cut a section of C265T equalto 10-1/4 inches plus the length of the damage(see Figure 27J . The skin rivets at the ends ofthe spl ice must have an edge distance of at least3/8- inch, and this may necessitate making thesplice member slightly longer than noted. Ateach side of the damage, dri I the affectedskin rivets. In order to fit the splice member,bend up the flange of the damaged stringer(see Figure 27 J . Clamp the spl ice member to thedamaged stringer. At each side of the damage inthe upstanding flange of the splice member, centerpunch eight rivet locations at an average spacingof 5/8-inch on centers. With a No. 21 (. 159) dri I I,dri I center-punched rivet locations. Redri Ithe existing skin rivet locations through the splicemember with the same skin rivet hole dri Ioriginal ly used. Remove the spl ice member and burrall the rivet holes. Apply one coat of zincchromate to all surfaces of the splice member.Again clamp the member to the proper positionon the damaged stringer. At each side of thedamage in the upstanding flange, insert and driveeight AN442-AD5 rivets. Rivet the splice memberthrough the skin with the same type and sizerivets originally used through the skin (seeFigure 21).18. STRINGER TYPE C27HT.The Type C274T T-angled bulb extrusion is employedas the leading edge stringer oh theupper surface of the centersect ion. The stringeris fonned to a 24ST extruded aluminum alloy shapehaving the equivalent Alcoa die number L23966.If the stringer is severely damaged, Cut out thedamaged material with a hack saw, locating thecut between two skin rivets at each side of thedamage. In order to splice the stringer properly,cut two lengths of C204T extruded sections, one havinga length of 9 inches plus the length of thedamage and another having a length of 12-3/4 inchesplus the length of the damage. Chamfer the outsidecomers of the spl ice members to fit the insideradius of the damaged stringer. Trim the bulbof the damaged stringer for a length of 4-1/2inches on each side of the damage in order tofit the C204T spl ice member on the one sideof the stringer. Drill out the affected skinrivets at each side of the damage and clampthe splice members to the positions indicatedL 109]RESTRICTED

surfaces-theI,.FIXED SURFACESPAR. 18 TO 19RESmiCTED•^-^ 12% PLUS LENGTH OF DAMAGE-LOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGEEXISTING SKINRIVETSAN442-AD5 RIVETS(10 REQ. ON EACH SIDE)TRIM BULB TO FITC204T SPLICE MEMBERON THIS SIDESit—2-C204T SPLICE MEMBERSxis:.093n-^vHEXISTING SKINRIVETShaving the equivalent Alcoa die number L249I0.If the stringer is severely damaged, cut out thedamaged material with a hack saw, locating thecut between two skin rivets at each side of thedamage. Prepare a f i I ler section of Type C366Textrusion to match the removed material. Cuta length of Type C366T extruded section equalto 9 inches plus the length of the damage (seeFigure 2q) . Clamp the splice member to thedanaged stringer. At each side of the damage,center punch six rivet locations in the upstandingflange at an average spacing of 3/4- inch oncenters. With a No. II ( . 19 I ) dri I dri I I thecenter-punched rivet locations. With a drillequal to the size of the existing rivet skinrivet holes, drill new rivet locations throughthe skin and splice member. Remove the splicemember and burr al I rivet holes. Apply onecoat of zinc chromate primer to all surfacesof the spl ice member. Again clamp the spl iceto the damaged stringer. In the upstandingflange at each side of the damage, insert anddrive six AN442-AD6 rivets. Using the samesize rivets as existing skin rivets, rivet thesplice member flange to the skin (see Figure 2g)Figure 28— Spl ice for Stringer Type C274T(see Figure 28). At each side of the damage,center punch ten rivet locations in the upstandingflanges at an average spacing of 5/8-inch on centers. With a No. 21 ( . 159) dri I I,drill the center-punched rivet locations. Witha drill equal to the size of the existing holesin the skin, redrill the skin holes throughthe splice members. Remove the splice membersand burr all the rivet holes. Apply one coatof zinc chromate primer to al I of thesplice members. Again clamp the splice membersto the damaged stringer. At each side ofthe damage in the upstanding flanges, insertand drive ten AN442-AD5 rivets. Secure thesplice members to the skin with the same typeand size rivets previously drilled out (seeFigure 28) .19. STRINGER TYPE C366T.The Type C366T stringer is employed extensivelyin the upper surface of the outer wingand in several locations in the upper surfaceof the centersect i on. The stringer is a bulbangle extrusion of 24ST aluminum alloy materialEXISTINGSKIN RIVETS-LOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGE.BUTT FITC366T FILLER.062Figure 29— SpliceM^Re^Q~rAN442-AD6RIVETS (6 REQ.ON EACH SIDE)EXISTINGSKIN RIVETSCI37T OR C366TSPLICE MEMBERSAME SIZE ANDSPACING AS EXIST-ING SKIN RIVETSfor Stringer Type C366TRESntlCTED[110]

I 1-1/2I fI equallerstheI,IIRESTRICTEDFIXED SURFACESPAR. 20 TO 2120. DOUBLED C366T TYPE STRINGERS.The doubled combination of C366T stringersis employed in the root of the upper surfaceof the outer wing. It is to be noted that usuaMy only a short length of C366T stringeris used to reinforce another C366T stringer in theroot of the outer wing. If damage occurs atthe root of these stringers, replace the entirelength of doubled stringers, which is usuallynever longer than one or two bays in length,and splice the remaining C366T stringer asoutlined in the preceding paragraph. Replacementof this damaged combination is recommended;but the splice outlined herein may be employedH^/sf^HLOCATE CUTBETWEEN 2SKIN RIVETSON EACH SIDEOF DAMAGE.BUTT FITC366T FILLERSUy^ PLUS LENGTH OF DAMAGE5XFigure 30 Splice forAN442-AD5 RIVETS(9 REQ. ON EACH SIDE)EXISTING SKIN RIVETSK77A SPLICEMEMBERS. MODIFYAS SHOWNDoubled C366TType Stringerswhen the C366T combination is longer than twobays in length and immediate repair is necessary.This condition occurs only in one locationon the upper surface of the right outer wingof the AT-6B and AT-6C Airplanes, just aft ofthe gun compartment (see Figure ii). Proceedas follows: Cut out the damaged material,locating the cut between two skin rivets ateach side of the damage. Prepare fillers ofC366T material to match the removed sectionof the stringer. For the splice members, cuttwo lengths of Type K77A extruded sections,each equal to the length of the damage plusinches (see Figure 30). The skin rivetsat the ends of the spl ice must have at least3/8- inch edge distance and this may necessitatemaking the splice members slightly longer thanthat noted. Cut and modify the splice membersand chamfer the outside corners of thesplice members to fit the inside radius of thestringers. Drill out the affected skin rivets.lers to the damagedClamp the spl ice rrembers and f i Istringers. At each side of the damage in theupstanding flanges, center punch nine rivetlocations at an average spacing of 5/8-inchon centers. With a No. 21 ( . 159) dri Idri Ithe center-punched rivet locations. With adrill equal to the size of the existing skinrivet holes, drill through the splice membersand fillers, using the existing rivet holes asa guide. Remove the spl ice members and fi I lersand burr all the rivet holes. Apply one coat ofzinc chromate primer to all surfaces of the splicemembers and reclamp the spl ice members and f i I lersto the stringers. At each side of the damage inthe upstanding flanges, insert and drive nineAN442-AD5 rivets. Replace the skin rivets withrivets of the same size and type as previouslyremoved. Rivet the fi to the stringersIas required (see Figure 30).21. STRINGER TYPE C373LT.The Type C373LT stringer is employed in theconstruction of the stabilizers of the AT-6Band AT-6C Series Airplanes. The stringer isformed of rolled alclad sheet .025 inch thick.the damage is si ight, spl ice members 7-1/2inches long will suffice (see Figure 31). Ifthe damage is extensive, cut out the damagedmaterial, locating the cut between two skinrivets at each side of the damage. Bend up thetwo .032 inch thick splice members, each havinga length equal to 7-1/2 inches plus the lengthof the damage (see Figure 31). Drill out theaffected skin rivets at each side of the damage.Clamp the splice members to the stringer. Inthe upstanding flange, center punch five rivetlocations at an average spacing of 3/ 4^ inch oncenters. With a No. 30 (. 1285) dri 1I, dri Ithe center-punched rivet locations. With adri to the size of the existing skinIrivet holes, drill through the splice members,using the existing rivet holes as a guide. Removethe splice members, and burr al Irivetholes. Apply one coat of zinc chromate primerto all overlapping surfaces. Reclamp the splicemembers to the damaged stringer. In the upstandingflange of the stringer at each side ofthe damage, insert and drive five AN442-AD4rivets. Rivet the splice member to the skinwith the same size and type rivets previouslyremoved (see Figure 31).LIN JRESTRICTED

.OFI equalI lerI outI.FIXED SURFACESPAR. 22 TO 23RESTRICTED-*-W\/\/^— 71^2"-UOCATE CUTBETWEEN 2SKIN RIVETS ONEACH SIDE OFDAMAGEPLUS LENGTH OF DAMAGEI." ,".^4\k;'R8.025"EXISTING SKIN RIVETSAN442-AD4 RIVETS( 5 REQUIRED ON EACH SIDE )032- INCH 24ST ALCLADFigure Sl—Splice for Stringer Type C373LT11, STRINGER TYPE K77A.The Type K77A Alcoa stringer is employed inseveral locations in the lower surface of theouter wing. The stringer is formed of 24STaluminum alloy. If the damage is extensive,cut out the damaged material, locating the cutbetween two skin rivets at each side of the\Z\ PLUS LENGTH DAMAGE_I* 6^/6 -7(1 Q Q Q Q Q Q G O QJl .:; . .. ibH^HLOCATE SPLICEBETWEEN 2SKIN RIVETS ONEACH SIDE OFDAMAGE. BUTTFIT K77A FILLEREXISTING. . ,. T ; _." . _.. , .....- -,,.,,fe^SKIN RIVETS-^^8TtH'/akAN442-AD5 RIVETS(10 REQUIRED ON BACH SIDE)EXISTINGSKIN RIVETS040 X '/a ALCLAD SHEETK77A SPLICE MEMBERSAME SIZE ANDSPACING AS EXISTINGSKIN RIVETSFigure 32^Splice for Stringer Type K77Adamage. Prepare a f i of K77A extrusion tomatch the removed material. For the splicemembers, cut a length of Type K77A extrusionequal to the length of the damage plus 12-3/4inches and bend up the same length of .040inch thick ale lad sheet (see Figure ^2). Ateach side of the damage, dri I the affectedskin rivets. Clamp the splice rrembers and fillersto the damaged stringer. At each side ofthe damage in the upstanding flange, centerpunch ten rivet locations at an average spacingof 5/P-inch on centers. With a No. 21 (.159)drill, drill the center-punched rivet locations.With a drill equal to the size of theexisting skin rivet locations, redrill skinrivet holes and drill new rivet locations. Removethe splice members and burr the rivet holes.Apply one coat of zinc chromate primer to allsurfaces of the extruded splice member andto the overlapping surfaces of the ale lad sheetsplice member. Reclamp the splice membersto the damaged stringer. At each side of thedamage in the upstanding flange, insert anddrive ten AN442-AD5 rivets. Replace the skinrivets with rivets of the same size and typeas previously used (see Figure 32).23. DOUBLED STRINGER TYPES K77A.The Alcoa doubled K77A stringers are employedonly in the upper surface of the centersection.Each stringer is formed of 24STaluminum al loy extruded stock. If the damageis extensive, cut out the damaged material,locating the cut between two skin rivets ateach side of the damage. Prepare fillers ofK77A extrusion to match the removed aiaterial.For the splice members, cut two lengths of K77Aextrusion, each equal to the length of the damageplus 12-3/4 inches (see Figure 33J . Skinrivets at the ends of the spl ice must have atleast 3/P-inch edge distance and this may necessitaterraking the splice members slightly longerthan that noted. Drill out the affected skinrivets at each side of the damage and chamferthe outside corners of the splice members tofit the inside radii of the stringers. Clampthe splice members and fillers to the properpositions on the stringers. At each side ofthe damage in the upstanding flanges, centerpunch ten rivet locations at an average spacingof 5/B-inch on centers. Drill the centerpunchedrivet holes with a No. 21 ( . 159) dri IUsing a dri I to the size of the existingskin rivet holes, redrill the skin rivetholes through the splice members and fillers.Remove the splice members, and burr and cleanthe rivet holes. Reclamp the splice membersRESTRICTEDL 112]

^I,ingI outskinRESTRICTED-—12^" PLUS LENGTH OF DAMAGE6^8"FIXED SURFACESPAR. 23 TO 26AN442-AD5 RIVETS(10 REQUIRED ON EACH SIDELOCATE CUTBETWEEN 2EXISTING SKIN RIVETSSKIN RIVETSAT EACH SIDE OFK77A SPLICE MEMBERSDAMAGE. BUTTFIT K77A FILLERSAFigure JS— Splice for Doubled K77AType Stringersand the fi llers to the damaged stringers. Ateach side of the damage in the upstanding flanges,insert and drive ten AN442-AD5 rivets. Replacethe skin rivets through the splice membersand f i I lers with rivets of the same typeand size as previously used (see Figure 33).LS1I27-5-6>>r

'I equal7F IXED SURFACESPAR. 25 TO 27RESTRICTEDfrom the affected skin panel. In many cases,the rib attaching angles on the spar must beremoved in the affected area and then replacedafter the spar repair is completed. Modifythe rib attaching angles as required to permitreplacement.26. OUTER WING MAIN SPAR - ROOT TO 106 INCHESOUTBOARD.Depending upon the extent of damage, allor part of the complete splice illustratedmay be used (see Figure ^6). For example,if only the spar cap has been damaged, useonly the spar cap part of the spl ice; if onlythe spar web is damaged, use a .064 inch thicksheet of 24ST alclad overlapping the trimmedoutround hole by 1-1/4 inches all around,secured by two rows of AN442-AD6 rivets at7/8-inch spacing in each rcw. However, theprocedure for a complete splice rmy be summedup as follows: With a spiral reamer, cut outthe damaged area to a round-cornered shape orcut out the entire damaged spar section. Forthe spl ice members, cut two flat sheets of .064inch thick 24ST alclad having a length of 10-1/2 inches plus the length of the damage and anapproxitrcite width of 8-3/4 inches. Along thelength of each of the splice members, brake a 1/2-inch right-angled flange; then brake a T-3/8inch right-angled flange measured from thefirst bend. A hand brake should be used ifavailable, or the sheet may be clamped betweentwo blocks of wood and bent by means of handpressure; but in either case, obsen/e a minimumbend radius of 1/4-inch. Trim the splice membersto the shape shown (see Figure 36). Ateach side of the damage, drill out the affectedskin rivets with a No. 50 (.1285) drill, tak-I iing particular care to avoid elongating theexisting holes. For the web stiffeners, cutand bend four l-f/4 inch wide strips of .064inch thick 24ST alclad equal to the spar depthat the affected area. F^nd up any additionalweb stiffeners to replace existing ones whichrTBy have been damaged. Securely clamp the splicemembers and web stiffeners to the spar. Centerpunch the required rivet locations as noted,and drill these locat ions with a No. II (.191)drill. It is to be noted that all rivets atthe ends of the spl ice must have at least 3/8-inch edge distance, measured from the centerneof the hole. With a dri I to theexisting skin rivet holes, drill the splicemembers through the existing skin rivet holes.Remove the splice members and burr all therivet holes. Apply one coat of zinc chromateprimer to all overlapping surfaces of the splicemembers and web stiffeners and reclamp themembers to the proper positions on the spar.Into the No. If (. 191 ) holes, insert and driveAN442-AD6 rivets. To prevent the sheet fromcreeping, rivet first at the corners and thenat tne intermediate positions. On each sideof the damage through the upper skin, insertand drive at least ten AN426-AD rivets of thesame size as previously used. On each sideof the damage through the lower skin, insertand drive at least ten B!227-AD rivets of thesame size as previously used (see Figure 36).27. OUTER WING MAIN SPAR -OUTBOARD OF ROOT.106 TO 151 INCHESIf only the spar cap has been damaged, useonly the spar cap part of the splice; if onlythe spar web is danaged, use a .040 inch thickpatch of 24ST alclad overlapping the hole byt- STA.OUTER. WING, ROOTROLLED SHEETANGLE WEB•STIFFENERS106 INCHES OUTBOARD.064 INCH THICKALUM. ALLOY SHEET SPARALCLAD ON AT-6B aC)|151 INCHESOUTBOARDEXISTINGSPAR SPLICE\7SPAR—TIP VBENTUP SHEET CAPS.040 INCH THICKALUM. ALLOY SHEET SPAR(ALCLAD ON AT-6B 8 0)LIGHTENINGHOLESFigure 35—Outer Wing Main SparRESTRICTED[114]

RESmiCTED FIXED SURFACESFIGURE 36AN 426 SKINRIVETS (10 REQ.)SISKIN(REF)TOP V/EW (SKIN REMOVED)064 SPLICE PLATE24ST ALCLAD (2 REQ)AN 442-AD6RIVETS (12 REQ.)/-.064 X 1-1/4" X SPAR DEPTH/ 24ST ALCLAD SHEET ANGLE (4 REQUIRED)^Ctl flr> «-» in im

2-inchI overlappingI outFIXED SURFACESPAR. 27 TO 28RESTRICTEDII-I/4 inches all around. Secure the patch withtwo rows of AN442-AD5 rivets at 3/4-inch spacingin each row. However, the procedure fora complete spl ice may be summed up as fol lews(see Figure 97^. With a spiral reamer, cutout the damaged area to a round -cornered shape,or cut out the entire damaged spar section.For the splice members, cut two flat sheetsof .040 inch thick 24ST ale lad having a lengthof 7-1/2 inches plus the length of the damageand an approximate width of 6-1/4 inches.Along the length of each of the splice members,brake a right-angled flange,and then brake a 1-3/8 inch right-angled flangemeasured from the first bend. A hand brakeshould be used if available, or the sheet maybe clamped between two blocks of wood and bentby means of hand pressure; but in either case,observe a minimum bend radius of 5/32-inch.Trim the splice members to the shapes shewn (seeFigure 97J. At each side of the damage, drillout the affected skin rivets with a No. 40( .098) drill, taking particular care to avoidelongating the existing holes. For the webstiffeners, cut and bend up four lengths of1-3/8 inch strips of .051 inch thick 24ST alclad equal to the spar depth at the affectedarea. Bend up any additional web stiffenersto replace existing ones which may have beendamaged. Securely clamp the splice membersand web stiffeners to the proper positions onthe spar. Center punch the required rivet locationsas noted, and drill these locations witha No. 2! (.159) drill. With a No. 30 (.1285)drill, drill the splice members through theexisting skin rivet holes. It is to be notedthat all rivets at the ends of the splice membersmust have an edge distance of at least3/8-inch measured from the centerline of thehole. Remove the splice members and web stiffenersand burr all rivet holes. Apply one coatof zinc chromate primer to a Isurfaces.Rec lamp the members to the proper positionson the spar. Into the No. 21 holes,insert and drive AN442-AD5 rivets. To preventthe sheet from creeping, rivet first atthe corners and then at the intermediate positions.On each side of the damage throughthe upper skin, insert and drive at least eightAN426-AD4 rivets. On each side of the damagethrough the lower skin, insert and driveat least eight BI227-AD4 rivets (see Figure 37;.28. OUTER WING MAIN SPAR - 151 INCHES OUT-BOARD TO TIP.Darmge to the outer panel main spar from 151inches outboard to the tip of the wing may berepaired as shown (see Figure ^8). The spliceis applicable only to areas between flangedweb cutouts. Where flanged web cutouts arepresent in the affected area, trim the doublersaround the web cutouts as requ i red,makingcertain that the required number of rivets ispicked up at each side of the damage. Witha spiral reamer, cut out the damaged area, locatingthe cut between two skin rivets at eachside of the damage. For the splice members,cut two flat sheets of .040 inch thick 24STale lad, having a length of 6 inches plus thelength of the damage and having an approximatewidth of 4-7/8 inches. It is to be notedthat if the damage occurs -in areas adjacentto web cutouts, the length of the spl ice membersmay have to be increased to pick up therequired rivets at each side of the damage.Along the length of each of the splice members,brake a 1/4-inch right-angled flange,and then brake a 1-inch right-angled flangemeasured from the first bend. A hand brakeshould be used if available, orthe sheet maybe clamped between two blocks of wood and bentby means of hand pressure; but in either ease,a minimum bend radius of 5/32- inch should beobserved. Trim the spl ice rremberB to the shapesshown (see Figure 98J , or trim the membersaround cutouts as required. At each side ofthe damage, dri I the affected skin rivetswith a No. 40 (.098) drill, taking particularcare to avoid elongating the existingholes. For the web stiffeners, bend uptwo 1-3/8 inch strips of .051 inch thick 24STa Id ad equal to the spar depth at the affectedarea. Bend up any additional required webstiffeners to replace existing ones which mayhave been damaged. Securely clamp the splicemembers and the web stiffeners to the properpositions on the spar. Center punch the requiredrivet locations as noted and drill theselocations witha No. 21 (.159) drill. Witha No. 30 (.1285) drill, drill the spl ice merr^bers through the existing skin rivet holes.It is to be noted that all rivets at the endof the splice must have an edge distance of3/8-inch measured from the centerline of thehole. Remove the spl ice members and the webstiffeners, and burr all rivet holes. Applyone coat of zinc chromate primer to all overlappingsurfaces. Again clamp the members tothe spar. Into the No. 21 (.159) holes, insertand drive AN442-AD5 rivets. To preventthe sheet from creeping, rivet first at thecornerB and then at the intermediate positions.On each side of the damage through the upperskin, insert and drive at least four AN426-AD4rivets. On each side of the damage throughthe lower skin, insert and drive at least fourBI227-AD4 rivets (see Figure ^8).RESTRICTEDLm6]

RESTRICTEDFIXED SURFACESFIGURE 37AN426-AD4RIVETS (8REQ.)^^SKIN(REF)AN442-AD5RIVETS (lOREQ)040 SPLICE PLATE, 24 ST ALCLAD(2 REQ).051 X 1-3/8 ANGLE(4 REQ)sn OD fo anJ k .lE^—flD ID C~)L2 VERTICAL ROWS OFAN 442-r"ADS RIVETSAT 3/4" O.C.SPAR^FILLERTOP VIEW (SKIN REMOVED)AN 442- ADSRIVETS (10 REQ.),1*^BI227-AD4 RIVETS(8REQ.)-SKIN(REF)SIDEVIEWNOTE; ALL RIVETS REQUIREDMUST BE USED ONEACH SIDE OF DAMAGEFigure 37— Outer Wing Main Spar Splice - 106 to 151 Inches Outboard of RootLM7jRESTRICTED

AFIXED SURFACESFIGURE 38RESTRICTED_LiB^ -dXL ^^-,051 XIV. ANGLE(2 REQ)SPAR.040 DOUBLER24ST ALCLAD(2 REQ.)SECTION A-SKIN(REF.)AN442-AD5RIVET (4 REQ.)2 ROWS\ AN442-AD5RIVET (8 REQ.)NOTEALL RIVETS REQUIREDMUST BE USED ATEACH SIDE OF DAMAGEAN442- ADSRIVET (4 REQ.)SKIN(REF.)BI227- AD4RIVET (4 REQ.)Figure 38-~0uter Wing Main Spar Splice - 151 Inches Outboard to TipRESTRICTED[ 118]

6-I,/8-inchin the outer wing trailing edge structure inboardof the aileron) may be repaired as shown(see Figure ggj . With a spiral reamer, cutout the damaged material, locating the cut betweentwo skin rivets at each side of the damage.For the splice members, cut two flatsheets of .040-inch 24ST ale lad, having a lengthof 6-1/2 inches plus the length of the damageand an approximate width of 3-1/2 inches.It is to be noted that, if the damage occursbetween two lightening holes in the spar, thelength of the splice members may have to beextended to pick up the required number of rivets.Along the length of the upper spl ice member,brake a 77-degree, 3/4- inch flange. Alongthe length of the lower splice member, brakean 86-degree, 7/ 1 Inch flange. A hand brakeshould be used if available, or the sheet maybe clamped between two blocks of wood and bentby means of hand pressure; but in either case,a minimum bend radius of 1/8-inch should beobserved. Trim the splice members so that theywill fit around the lightening holes. At eachside of the damage, drill out the affected skinrivets with a No. 40 (.098) drill, taking particularcare to avoid elongating any existingholes. For the web stiffeners, bend uptwo 1-1/4 inch strips of .032 inch thick 24STalclad equal to the spar depth at the affectedarea. If damage occurs to any of theextruded stiffeners employed at intervals alongthe spar length, remove and replace the stiffenerswith similar extrusions. Securely clampthe splice members and the web stiffeners tothe damaged spar. Center punch the requiredrivet locations as noted and drill these locationswith a No. 30 ( . 1285) drill. With thesame dri I I, dri 1 I the cap of the upper spl icemember through the existing skin rivet holes.With a No. 40 (.098) drill, drill the cap ofthe lower splice member through the existingskin rivet holes. It is to be noted that allrivets at the ends of the splice must have anedge distance of 3/8-inch measured from thecenterline of the hole. Remove the splice merr^bers and the web stiffeners, and burr all rivetholes. Apply one coat of zinc chromateprimer to all over lappi ng surfaces. Againclamp the members to the proper positions onthe spar. Into the spar web rivet holes, insertand drive AN442-AD4 rivets. Replace theskin rivets with BI227-AD4 rivets on the uppercap and BI227-AD3 rivets on the lower cap.To prevent the sheets from creep ing, rivetfirst at the corners and then at the inter-.RESTRICTED FIXED SURFACESPAR. 29 TO 3129. OUTER WING FLAP SPAR.mediate positions. This completes the repair(see Figure gojDamage to the outer wing flap spar (located30. OUTER WING AILERON SPAR.iDamage to the outer wing ai leron spar ( locatedin the outer wing trailing edge structureoutboard of the flap) may be repaired asshown (see Figure 40). With a spiral reamer,cut out the damaged material, locating the cutbetween two skin rivets at each side of thedamage. For the splice members, cut two flatsheets of .052 inch thick 24ST alclad, eachhaving a length of 7-1/2 inches plus the lengthof the damage and an approximate width of 4inches. It is to be noted that the skinrivets at the end of the splice must havean edge distance of 3/8- inch and this maynecessitate making the splice members slightlylonger than that noted. Along the lengthof the upper splice member, brake a 100-degree,3/4-inch flange. Along the length ofthe lower splice member, brake a 95-degree,3/4- inch flange. A hand brake should be usedif available, or the sheets may be clamped betweentwo blocks of wood and bent by means ofhand pressure; but in either case, a minimumbend radius of should be observed.Trim the splice members so that they will fitaround the lightening holes in the damagedspar. At each side of the damage, drill outthe affected skin rivets with a No. 40 (.098)dri I taking particular care to avoid elongatingthe existing rivet holes. For the webstiffeners, bend up two 1-1/4 inch strips of.032 inch thick 24ST alclad equal to the spardepth at the affected area. Securely clampthe splice members and the web stiffeners tothe proper positions on the damaged spar.Center punch the required rivet locations asnoted and drill these locations with a No. 30(. 1285) dri 1 I . With the same drill, dri IIthe splice members through the existing skinrivet holes. Remove the splice members andthe web stiffeners and burr all rivet holes.Apply one coat of zinc chromate primer to al Ioverlapping surfaces. Again clamp the membersto the proper positions on the spar. Into thespar web rivet holes, insert and drive AN442-AD4 rivets. Replace the skin rivets with BI227-AD4 rivets (see Figure 40). To prevent thesheets from creeping, rivet first at the comersand then at the intermediate positions.31. CENTERSECTION FLAP SPAR.Damage to the centersect ion flap spar (locatedin the centersect ion structure forward[ 119]RESTRICTED

FIXED SURFACESFIGURE 39RESTRICTEDBI227-AD4(3 REQ.)RIVETSSKIN (REF.)AN442-AD4(4 REQ.)RIVETSALL RIVETS REQUIRED ADDITIONAL AN442-AD4\0"'^^^^MUST BE USED AT^^^ REQUIRED-OF DAMAGEEACH SIDE2 VERTICAL ROWSOF AN442-AD4RIVETS AT 7/8ON CENTERSAN442-AD4 RIVETS(5 REQ.)BI227-AD3 RIVETS-\ ^(3 REQ.) \^pSECTION BBSKIN(REF.).032 X H/4 ANGLE, 24ST ALCLAD(2 REQ.)SPARI •^iiiiiimmw ,^«'!,lk»M,yi :\..." "" V iii "«r-m'SECTION AA.

ARESTRICTED FIXED SURFACESFIGURE ^0-rsa—fHHp^^H^^ mmrSPAR.032 X PA ANGLE24ST ALCLAD (3REQ.)032 X 4 SPLICE SHEET24ST ALCLAD (2 REQ.)SECTION A-ADDITIONAL AN442-AD4RIVETS AS REQUIREDEXISTING TYPELIGHTENING HOLEBI227-AD4 RIVETS(3 REQ.)AN442-AD4RIVETS (5 REQ.)NOTE:ALL RIVETSREQUIRED MUSTBE USED AT EACHSIDE OF DAMAGE2 VERTICALROWS OFAN442-AD4RIVETS AT3/4 ON CENTERSAN442-AD4RIVETS (5 REQ.)^BI227-AD4(3 REQ)RIVETSFigure 40— Outer Wing Aileron Spar SpliceL 121]RESTRICTED

orI outI.I ingFIXED SURFACESPAR. 31 TO 32RESTRICTEDof the flap) may be repaired as shown (see Figure41)' The complete splice of the member isshown; but depending upon the extent of damage,al I part of the complete spl ice may be used.With a spiral reamer, cut out the damaged material,locating the cut between two skin rivetsat each side of the damage. For the spl icemembers, cut two flat sheets of .032 inch thick24ST alclad, having a width of 3-5/16 inchesand sufficient length to pick up the requirednumber of rivets at each side of the damage.In the repair illustrated, the length of thesplice member is equal to the length of thedamage plus 10-1/2 inches. Along the lengthof the upper spl ice member, brake a 78-degree,3/4- inch flange. Along the length of the lowersplice member, brake an 86-degree, 7/16-inchflange. A hand brake should be used if available,or the sheets may be clamped between twoblocks of wood and bent by means of hand pressure.Observe a minimum bend radius of 1/8-inch. Trim the splice members so that theywill fit around the lightening holes. At eachside of the damage, dri I the affected skinrivets with a No. 40 (.098) drill, taking particularcare to avoid elongating the existingrivet holes. For the web stiffeners, bend upthree strips of .032 inch thick 24ST alclad,1-1/4 inches wide and 5 inches long. If damageoccurs to any of the extruded angles usedat intervals along the spar length, remove andreplace the angles with similar extrusions.Securely clamp the spl ice members and the webstiffeners to the spar. Lightly center punchthe required rivet locations as noted and drillthese locations with a No. 30 ( . 1285) driIWith the same drill, drill the cap of the uppersplice member through the skin rivet holes.With a No. 40 (.098) drill, drill the cap ofthe lower splice member through the existingskin rivet holes. Remove the splice membersand the web stiffeners, and burr all rivet holes.Apply one coat of zinc chromate primer to alloverlapping surfaces. Again clamp the membersto the proper positions on the spar. Into thespar web rivet holes, insert and drive AN442-AD4 rivets. Replace the skin rivets with BI227-AD4 rivets in the upper cap and BI227-AD3 rivetsin the lower cap (see Figure 41).32. CENTERSECTION REAR SPAR.Damage to the centersect ion rear spar (locatedin the centersect ion structure just aftof the fuel tank compartments) may be repairedas illustrated (see Figure 4^) . The completesplice of the member is illustrated; but dependingupon the extent of the damage, all orpart of the complete spl ice may be used. Forexample, if only the spar web is damaged, roundout the hole to an oval shape; prepare an .051inch thick 24ST alclad patch with a 1-1/4 inchoverlap and secure the patch with two rows ofAN442-AD5 rivets at 3/4-inch spacing all aroundthe hole. With a spiral reamer or a hack saw,cut out the damaged material, locating the cutbetween two skin rivets at each side of thedamage. Remove any damaged stiffeners in thearea. With a No. 30 ( . 1285) dri I I, dri I I outthe affected skin rivets at each side of thedamage, taking particular care to avoid elongatingany existing rivet holes. On the lowercap of the rear spar on the AT-6A Series Airplanes,drill out the rivets securing the existingBI083 stop nuts in the affected area. Onthe AT-6B and C Series Airplanes, the singlesteel nut channel assembly ( NA O/g. 84-13332)spanning the entire length of the rear spar lowercap must be removed by dri I out the securingrivets. Remove the channel and, with a hack saw,cut out the damaged section of the channel 5-1/4inches from each side of the damage betweentwo of the channel nuts. At each side of thedamage, rerivet the two remaining sections ofthe nut channel to the lower spar cap, discardingthe damaged section. To replace the damagedsection of the nut channel, cut a lengthof 22GCH-048 steel channel to match the damagedsection and snap in the required numberof 22G 1-048 nuts. (Channel and nuts may be purchasedfrom the Boots Nut Corporation, Waterbury,Conn. ) For the spl ice members, cut twosheets of .051 inch thick 24ST alclad having alength equal to the length of the damage plus10-1/2 inches and a width of 7-3/16 inches. Theskin rivets at the ends of the spl ice must havean edge distance of at least 3/8-inch and thebolts must have 1/2- inch edge distance; this maynecessitate making the splice members slightlylonger than that noted. Along the length ofthe upper spl ice member, brake a 90-degree,1/2- inch flange and then brake a 99-degree,1-3/4 inch flange measured from the first bend.Along the length of the lower splice member,brake a 90-degree, 1/2- inch flange and thenbrake a 92-degree, 1-3/4 inch flange measuredfrom the first bend. A hand brake should beused if available. Observe a minimum bend radiusof 3/16-inch. Trim the splice members to theshapes shown (see Figure 4q) . For the extrudedstiffeners, cut two lengths of Type C204T extrusion,each having a length of 9-5/8 inches.Cut similar lengths of the extrusion to replaceany of the existing stiffeners which may havebeen damaged. Clamp the splice members and theweb stiffeners to the proper positions on theRESTRICTEDI 122]

RESTRICTED FIXED SURFACESFIGURE 41032 X 1-1/4 X 5 ANGLE24 ST ALCLAD (3 REQ,)SPARiJ—rV^^^arm^^im^^mSECTIONAA.032 X 3-5/1624 ST ALCADSPLICE SHEET(2 REQ.)SKIN (REF.)(REQ)ALL RIVETS REQUIREDMUST BE USED AT EACHSIDE OF DAMAGEB I227-AD3 RIVETSSECTION j^nFigure 41— Center sect ionFlap Spar SpliceLl23]RBSTRICTtD

FIXED SURFACESPAR . 32 TO 33.051-INCH THICK24ST ALCLAD SHEETmSTRICTED^'4REARVIEW

RESTRICTED FIXED SURFACESFIGURE 43SPARFILLERBI227- ADS RIVETS( 10 REQUIRED ON EACHSIDE OF DAMAGE )SKIN (REF.)PLUS LENGTHOF DAMAGESKINFUEL TANK COMPARTMENT(REF.)DOORBI25I -428-14 SCREW(3 REQUIRED ON EACHSIDE OF DAMAGE)BI227-AD5 RIVETS( 5 REQUIRED ON EACHSIDE OF DAMAGE)AT AFFECTED AREA, REPLACEEXISTING 22 GCH - 048 ( BOOTS)STEEL NOT CHANNEL AND22GI-048 NUTS ON AT-6BaC.SECURE CHANNEL WITH AN426-AD4RIVETS (BI083 STOP NUTS ONAT-6A SECURE WITH AN426-AD3RIVETS )C 204T-9-fO-STIFFENER(2 REQUIRED)BOTTOMVIEWFigure 43— Centersect ion Rear Spar Splice[125]RESTRICTED

holes032I twentyinerecessedFIXED SURFACESRESTRICTEDPAR. 33 TO 3>^Icap, and dri I No. M ( . 191 ) holes throughthe plate, using the existing holes as a guide.Al at the ends of the splice must havean edge distance of at least 3/ 8- inch measuredfrom the center I of the hole to the edge ofthe material involved. Drill additional holesler to the skin.as requi red to rivet the f i IFtemove the splice members and burr all the rivetholes. Apply one coat of zinc chromate primerto all surfaces of the spl ice members and f i I ler.Again secure the members to the damaged sparcap; and nto the e ighteen No. II (.191) holesithrough the skin, insert BI25I-I052-I4 screwsand secure with AC365-I032 nuts. Through thetwenty No. II (.191) holes in the heel of thecap, insert B 1251-1032-17 screws and securewith AC365-I032 nuts. In areas where flushrivets were originally employed, substitute81 248- flush screws1for 81 25 1buttonhead screws, after cut countersinking the holes100 degrees x 5/8-inch diameter. Tighten allnuts with an approximate torque of from 20 to25 inch-pounds. Bolt threads must not bear inany part of the splice members. Although thescrew threads must cut completely through thefiber of the self-locking nut, there should beno more than approximately two screw threadsshowing beyond the fiber. Normally, no washersshould be necessary under the nuts; but if theuse of washers is necessary in certain locationswhere the screw passes through a web stiffener,use AN960-I0 washers in the quantity required,and use the next greater length screw. Rivetthe C245T spar filler to the splice memberswith the same size and type rivets as originallyemployed in the area (see Figure 4^).3^. CENTERSECTION FRONT SPAR LOWER CAP.In an emergency, damage to the centersect Ionfront spar lower cap (formed of Type C245Textrusion) may be repaired as shown (see Figure46). It is to be noted that instructionsoutlined herein cover damage to the lower extrudedspar cap only. If the spar web is damaged,reference should be rrade to the paragraphscontaining information pertinent to an applicablerepair. This spar cap repair should beused only in an emergency where immediate repairis necessary. Replacement of a severelydamaged spar cap is recommended. Remove the1/4-inch diameter bolts securing the fuel tankcompartment doors. If the cap has been completelysevered, cut out the damaged cap materialwith a hack saw, locating the cut betweentwo fuel tank bolt holes at each side of the064- INCH THICK 24ST ALUM.ALLOY SHEET WEB:-\:S'o :;FRONT VIEW66-13198LOWER CAP49-13105-2UPPER CAP66-13090FITTINGSREARVIEWREF. DWG. 77-13003 AT-6AaB88-13003 AT-6CFigure 44- Centersect ion Front SparRESTRICTEDL 126]

1t|IRESTRICTED FIXED SURFACESPARAGRAPH 34SKIN(REF)TOPVIEW-DRILL NO.II CI9I) DIA. 38 HOLES THROUGH. BI25I - 1032- 14 SCREW (18 REQ.)B 1251 -1032- 17 SCREW (20 REQ.) (USE B 1248-1032 SCREWS IN AREAS EMPLOYINGFLUSH RIVETS) AC365- 1032 NUT (38REQ.)H>^£SKIN_(REF)-^C 245T Fl LLER*-2'30'(REF.)^2 (RER)C245T CAP(REF215 (REF.)SPAR WEB (REF.)END VIEW\^USE EXISTINGSPACING9-^ PLUS LENGTH OF DAMAGE. APPROX.C 245 T SPLICE MEMBER . MODIFY AS SHOWN-FRONTVIEWSPARWEBFigure 45— Centersect ion Front SparUpper Cap Splicedamage. Prepare a f i I ler section of Type C245Textrusion to match the removed material. Forthe splice members, cut two lengths of TypeC245T {KI6862 Alcoa die) extruded sections.The lengths of the C245T extruded members mustbe determined by the existing rivet spacingat the affected area multiplied by the numberof attachments required at each side of thedamage. In most cases, the length of the spliceshould not exceed 13-1/2 inches plus the lengthof the damage. On one of the extruded splicemembers, open the existing angle of 92 degrees30 minutes to an angle of 95 degrees.On the other splice member, cut off one legto form a 1-3/8 inch wide splice plate. Witha No. 21 (.159) drill, drill out the fourteenrivets at each side of the damage which securethe spar cap to the web. Take particular careto prevent the elongation of any of the rivetholes. If any rivets are used through theskin side flange of the spar cap, dri I I therivets out with a No. 30 (.1285) drill. Onthe AT-6A Series Airplanes, drill out the rivetssecuring the existing BI083 stop nuts inthe affected area. On the AT-6B and AT-6CSeries Airplanes, the single steel nut channelassembly ( KA DA/g. 84^13330) spanning the entirelength of the cap must be removed by drillingout the securing rivets. Remove the channel;and with a hack saw, cut out the damaged sectionof the nut channel 6-3/4 inches from eachside of the damage between two of the channelnuts. At each side of the damage, rerivetthe two remaining sections of the nut channelto the cap, discarding the damaged section.To replace the damaged section of the nut channel,cut a length of 22GCH-048 steel channelto match the damaged section and snap in therequired number of 22GI-048 nuts. (Channel andnuts may be purchased from the Boots Nut Corporation,Waterbury, Conn.) Clamp the C245Tangle splice member to the proper position onthe spar cap. With a No. F ( .257) dri I I, dri Ithe ten fuel tank compartment door bolt locations,using the existing holes in the capas a guide. With a No. II (.191) drill, drillthe 34 holes as shown (see Figure 46), usingthe existing rivet holes in the cap as a guide.Clamp or bolt the modified C245T splice memberplate to the heel of the cap and drill therequired holes through. All holes at the endsof the splice must have an edge distance ofat least 3/8-inch measured from the centerlineof the hole to the edge of the materialinvolved. Drill additional holes as requiredto secure the f i I ler. Remove the spl ice members[127]RESTRICTED

I through25-428-7IFIXED SURFACESPAR. 31 TO 35—RESTRICTED)iand burr all the rivet holes. Apply onecoat of zinc chromate primer to all surfacesof the splice members and filler. Again securethe members and filler to the spar; andnto the twenty-e ight No. M ( . 191 ) ho les inthe heel of the cap, insert Bl 25 1-1032-1 7 screwsand secure with AC365-I032 nuts. Into the sixNo. II (.191) holes in the skin side flangeof the cap, insert BI25I- 1032-1 4 screws andsecure with AC365-I032 nuts. Tighten all nutswith an approximate torque of from 20 to 25inch-pounds. Bolt threads must not bear inany part of the splice members. Although thescrew threads must cut completely through thefiber of the self-locking nut, there shouldbe no more than approximately two screw threadsshowing beyond the fiber. Normally, no washersshould be necessary under the nuts; but if theuse of washers Is necessary in certain locationswhere the screw passes through a webstiffener, use AN960- 10 washers in the quantityrequired and use the next greater length screw.On the lower cap of the rear spar on the AT-6A Series Airplanes, line up BI083 stop nutswith the No. F (.257) holes and flush rivetthe stop nuts to the splice member with AN426-AD3 rivets passed through the existing rivetholes In the skin. On the AT-6C Series Airplanes,instead of the BI0R3 stop nuts, flushrivet the replacement 22GCH-048 steel channelto the splice member with AN426-AD4 rivetspassed through the existing rivet holes In theskin. When installing the fuel tank compartmentdoors, pass B 1 I 1 screws throughthe splice members; five of these screws arerequired at each side of the damage (see Figure46).35. PATCHING CENTERSECTION FRONT SPAR WEB.Damage to the centersect Ion front spar webmay be repaired in most cases by patching; orIf the damage is severe, the web may be completelyspl Iced as out I i ned in the f o Mowi ngparagraph. Both types of repairs are illustrated(see Figure 4^). If the spar web is to bepatched, carefully scribe a circle around thedamage and with a spiral reamer cut around thescribe line. With a file, smooth out the roughedges of the material. Cut two round sheetsof .064 inch thick 24ST ale lad that will overlapthe spar web hole by 1-1/2 inches alaround. Position one of the patches on eachside of the spar web. With a No. II ( . 191drill, dri I the patches and spar webin several locations and temporarily fasten@-DRILL NO F (257) DIA 10 HOLES THROUGHBI25I-428-I7 SCREW (10 REQ.)® @^ ®./+ ©SAME SIZE AND TYPE ATTACHMENTPREVIOUSLY USED^® @SKIN^(REF)Tgp igil \Qt — (^yp lyf'" ''iij rO—tgnptgily"14J —tgi—[grBOTTOM VIEW \AT AFFECTED AREA, REPLACEEXISTING 22 GCH-048 (BOOTS)STEEL NUT CHANNEL AND 2201-048NUTS ON AT-6B S C. SECURE CHANNELWITH AN426-AD4 RIVETS.(81083 STOP NUTS ON AT-6A. SECUREWITH AN426-AD3 RIVETS)-DRILL NO. II (.191) DIA. 34 HOLES THROUGHBI25I-I032-I4 SCREW (6 REQ)BI25I-I032-I7 SCREW (28 REQ)AC365-I032 NUT (34 REQ)-USE EXISTING SPACING— 6i APPROX4-C245T SPLICEMEMBER ONELEG REMOVED13^ PLUS LENGTH OF DAMAGE, APPROXnlk^-^k„* __.„„ii..:. j t,, —SPAR WEB (REF)£/V0 VIEWC245T FILLERFRONT VIEWFigure 46—Centersect ion Front Spar Lower Cap SpliceRESTRICTEDL 128]

RESTRICTED FIXED SURFACESFIGURE ^7p /-AN442-AD6 RIVETS(AS REQUIRED).064 PATCHES,24 ST ALCLAD(2 REQUIRED)Figure 47 — Center sect ionFront Spar Web RepairL 129]RESTRICTED

inchinchI theFIXED SURFACESPAR. 35 TO 38RESTRICTEDthe patches in place with skin fasteners. Drillthrough the remaining holes in the patches atan average spacing of I on centers. Renxjvethe patches and burr all the rivet holes.Apply one coat of zinc chromate primer to al I surfacesof the patch sheets. Again secure themembers to the proper positions on the spar web.Rivet the patch sheets to the spar web in a doubleshear joint, using AN442-AD6 rivets in the quantityrequired. If any of the existing web stiffenersin the area are damaged, replace them.36. SPLICING CENTERSECTION FRONT SPAR WEB.in cases where patching the spar web willnot suffice, the web sheet may be completelyspl iced as shown (see Figure 47) . It is tobe noted that the repair outlined herein coversdamage to the spar web only. If the extrudedspar cap is damaged, reference should be madeto the paragraph containing applicable repairinformation. With a hack saw, make two verticalcuts on each side of the damage, and removeand discard the damaged material. Forthe spl ice members, cut two sheets of .064inch thick 24ST a Id ad. One of the spl ice sheetsshould have a width of 3-1/2 inches plus thewidth of the damage, and a length of 9-1/4inches. The other spl ice sheet should have awidth of 3-1/2 inches plus the width of the dam-Iage, and a length of 12-7/32 inches. With a No.21 (.159) drill, drill out the affected rivetsin the area, and position one of the patcheson each side of the damage. Dri I I the spl icesheets through in several locations and temporarily fasten the sheets in place with skinfasteners. Center punch the locations for tworivet rows at each side of the damage at anaverage spacing of on centers. With aNo. II (.191) drill, drill the center-punchedlocations. Remove the splice members and burrall the rivet holes. Apply one coat of zincchromate primer to all overlapping surfacesof the spl ice members. Again fasten the membersto the spar web. Rivet the spl ice membersin a double shear joint to the spar web withtwo rows of AN442-AD6 rivets (52 required)at each side of the damage. If any of thespar web stiffeners in the area are damaged,remove them and replace with new stiffeners.37. VERTICAL STABILIZER FRONT SPAR.Damage to the front spar of the vertical stabilizermay be repaired as shown (see Figure48). Depending upon the extent of the damage,all or part of the complete splice illustratedmay be used. If the damage is extensive,cut out the damaged material, locatingthe cut between two skin rivets at eachside of the damage. Use a sprial reamer ora hack saw. For the spl ice members, cut twosheets of .032 inch thick 24ST alclad, eachhaving a length equal to the length of thedamage plus 6 inches, and a width of 3-1/4 inches.Along the length of each of the spl ice members,bend up a 3/4-inch flange, observing aminimum bend radius of 1/8-inch. Trim thesplice members to the required shape; and ifany flanged lightening holes are present inthe affected area, trim the spl ice members tomatch the existing lightening holes if necessary.For the web stiffeners, bend up two.032 inch thick 24ST alclad strips equal tothe spar depth in length and 1-3/8 inches wide.Drill out the affected skin rivets at each sideof the damage. Clamp the spl ice members tothe proper positions on the spar, and drillthrough and temporarily fasten the splice memberswith skin fasteners. At each side of thedamage, center punch the required rivet locationsat an average spacing of 3/4- inch oncenters. It is to be noted that all rivetsat the ends of the splice must have an edgedistance of at least 3/8-inch measured fromthe centerl ine of the rivet hole. With a No.30 (.1285) drill, drill center-punched rivetlocations at each side of the damage. Witha No. 40 ( .098) drill, dri I spl ice membersthrough the skin holes. Remove the splicemembers and stiffeners, and burr all the rivetholes. Apply one coat of zinc chromateprimer to all overlapping surfaces of the splicemembers and temporarily secure the members tothe spar. Rivet the splice members to thespar, using AN442-AD4 rivets through the weband replacing the dri Ned-out skin rivets withAN426-AD3 rivets (see Figure 48). To preventthe splice members from creeping, rivet firstat the corners and then at the intermediatepos i t ions.38. VERTICAL STABILIZER REAR SPAR AREA ABOVEROOT RIB.Damage to the rear spar of the vertical stabilizeron areas located above the root ribmay be repaired as shown (see Figure 49J. Ifthe damage is extensive, cut out the damagedmaterial, locating the cut between two skinrivets at each side of the damage. For thesplice members, prepare two sheets of .040 inchthick 24ST alclad having a length of 6 inchesplus the length of the damage and having a maximumwidth of 6 inches. Along the length ofthe spl ice sheets, bend a flange equal to theRESTRICTED[ 130]

RESTRICTED FIXED SURFACESFIGURE il8032 X 1-3/8, 24STALCLAD STIFFENERS(2 REQUIRED)TRIM SPLICE SHEETS TOMATCH EXISTING LIGHTENINGHOLES IN DAMAGED AREAF NECESSARY032 IN. THICK 24 ST ALCLADSHEET SPLICE MEMBERS(2 REQUIRED)3/47/8EXISTING TYPELIGHTENING HOLESECTIONAASKIN {REF.)v^SKIN { REF )3/43/4note: ALL RIVETS REQUIREDMUST BE USED AT EACHSIDE OF DAMAGEAN442- AD4 iD4 y'^^RIVETS (' 4 REQ.)^ iSECTIONBBAN442- AD4RIVETS (4 REQ.),2 HORIZONTAL ROWS OFAN442-AD4 RIVETS AT 3/4" ON CENTERSFigure 48— Vertical Stabilizer Front Spar SpliceLl3l]RESTRICTED

I overlappingityizer.IFIXED SURFACESPAR. 38 TO >*lRESTRICTEDItapered spar flange; then bend a 90-degree,29/32- inch flange measured from the first bend.Observe a minimum bend radius of 5/32-inch.Trim the splice members as shown (see Figure4q) . For the spar web stiffening angles, bendup two 1-3/8 inch wide strips of .040-inch24ST ale lad of a length equal to the spar depthat the damaged area. With a No. 40 (.098)drill, drill out the affected skin rivets ateach side of the damage, taking particular careto avoid elongating the rivet holes. Securethe splice sheets and the web stiffeners tothe spar. Drill through the splice sheets inseveral locations and temporarily secure thesheets to the spar with skin fasteners. Centerpunch the required sixty rivet locationsat an average spacing of 3/4- inch on centers.At the ends of the splice sheets, observe aminimum edge distance of 3/8-inch measuredfrom the center-punch marks. With a No. 21(.159) drill, drill the required rivet holes.As the spar depth decreases toward the spartip, omit the rivets on the spar web centerline.With a No. 30 (.1285) drill, drill thesplice members through the existing rivet holesin the skin. Remove the splice members andthe web stiffeners, and burr all the rivetholes. Apply one coat of zinc chromate primerto a surfaces of the spl iceImembers. Secure the splice members and webstiffeners to the spar and temporarily securethe members with skin fasteners. Into thesixty No. 21 (.159) holes, insert and driveAN442-AD5 rivets. Toward the spar tip, therivets on the spar centerl i ne and cap flangesmay be omitted. Replace the skin rivets withBI227-AD4 rivets in the quantity required (seeFigure 4g). Rivet first at the corners andthen at the intermediate positions to preventthe sheets from creeping.39. VERTICAL STABILIZER REAR SPAR AREA BELOWROOT RIB.The complete splice of the vertical stabilizerrear spar for the area located belowthe root rib is illustrated (see Figure ^o).Depending upon the extent of the damage, allor part of the complete splice may be used.Remove the stabilizer from the fuselage. Ifthe inner reinforcing angles on the spar capsare damaged, drill out the fifty-six AN442-AD4rivets securing these angles to the spar caps.Use a No. 40 (.098) drill. Remove these anglesand cut out the damaged material, locatingthe cut between two rivet holes at eachside of the damage. For the splice members,bend up two sheets of .040 inch thick 24STale lad having a length equal to the length ofthe damage plus 7-1/4 inches and having a maximumwidth of 6-1/2 inches. Along the lengthof each of the splice members, bend up a flangeequal to the tapered spar cap flange at thedamaged area; then bend a 29/32 inch, 90-degreeflange measured from the first bend. Observea minimum bend radius of 5/32-inch. Trim thesplice members to the shapes shown (see Figure^o) . If the two inner angles are damaged,replace the angles with new angles, each bentup from .040 x 3-5/16 x 20-1/4 sheets of 24STale lad material. Clamp the reinforcing angles,the splice members, and the spar fillersto the spar. At each side of the damage,center punch the required rivet locations atan average spacing of 13/16-inch on centers.Except as noted, observe a minimum edge distanceof 3/8-inch measured from the centerpunchmarks. With a No. 21 ( . 159) dri 1 I, dri Ithe center-punched rivet locations through thesplice members. With a No. 30 (.1285) drill,drill the existing rivet holes at each sideof the splice members through the inner reinforcingangles. Drill additional rivet holesas necessary. Remove the splice members, filler,and inner reinforcing angle, and burr allthe rivet holes. Apply one coat of zinc chromateprimer to all overlapping surfaces of thesplice members, filler, and inner angles. Securethe members to the spar and insert a numberof skin fasteners in the rivet holes alongthe length of the inner angle and in variouslocations in the splice members. At each sideof the splice member, replace the drilled-outrivets through the inner reinforcing angleswith AN442AD-4 rivets. Through the splicemembers, insert and drive eighty-eight AN442-AD5rivets in the pattern noted (see Figure ^o)To prevent the splice members from creeping,rivet first in the corners and then in the intermediatepositions.1^0. REPAIR OF WOODEN HORIZONTAL STABILIZER.The inaccessibi I of the stabi 1 prohibitsall types of repairs except skin repairs.Refer to Section II for skin repairsto wooden structures. Inasmuch as all membersof the stabilizer are glued together, replacementof parts is impractical. Damagedfittings may be replaced (see Figure ^) . Thefollowing paragraphs on horizontal stabilizerspar repairs are applicable to the aluminumst ructure on ly.^\. HORIZONTAL STABILIZER FRONT SPAR.The front spar of the aluminum horizontalstabilizer is formed of .032 inch thick 24STRESTRICTED[132]

RESTRICTEDFIXED SURFACESFIGURE 49TO MATCH SPARTAPER29/32 Ci_WIDTH DETERMINEDBY SPAR DEPTHAT DAMAGED AREAIN THE AREA 6 INCHESFROM SPAR TIP THESERIVETS MAY BEOMITTED ON BOTH CAPSSKIN (REF.)-SKIN (REF.)AS SPAR DEPTH DECREASESTOWARD SPAR TIP, OMITRIVETS ON SPAR ^.040X1-3/8 ANGLE J\24ST ALCLAD6 PLUSLENGTHOF DAMAGESECTION.040 SPLICESHEETS, 24STALCLAD(2 REG.)A-AFigure 49— Vertical Stabil izer Rear Spar Splice for Area Above Root Rib[ 133]RESTRICTED

FIXED SURFACESFIGURE 50RESTRICTEDI»Hsr.040 SPLICE SHEETS 24 ST ALCLAD (2 REO.)SECTIONBB040 X 3-5/8 X 20-1/424ST ALCLAD SHEETREPLACEMENT (2 REQ.)AN442-ADSRIVETS (88 REQ.)7-1/4 PLUSLENGTH OF DAMAGERIVETS MARKED THUS (^ARE IN LOCATIONS OFORIGINAL RIVETSAREPLACEEXISTINGRIVETSA.040 X 5-1/2 X (3 PLUS LENGTH OF DAMAGE)24ST ALCLAD FILLERSECTIONAAFigure SO^Vertical Stabilizer Rear Spar Splicefor Area Below Root RibRESTRICTEDL 134]

I.RESTHICTEDFIXED SURFACESPAR. m TO 42\— SPAR -i=;:::.032 X |-3^ I-3/8X SRAR DEPTH 24ST ALCLAD ANGLE (2 REQ.)SECTION A A-c*-LJ'-~n_HJIf-7^ PLUS LENGTH OF DAMAGEAN426-AD3RIVETS(4 REQ.)SKIN(REF.)AN442-AD4rivets(5req;2 VERTICALROWS OFAN442-AD4RIVETS AT3/4 INCH ONCENTERSAN442-AD4rivets(5req:5j| MAX.2iMIN.AN426-AD3RIVETS(4 REQSECTION BBEXISTING TYPELIGHTENING HOLE032 SPLICE SHEETS24ST ALCLAD (2 REQ.)TRIM SPLICE SHEETS TOMATCH ANY EXISTINGLIGHTENING HOLES INDAMAGED AREA, IFNECESSARYNOTE:ALL RIVETS REQUIREDMUST BE USED AT EACHSIDE OF DAMAGEFigure Sl—Horizontal Stabilizer Front Spar Splicealclad sheet. The splice of the member is illustrated(see Figure ^i) . If the damage isextensive, cut out the damaged material, locatingthe cut between two skin rivets at eachside of the damage. For the splice members,prepare two sheets of .032 inch thick 24ST alcladhaving a length of 7-1/2 inches plus thelength of the damage and having a maximum widthof 3 inches. Along the length of the splicesheets, bend a 13/16-inch, 97-degree flange.Trim the splice members to the proper shape.If the damage occurs in areas where flangedcutouts are located, trim the splice sheetsto match the holes in the area if necessary.For the spar web stiffener angles, bend up two1-3/8 inch wide strips of .032-inch 24ST alcladof a length equal to the spar depth. Drillout the affected skin rivets at each side ofthe damage. Clamp the spl ice members to thespar. Drill through in several locations andtemporari \y secure the spl ice members to thespar with skin fasteners. At each side of thedamage, center punch the required rivet locationsat an average spacing of 3/4- inch oncenters. With a No. 30 ( . 1285) dri II, dri Ithe center-punched rivet locations. At theedges of the sheets, observe a minimum rivetedge distance of 3/8- inch measured from thecenterline of the rivet hole to the edge ofthe material involved. Remove the splice membersand burr all the rivet holes. Apply onecoat of zinc ch romate primer to all overlappingsurfaces of the spl ice members and theweb stiffeners. Again secure the members tothe spar, and rivet the splice members withAN442-AD4 rivets in the quantity required. Replacethe skin rivets with AN426-AD3 rivetsin the quantity required (see Figure ^i)42. HORIZONTAL STABILIZER REAR SPAR AREA OUT-BOARD OF CENTER HINGE FIHING.The rear spar of the aluminum horizontal stabilizerin the area outboard of the centerhinge fitting may be spliced as shown (see Figure^2). Depending upon the extent of the damage,all or part of the complete spl ice maybe used. Carefully cut out the damaged material,locating the cut between two skin rivetsat each side of the daniage. For the splicemembers, cut two sheets of .051 inch thick24ST alclad, each having a length of 7-1/2inches plus the length of the damage, and amaximum width of 5 inches. Along the lengthof each of the splice members, bend up a 90-degree flange equal to the spar flange at theL I3b] RESmiCTED

6-I outadoftizerFIXED SURFACESPAR. 42 TO 44RESTRICTEDdamaged area. Measure 29/ 32- inch from the bendand then bend another 90-degree flange. Observea minimum bend radius of 3/ 1 inch. Inareas having flanged cutouts, trim the splicesheets around the holes as required. For thespar web stiffeners, bend up two sheets of.040 inch thick 24ST ale lad having a width of1-3/8 inches and having a length equal to thespar depth at the damaged area. Dri I theaffected skin rivets at each side of the damage.Secure the spl ice members to the spar.Drill through in several locations and temporarilysecure the members to the spar withskin fasteners. At each side of the damage,center punch the required rivet locations atan average spacing of 3/4-inch on centers.Observe a minimum edge distance of 3/8-inchmeasured from the center-punch mark to the edgeof the material involved. With a No. 30 (. 1285)drill, drill the center-punched rivet locations.Remove the splice members and applyone coat of zinc ch romate primer to all overlappingsurfaces of the members. Burr the rivetholes. Secure the members to the spar withskin fasteners. Rivet the splice members tothe spar with AN442-AD5 rivets in the quantityrequired (see Figure ^2)- Rivet firstat the corners and then at the intermediatepos i i ons.43. HORIZONTAL STABILIZER REAR SPAR AREA IN-BOARD OF CENTER HINGE FITTING.If serious damage occurs to the spar areainboard of the center hinge fitting of thealuminum horizontal stabi I rear spar, itis advisable to replace the entire spar.44. SKIN - GENERAL.Except for the 52S-I/2H aluminum alloy tipskin on the aluminum horizontal stabilizer,al I the skin covering the fixed surfacesconsists of 24ST aluminum alloy. The wing skinvaries from .020 to .064 inch thick 24ST alcI (see Figure 5

lestI theWhereinchsRESTRICTED FIXED SURFACESPAR. »m TO 46Figure 53— V/ing S^in ArrangementLOWER SURFACErepairs are perfectly acceptable. In areaswhere interior access is limited, use Cherryblind rivets in place of regular rivets, subjectto the restrictions set forth in the followingparagraphs.45. SMALL HOLES IN SKIN LESS THAN 1/2- INCHDIAMETER.Where the hole is located away from interiorstructure and is less than 3/16-inch ind iameter, fill the hole with the shank of arivet (see Figure ^^, Detail A). the hole,is located away from interior structure andis up to 3/8- inch in diameter, install a smallbacking sheet of 24ST alclad and fill the holewith the head of an AN442-A0 rivet (see Figure^^, Detail B) . Where the hole is located directlyover a stringer, dri t hole to the,size of the sma I rivet head of a flat headrivet up to and including a 3/ 16- inch diameterrivet, and insert and drive an AN442-AD rivetthrough the structure (see Figure 35, DetailC). If the rivet spacing is more than 8 timesthe diameter of the rivet, add a rivet on eachside of the damage, midway between the holeand the adjacent rivets. This repair is especiallyuseful where cracks appear in dimpledskin. It i to be noted that these repairsshould not be used within 8 inches of spar orexisting cutouts or reinforcement, or within15 inches of a bolting angle.46. SMALL HOLES IN SKIN 1/2- TO 1-INCH DI-AMETER.With a hole saw, cut out all small holesin the skin to a round shape (see Figure 56J . .If the largest dimension of the hole does notexceed Iin diameter, the damage usuallymay be adequately repaired by means of thesmall circular button patch illustrated (seeFigure 57>'.; Cut the patch from 24ST alcladof the same thickness as the damaged skin.Spot-weld or rivet the filler to the patch.Apply the patch to the skin and drill throughthe skin and patch in the rivet pattern noted.Use a No. 40 (.U98), a No. 30 (.1285), and aNo. 21 (.159) drill for 3/32-, I/8-, and 5/32-inch diameter rivets, respectively. Use the same[137]RESTRICTED

eotherinedFIXED SURFACESPAR. 16 TO 48RESTRICTEDtype rivet (countersunk or brazier head) asis used in the area surrounding the damage.If the hole is located near the intersectionof a rib and a stringer, extend the patch topick up the surrounding rivets in the area(see Figure 58, Details A and B) . If the holeis directly over a stringer, the damage may beadequately repaired as shown (see Figure 58,Detail C) . For the repair of holes near adjacentstructure, the rivet data table is entirelyapplicable (see Figure ^l)• It is tobe noted that all the patches shown are ofthe flush type. Whenever the patch is to beslipped Detween the skin and structure, chamferedges of the backing sheet to provide a moregradual change in contour. Wherever the smoothsurface is not required, the backing sheet maybe applied to the outside surface of the skin,and the fi Her may be omitted. These reinforcementsshould not be used within 8 inches ofa spar, existing cutout, or reinforcement, orwithin 15 inches of a bolting angle.1^7. ISOLATED SKIN CRACKS.When an isolated crack is found in the skin,drill a No. 40 ( . 098 ) ho I at the ends of thecrack to arrest further growth (see Figure 59J.Cut a patch of 24ST alclad large enough tooverlap the damage by 1-1/4 inches all aroundthe crack. Fasten the patch in place, drillthrough the skin and patch, add rivet with twostaggered rows of rivets as out I for patchinglarge holes in skin.18. PREPARING LARGE HOLESING.IN SKIN FOR PATCH-Large holes in the skin may be repaired bypatching if the damage affects less than onethirdthe width of the skin panel and if thedamage is not more than 8 inches wide. Holeswhich exceed these restrictions should requirethe replacement of all or part of the sxinpanel. Extensive skin damage occurring nearor directly over existing doublers should requireskin panel replacement. Flush patches(see Figure 60) should be used on the flushriveted leading edge skin, because externalpatches introduce unfavorable stalling characteristics.However, for al I locations,the external repair (see Figure 62) is recommendedfor field repairs, because it requiresless time and skill to apply than the flushpatch. The procedure for both the flush andexternal skin repair is as follows: With a-FUSELAGEALL SKIN IS24 ST ALCLADEXCEPT ASNOTED.DECIMAL FRACTIONSINDICATE SKINTHICKNESS ININCHES.032 /^ 52 S '/? HHORIZONTAL STABILIZER VERTICAL STABILIZERFigure 54—Stabilizer Skin ArrangementRESTRICTEDL 138]

eamerr.IRESTRICTED FIXFD SURFACESPAR. ^Q TO 49ABFILL HOLEWITH RIVETSHANKHOLES UP TO ^y,e DIAMETER24ST ALCLADSHEETFILL HOLE WITHHEAD OF AN442-ADRIVETRIVET BACKINGSHEET TO SKINHOLES UP TO %' DIAMETERHOLEFILLWITH HEADOF AN442-ADRIVETHOLES UP TO %' DIAMETERDIRECTLY OVER STRINGERFigure 55— Skin Repair for HolesLess Than 1/2' inch Diameterout the existing rivets to be covered by thepatch (see Figure 6^). Apply the patch tothe inside or outside of the skin as required,center punch and drill through the corners ofthe patch, and temporarily fasten the patchin place with skin fasteners (see Figure 63JIf interior access can be gained, drill anoutside patch through the existing rivet holesto be covered by the patch. Use the dri I I fromthe interior. For an outside patch where it isnot possible to manipulate the dri 1 1 in the interiorof the structure, remove the patch and apply asheet of paper over the area. Fasten the paperby means of masking tape. To locate the rivetholes beneath the paper, blacken the area overeach of the holes with a soft penci I, thus out I iningthe holes on the paper. On the paper, blackenalso the locations of the four holes previouslydrilled to temporarily secure the skinpatch (see Figure 64}. Next, remove the paperand place the paper on the patch, lining upthe four corner holes on the paper and the patch.Center punch the rivet locations on the patchthrough the paper. Remove the paper, and dri Ithe patch with the size drill required for thethickness of the skin as outlined below. Securethe patch to the skin with skin fastenersand mark a line around the patch 3/R-inch infrom the patch edge; then mark a rivet linearound the patch 3/8-inch in from the edge ofthe cutaway damage. It is to be noted thatthe procedure may have to be varied where theI igrease pencil, mark a rectangle around thedamaged area with the sides parallel to thelongitudinal stringers. With a pair of dividers,mark each of the four corners witha 1-1/2 inch radius. Cut along the markedne wi th a spi ra I or a pa i of met a Isnips, taking part icu lar care to cut smoothlyat the radii of all the corners (see Figure61). Remove the damaged material and smoothout the rough edges of the cut with a file.Obtain a sheet of 24ST ale lad of the same thicknessas the affected skin panel. Cut the patchto overlap the hole a minimum of 1-1/4 inchesall around; but if the hole is located nearstringers, the overlap may have to be increasedin order to pick up the stringer rivets. Provideall rivet locations around the edges ofthe patch with an edge distance of 3/8-inch.Chamfer the edges of the patch.1^9. PATCHING LARGE HOLES IN SKIN.After the damage is trimmed and the 24STa Ic lad patch is cut to the proper s ize, drill Figure 56 — Using Hole Saw To Trim Skin DamageL 139]RESTRICTED

FIXED SURFACESPAR. 1^9 TO 50RESTRICTEDFOR .040 ANDTHINNER SKINFOR .051 ANDTHICKER SKINVIEW OF INSTALLED PATCH24 ST ALCLADBACKING SHEETOF SAMETHICKNESSAS SKIN..FILLER(RIVET ORSPOTWELDTO BACKINGSHEET.)^OUTER SKINSURFACE(REF.)tHATl..

I equalI throughI outRESTRICTEDFIXED SURFACESPAR. 50 TO 513/4-inch wide and 6 inches long, riveted togetherat one end with several AN442-AD4 rivets.At the other end of the strips, a No. 40 (.0Q8)pilot hole is line-drilled through ooth sheets.In one of the strips, dri Ithe pi lothole with a dri Ito the size of the bl indholes to be located. It is to be noted thatIt is desirable to make a separate tool foreach of the four common rivet diameters encountered;namely, 3/32, 1/8, 5/32, and 3/16.Into the hole in one of the strips of the samesize as the blind holes, insert an AN442-ADrivet of the proper diameter. Stake the shankof the rivet to the sheet. In application,slip the patch sheet between the two strips ofthe tool, mnipulate the tool so that the rivetshank in the tool drops into the rivet holeDRILL HOLE AT ,,ENDS OF CRACK'^-'^l^^— RIB (REF.)SKIN (REF.)FILLERFigure 59~Arresting Growthof Crack24ST ALCLADSHEETA CORNER REPAIRSKIN(REF)FILLERSTRINGER(REE)covered by the patch sheet (see Figure 6y).When this occurs, mark rivet location on thepatch sheet through the guide hole in the topsheet of the tool. Certain types of blind rivethole locating tools employ a pilot bushinginstead of a pilot hole (see Figure 68).51. SPLICING SKIN PANELS.24ST ALCLADSHEETREPAIR ADJACENT TO STRINGER, FRAMEOR RIBSKINFILLER(rEF.)24ST ALCLADSHEETC REPAIR DIRECTLY OVER STRINGERFigure 58— Patch for 1-inch DiameterSkin Holes Near Adjacent StructureHoles in the skin that exceed one-third thewidth of the skin or exceed 8 inches in diametershould require the complete replacement of thedamaged panel. However, if the damage is Ichcalizedand complete replacaiient is not warranted,it is sometimes permissible to cut the entiredamaged panel and to splice a replacement skinpanel into position. It is to be noted thatno spl ice should be made in the same bay withthe end of the sheet. Never should two splicesbe made in the same bay, and it is not permissibleto make a spanwise splice within 15inches of the spars. Either a spanwise or achordwise splice may be employed; but sincepractical ly all the skin panels are considerablyshorter in a chordwise direction, a chordwisesplice is recorrmended for general repair,for it entails less work (see Figure 6q) . Properlysupport the affected section of the structure,and driI al I the rivets securingthe damaged portion of the skin panel to thesurrounding structure (see Figure 6^). Takeparticular care to prevent the elongation ofL 141]RESTRICTED

theFIXED SURFACESPARAGRAPH 51RESTRICTED-DOUBLE THE NUMBER OF RIVETS IN THE STRINGERTHROUGH THE PATCH, IF EXISTING RIVETS ARESPACED GREATER THAN |-|/4" O.C.SAME TYPE RIVETS AS USED IN SURROUNDINGAREA. SIZE OF RIVETS AND SPACING TO BEDETERMINED BY THICKNESS OF SKIN AS OUTLINEDIN TEXT..040 inch thick skin. Drill through the oppositeends of the splice, and fasten thesplice overlap in place with skin fasteners.Along the three marked rivet lines of. theskin overlap, center punch the required rivetlocations as noted below for the thicknessof the skin involved. Drill the overlaprivet locations with the size drill notedbelow. Remove the replacement skin panel andburr al I rivet holes. Apply one coat ofzinc chromate primer to all overlapping surfaces.Again secure the skin replacementpanel to the structure. Rivet the replacementskin portion to the surrounding structurewith the same type and size rivets previouslyremoved from the damaged skin. Rivet thesplice overlap with the rivet size, rivetspacing, and rivet rows noted below (see Figure6g) . Do not stagger the rivets in the rows.THICKNESS RIVETOF SKIN TYPE ANDIN INCHES SIZERIVETPATTERN(ON CENTERS)SPACEBETWEENROWSDRILL,020,025032040,051*.064*AD3 3 ROWS AT 5/8-1 N.AD4 3 ROWS AT 5/8- IN.AD4 3 ROWS AT 5/8- iN.AD5 3 ROWS AT 3/^-1 N.1/2- IN. NO. 401/2-1 N. NO. 301/2-IN. NO. 305/8-IN. NO. 21Figure 60— Typical Flush Patchfor Large Skin HolesI ithe existing rivet holes. Pull back the damagedportion of the skin panel; and with apair of snips, cut the entire damaged skin panelin a chordwise direction. Cut a sheet of24ST ale lad of the same thickness and dimensionsas the damaged portion of the skin panel,but extend the dimension of the replacementsheet a minimum of 2 inches beyond the cutedge of the damaged skin panel on .040 inchthick skin. On .032-inch skin and thinner,the splice overlap may be decreased to I-3/4 inches. Place the new skin section beneaththe damaged skin section that was removed.Drill the rivet holes in the new skinsection, using the existing rivet holes in thedamaged panel as a guide. Secure the new skinpanel to the structure and fasten with skinfasteners. Where the replacement skin panelsection overlaps the cut edges of the originalskin panel, mark a ne in from the patchedge 3/8- inch and then mark the two remainingrivet lines 1/2- inch apart on .032 inchthick skin and thinner or 5/8-inch apart onBecause of the shape and location of thesepanels on the wing cent ersec t i on , completereplacement is recommended.Figure 61 — Using Spiral Reamer To Trim Skin DamageRESTRICTED[ 142]

RESTRICTEDFIXED SURFACESPARAGRAPH 52/CHAMFER/ PATCH/ EDGESrA.ROUND OUT DAMAGE AND/CAREFULLY RADIUSCORNERS (1-1/2 R.)SECTIONA AFigure 62— Typical External Patchfor Large Skin Holes52. PREPARING WING LEADING EDGE SKIN FOR SPLIC-ING.Figure 64—Locating Blind Rivet Holesby Improvised Methodlocating the cut within 2 inches of the nearestundamaged nose rib. With a No. 40 (.09^)drill, drill out all the rivets securing thedamaged skin to the structure. Take part icularcareto avoid elongating any of the existingrivet holes in the skin. Lift the damagedportion of the skin free from the surroundingstructure. If any of the nose ribsare damaged, see the applicable paragraph forIf the skin forward of the leading edge stringeris severely damaged, the damage should berepaired by splicing in an entire leading edgeskin replacement sheet as outlined herein (seeFigure yo). At each side of the damage, cutinto the leading edge skin with a hack saw.fDRILL THROUGH PATCH AND SKIN^SECURE PATCH WITH SKIN FASTENERS.Figure 63— Temporarily Securing Skin PatchPrior to RivetingFigure 65—Drilling Out RivetsCovered by Skin PatchL 143]RESTRICTED

I a/!^-FIXED SURFACESPARAGRAPH 52No. 40 PILOT HOLEPILOT BUSHINGFOR No40 DRILLRON PLATE FOR BUSHINGNSTALLATIONRESTRICTEDSTAKEDBOTTOM VIEWDOWN RIVET SHANKFigure 66— Blind Rivet Hole Locating ToolI( .098)rib repairs. For the splice members^ cut twosheets of .040 inch thick 24ST alclad, eachhaving a maximum length of 27 inches and awidth of 3 inches. Along the length of thesplice sheets, bend up 1/4-inch flanges, observinga minimum bend radius of inch. Inthe center six inches of the splice sheets,taper cut the flanges and dri I No. 40relief hole at the termination of the flanges.Form the splice sheets to the contour of theleading edge as sh6wn (see Figure no). At eachBOTTOM VIEWSTAKED DOWN RIVET SHANKFigure 68-^Blind Rivet Hole Locating ToolWith Pilot Bushingof the cuts in the leading edge, clamp thesplice sheets, allowing half of the width ofthe sheets to extend beyond the cut. On eachof the splice locations, mark a line 3/8- inchin from the edge of the cut and then mark anotherline around the leading edge contour7/8-inch in from the cut. Along each of theselines, center punch rivet locations at an averagespacing of 3/4- inch on centers and staggerthe rivet locations in the rows. With aNo. 30 (.1285) drill, drill the center-punchedrivet locations through the skin and the splicesheets. Countersink the rivet holes 100 degreesby 7/32- inch. Into these holes, insertand drive AN426-AD4 rivets in the quantity required.Thus far in the repair, the damagednose skin is cut away and the splice sheetsare riveted to the stubs of the leading edge.REPLACE EXISTINGRIVETS THROUGHSKIN.SECTIONAAI227-A03 ON 020BI227-A04 ON 032BI227-AD5 ON 040Figure 67—Using Blind Rivet Hole Locating Tool Figure 69 — Typical Skin SpliceRESTRICTEDL 144]

ef,IRESTRICTED FIXED SURFACESPARAGRAPH 5353. SPLICING WING LEADING EDGE SKIN.IiBy following the procedure in the precedingparagraph, the damaged leading edge skinis prepared for splicing. Next, for the noseskin replacement panel, cut a sheet of .040inch thick 24ST alclad having a maximum widthof 27 inches and a length equal to the lengthof the damage. Roll the replacement skin panelto the contour of the leading edge, andtrim the sheet to fit the leading edge. Ifseveral bays of the leading edge are crushedin, check the contour of the leading edge replacementskin with the correct leading edgecontours shown (see Figure li) . Note the stationlocation of the damage (see Figure lo)and then check the exact contour at the station(see Figure 72 J. Lay out the noted contourin full size upon a panel of 1/2- inch plywoodand cut around the contour with a bandsaw. Use this ful l-size template to check thecontour of the leading edge skin replacement.Position the replacement skin panel in the correctlocation on the upper side of the leadingedge and drill upward through the stringer and285) drill. the areasheet wi th a No. 30 ( . Iis otherwi se naccessi b I for the driI I i ng,the lower edge of the replacement panel mayI thebe pulled away from the structure to provideaccess. Insert skin fasteners through the upperskin and stringer. Along the lower contourof each end of the skin replacement panel,mark a line 3/8- inch in from the edge ofthe sheet and then mark a second line alongthe contour 7/8- inch in from each edge of thereplacement panel (see Figure no). Along eachof the lines, lightly center punch rivet locationsat an average spacing of 3/4- inch oncenters, and stagger the rivet locations in eachrow. With a No. 30 ( . 1285) drill, dri Icenter-punched rivet locations, and insert skinfasteners. With a No. 30 ( . 1285) dri I I, dri Idownward through the lower stringer and skin,using the existing rivet holes in the stringeras a guide. If it is necessary to gain accessto the lower stringer for drilling, removethe skin fasteners securing the upper edgeof the skin to the upper stringer and pull backthe upper edge of the sheet. Along the uppercontour of each end of the skin replacement panel,mark a line 3/8- inch in from the edge ofthe sheet and then mark a second line alongthe contour 7/8- inch In from each edge of thereplacement panel. Center punch the rivet locationsand dri M No. 30 ( . 1285) holes in eachrow at an average spacing of 3/4- inch onRIB (REF.) RE-DRILL EXISTING SKIN RIVET LOCATIONS THROUGHREPLACEMENT SKIN. REPLACE AN426-AD4 RIVETS3/4'RIB(REF.)LEADING EDGESTRINGER (REF.)TAPER FLANGES AS REQUIREDAND DRILL NO. 40 (.098) RELIEFHOLES AT END OF TAPERSECTIONA AFigure 70 — y/ing24ST ALCLAD SKINREPLACEMENT ,040X 27" X LENGTH OFDAMAGE1/4'Leading Edge Skin SpliceDRILL NO. 30 (.1285)CSK. 100° X 7/32" DIA.AN426-AD4 RIVETSAS REQUIRED FORNOTED PATTERNAN426-AD4 RIVETSSTAGGERED (REF)24ST ALCLAD SPLICE SHEET.040 X 3"X 27' (2 REO )L 145]RESTRICTED

FIXED SURFACESFIGURE 71RESTRICTED20.12525.156-30.178MAIN SPAR (^25 L-70. 406-75. 445-85. 250-90. 139105. 625•110.656-115. 688-125.750-135. 80260. 35765. 37595. 313 -0-5.03110.063-15.089-35.21940 . 25045.26880.344-100.496-120.713130.781

RESTRICTED FIXED SURFACESPAR. 53 TO 54i33aBaaa-x,„^MAIN SPARLOWER CAPFigure 72—Gaining Access to theOuter Wing Leading EdgecenterB; stagger the rivet locations in each row.Remove the skin panel replacement, burr allthe rivet holes, and apply one coat of zincchromate primer to all overlapping surfaces.Again secure the skin replacement panel to theleading edge with skin fasteners. Rivet theskin panel to the lower leading edge stringerand to the lower areas of the two spl icesheets with AN426-AD4 rivets in the quantityrequired. If it is necessary to gain access,pull the top edge of the replacement panelaway from the structure. Next, rivet the up-per area of one of the spl ice sheets and theupper stringer to the skin with AN426-AD4 rivets.Remove the skin fasteners as the dril I-1ng progresses. If it is necessary to gainaccess to buck the rivets, pull the corner ofthe skin replacement away from the structure.Lastly, rivet the upper area of the skin panelto the rerraining splice sheet with AN426-AD4rivets. If the area is inaccessible for rivetbucking. Cherry blind rivets may be substitutedfor AN426-AD4 rivets for riveting theupper area of the remaining splice sheet and

I required,IFIXED SURFACESPAR. m TO 55RESTRICTEDbe avoided if possible by partial splicing orpatching. However, if over one-half the skinpanel is damaged, removal and replacement ofthe complete skin panel may have to be accomplished. If this is the case, careful ly dri I Iout all attaching rivets after properly supportingthe structure. Use a drill one sizesmal ler than the diameter of the rivet, andplace the point of the drill into the depressionin the rivet head. If a power tool isused, locate the drill on the rivet first, thenapply the power in short bursts. Drill untilthe rivet head twists from the shank, and thendrift the remainder of the rivet free. Particularcare must be exercised to prevent theelongation of the existing rivet holes. Removethe damaged skin panel and cut a replacementpanel, using the damaged skin as a template ifpossible. Place the panel beneath the originaldamaged panel previously removed. Drill throughthe two panels and temporari ly fasten with skinfasteners. Drill most of the skin and stringerrivet locations in the replacement panel, usingthe existing rivet locations in the originaldamaged panel as a guide. Install access coversand make other modifications on the replacementskin panel to rmtch the original skin panel.Figure 75— Welding a Rod to a Dent in theHorizontal Stabilizer Tipback the skin as required. If access is desiredto the interior structure forward of the outerwing main spar, the skin panel forward of theclosing strip of skin must be removed in certainareas (see Figure 72 J. When re riveting thestrip, prevent the sagging and bulging of theadjacent sheets by inserting a narrow driftpunch in the rivet hole just ahead of the rivetbeing driven. Whi le driving the rivet, tightenthe adjacent sheet by pressing the top of thepunch forvvard or aft, as required (see Figure j^).Where the fabrication of a new portion of theclosing strip becomes necessary, the generalprocedure outlined for the removal and replacementof the skin panels is entirely applicableexcept for the following: Using the damagedsection of skin as a template, center punchthe hole locations and drill the rivet locationswith the size dri I ti Iting the dri Iforward fO degrees at the rear edge and tiltingthe drill aft 10 degrees at the front edgeof the strip, measured from the normal perpendicularto the skin surface (see Figure 74J.2S0 ALUMINUMWELDING ROD3/16" DIAMETERRESTORE TIPCONTOUR WITHFIRM PULL,THEN CUT RODFROM TIP55. REMOVING AND REPLACING SKIN PANELS.Replacement of an entire panel of skin shouldFigure 76— Removing a Dent in theHorizontal Stabilizer TipRESTRICTEDL 148]

nI aRESTRICTED FIXED SURFACESPAR. 55 TO 57Burr all the rivet holes and apply one coatof zinc chromate primer to all overlapping surfacesof the replacement skin panel. Placethe skin panel in the proper position on thestructure and fasten with a sufficient numberof skin fasteners to prevent the panel fromcreeping while riveting. Using the same typeand size rivets as were used in the originalpanel, rivet the replacement panel, first atthe corners and then at the intermediate positi ons.a 2-foot length of 3/ 6- inch1 2S0 aluminum weldingrod. Bend up the rod at the center to forma doubled length. Place the ends of the rodat the center of the dent and weld the rod tothe dent with a welding torch (see Figure l^)- .Hydrogen welding facilities are desirable, butoxy acetylene equipment may be used with a tipsize slightly larger than would ordinarily beused with steel of the same thickness. Adjustthe torch to produce a slightly carbonizingflame, soft and bluish and not "blowy". UseFRONTSPAR.040-INCHTHICK 24 STALCLAD162 X2'/4 X32V8PLATE 24 STALCLAD

I 3/6-outad,orFIXED SURFACESPAR. 57 TO 58RESTRICTED.064- - INCH THICK , 24 STALUM. ALLOY WEBINBOARDVIEWOUTER WINGBOLTING SURFACEOUTBOARDVIEWREF. DWG. 77 - 13016R. H. SHOWNFigure 79 ^Centersect ionEnd Ribsplice member 1-5/8 inches wide and 4-1/2 incheslong plus the length of the damage. Along thelength of the 4-1/2 inch wide sheet, bend a 90-degree, 5/ 16- inch flange; then bend a 90-degree,inch flange1measured from the first bend.Along the length of the remaining splice sheet,bend a 90-degree, 5/8-inch flange. Trim thesplice members around any stringer cutouts inthe affected area. Dril I the affected skinrivets at each side of the damage. Clamp thesplice sheets to the former, and center punchthe required rivet locations at each side ofthe darage. With a No. 3C) (.1285) drill, drillthe splice member through the existing rivetholes in the skin, and drill the remaining rivetlocations as required. Remove the spl ice membersand burr the rivet holes. Apply one coatof zinc chromate primer to all overlapping surfaces.Reel amp the spl ice members to the former;and through the skin, insert and drive BI227-AD4 rivets in the quantity required. Throughthe remaining holes, insert and drive twelveAN442-AD4 rivets (see Figure 77 J.58. CENTERSECTION INTERMEDIATE RIBS.The centersect ion intermediate ribs, locatedon the upper surface of the centersect ion betweenthe front spar and the rear spar, may bespliced (see Figure 80). If the damage is extensive,cut out the damaged material with ahack saw, locating the cut between two skinrivets at each side of the damage. Dependingupon the extent of the damage, al I part ofthe complete spl ice may be used. For example,if only sheet material of the rib is damaged,disregard the extruded member splice and usethe spl ice of the sheet material only. For acomplete splice, cut a length of Type K78C extrusionequal to 6 inches plus the length ofthe damage. Also cut two sheets of .051 inchthick 24ST ale I each having a width of 2-3/4inches and a length of 6 inches plus the lengthof the damage. Along the length of one of thesheet splice members, bend a 90-degree, 5/8-inch flange, observing a minimum bend radiusof 5/32-inch. Along the length of the otherRESTRICTEDL 150]

.I.RESTRICTED FIXED SURFACESPAR. 58 TO 59sheet splice member, bend a 90-deyree, 9/16-inch flange. Chamfer the outside comer of theextruded K78C splice member to fit the insideradius of the existing extruded flange on therib. With a No. 40 (.098) dril I, dri II out theaffected skin rivets at each side of the damage.Clamp the spl ice members to the proper positionon the rib. With a No. 30 ( . 1285) dri I I, dri Ithrough the extruded splice member, using theexisting rivet holes in the rib as a guide. Centerpunch and drill new rivet locations through thesheet spl ice members as required to obtain thenoted pattern. Remove the spl ice members and burrall the rivet holes. Apply one coat of zincchromate primer to all overlapping surfacesof the sheet spl ice members and to al I surfacesof the extruded member. Again clamp thespl ice members to the rib. Through the skinrivet holes, insert and drive BI227-AD4 rivetsin the quantity required. Through the remainingtwenty-six holes in the rib, insert anddrive AN442-AD4 rivets (see Figure 8o)59. WING TRAILING EDGE RIBS AND CENTERSECTIONLEADING EDGE RIBS.TTie wing trailing edge ribs, inboard of theTRIM SPLICE SHEETSAROUND CUTOUTSAN442-AD4 RIVETS(26 REQ.)aileron cutout, and the centersect ion leadingedge ribs may be spl iced as shoM^ (see H^ure 8i)[Replacement of these ribs may be very readilyaccomplished; but if spare parts are not available,it is entirely acceptable to splice therib. If the damage is in the form of a crack,drill a No. 40 (.098) hole at the ends of thecrack. If the damage is extensive, cut outthe damaged material between skin rivets. Forthe spl ice members, bend up two sheets of .040inch thick 24ST alclad having a typical widthof 3 inches and a length of 2-1/2 inches plusthe length of the damage. Along the lengthof one of the sheets, bend up a 90-degree, 1/2-inch flange. Along the length of the othersplice rrenber, bend a 90-degree, 5/8-inch flange.Trim the splice sheets as required. With aright-angled drill, drill out the affected skinrivets at each side of the damage. Clamp thesplice sheets to the rib and center punch twostaggered rivet rows at each side of the damage.On the lower rib flange, center punchtwo rivet locations at each side of the damage.With a No. 30 (.1285) drill, drill the centerpunchedrivet locations and drill the splicemembers through the skin rivet holes on theREPLACE EXISTING RIVETS THROUGH SKINUPPER SURFACESKIN (REF.)-7 2-1/2 PLUS LENGTHOF DAMAGE05IX2-3/4"x(6"PLUS DAMAGELENGTH) 24STALCLAD SHEET(2 REQ.)-EXISTING HOLEIN WEB (REF)LAN442-AD4RIVETS (14 REQ.)SECTION A AK77C SPLICEMEMBER (I REQ.).040 X 3" X (2-1/2"PLUS LENGTH OF]DAMAGE) 24STALCLAD SHEET(2 REQ.)4" MAX.^fkSECTION A AIFigure 80 —Centersect ion Intermediate Rib SpliceFigure Si — Trailing Edge Rib Splice[ 151 JRESTRICTED

FIXED SURFACESPAR. 59 TO 61RESTRICTEDinsert and drive AN442-AD4 rivets in the quantityrequired to obtain two vertical rivet rowsat each side of the damage (see Figure 8i).60. DAMAGED RIB BEADS.Figure82-^Vxew of Outer Wing SectionShowing Ribs Ins ta 1 ledupper cap of the rib. Remove the spl ice nierrvbersand burr all the rivet holes. Apply onecoat of zinc chromate primer to all overlappingsurfaces. Again clamp the splice members tothe rib; and through the skin rivet holes, insertand drive BI227-AD4 rivets in the quantityrequired. Through the remaining rivet holes,The repair of damage located on and aboutbeads in the rib web is shown (see Figure 84).If the damage is in the form of a crack, dri IIa No. 40 (.098) hole at the ends of the crack.For the reinforcement, bend up a section of.040 inch thick 24ST aid ad having a typicaldimension of 2-5/8 x 1-1/2 inches. Along thevertical edges of the reinforcement, bend up90-degree, 1/4-inch flanges. Observe a minimumbend radius of 1/8- inch. Apply the reinforcementto the rib side opposite the bead.Center punch the four rivet locations and drillthrough the reinforcement with a No. 30 (.1285)dri I I. Remove the reinforcement and burr al Ithe rivet holes. Apply one coat of zinc chromateprimer to all overlapping surfaces andagain secure the reinforcement to the rib. Intothe rivet holes, insert and drive four AN442-AD4 rivets (see Figure 84).61. DAMAGED RIB CUTOUTS.The repair of damage located on and aboutSTIFFENINGBEADSi^lLliJiiTZTjU-V•^^ -^w 'vr.:»m'm7yj

I the'ine.RESTRICTEDFIXED SURFACESPAR. 61 TO 63TYPICALHOLELIGHTENINGTYPICAL BEADAN 4 42- A D4 RIVETSinch thick 24ST alclad having a typical dirrensionof 4 X I-I/4 inches. Depending upon theextent of the damage, one or several reinforcingangles may be used in the affected area. Cutoff the legs of the angles to a gradual taper,as shown. Apply the reinforcement to the ribweb and securely clamp in place to restore theshape of the rib web. Center punch four rivetlocations and drill through, using a No. 30(.1285) drill. Remove the reinforcement, andburr a I rivet holes. Apply one coat ofzinc chromate primer to all overlapping surfaces,and again clamp the reinforcement tothe rib web. Into the rivet holes, insert anddrive AN442-AD4 rivets (see Figure 86).63. DAMAGED RIB FLANGES.-.040 X 2-5/8 X 1-1/224ST ALCLAD SHEET(I REQ.)l-DRILL NO. 40 HOLE ATENDS OF CRACKSECTION XXwith regard to rib flanges, there are twotypes of ribs employed in the construction ofthe fixed surfaces; one type having the ribflanges riveted to the skin and cut around thestringers, and the other type having the unbrokenflange located approximately 3/4-inchwithin the skin mold I The repa i r of damagedrib flanges riveted to the skin is shown(see Figure 8j) . The repair of damaged ribFigure 84— Typical Repairof Broken Rib Beads040 X 3-3/4 "x 1-3/4"24 ST AUCLADSHEETlightening cutouts in the rib webs is shown(see Figure 85). If the damage is in the formof a crack, drill a No. 4-0 (.098) hole at theend of the crack. For the reinforcement, bendup a section of .040 inch thick 24ST alcladhaving a typical dimension of 3-3/4 x f-3/4inches. Along the length of one of the edges,bend up a 90-degree, 1/4— inch flange. Observea minimum bend ra"dius of 1/8- inch. Appl.y thereinforcement to the rib side opposite the cutoutflange. Center punch the nine rivet locationsand drill through the reinforcementwith a No. 30 (.1285) drill. Remove the reinforcementand burr all the rivet holes. Applyone coat of zinc chromate primer to all overlappingsurfaces and secure the reinforcementto the rib. Into the rivet holes, insert anddrive nine AN442-AD4 rivets (see Figure 85).62. BUCKLED RIB WEBS.Buckled rib webs that are not cracked or distortedbeyond restoration may be straightenedand repaired by the application of an angleperpendicular to the rib flanges (see Figure 86),For the reinforcement, bend up a sheet of .040AN442-AD4RIVETS (9 REQ.)1/4- f u-J u-i '^SECTIONA AFigure 85— Typical Repair of Broken Rib CutoutsL 153]RESTRICTED

I aFIXED SURFACESPAR. 63 TO 6^FESTRICTEDflanges not riveted to the skin is also shown(see Figure 88). These repairs may be usedon bent-up ribs up to and including .064 inchthick ale lad. The procedure is as follows:Dri I No. 40 ( .098) hole at the end of thebreak in the rib flange. Where the rib flangeis riveted to the skin, prepare a reinforcingangle of 24ST ale lad sheet of the same thicknessas the rib, and cut the angle of sufficientlength to extend between adjacent stringers.Where the rib flange is not riveted to the skin,cut a sheet of 24ST ale lad of the same thick-CUT ENDS AN442-AD4 .040X1-1/4X4 TYPICALOF AN6LE-7 RIVETS (4REQ.)7 e4ST ALCLAD LIGHTENINGANGLE -7 HOLE64. SPLICING TYPICAL RIBS.The centersect ion ribs between the rear sparand the flap spar, and all the outer wing noseribs and intermediate ribs may be spliced asoutlined herein. Cut out the damaged rib materialwith a hack saw, locating the cut atthe center of a lightening hole (see Figure8g) . For the splice members, cut two sheetsof .040 inch thick 24ST ale lad having a widthof 2-1/4 inches and a length of 4-1/2 inchesplus the length of the damage. Along one ofthe edges of the length of .each of the splicesheets, bend a 90-degree, 1/4- inch flange andthen bend a 90-degree, 5/8-ineh flange measuredfrom the first bend. Along the rema ining edgeof the length of each of the splice sheets,bend a 90-degree, 3/16-inch flange and then benda 90-degree, l/2-ineh flange measured from thefirst bend. Observe a minimum bend radius of1/8- inch. Clamp the splice members to the upperand the lower caps of the rib, and centerpunch the twenty-eight required rivet locations.With a No. 30 (. 1285) dri I I, dri I I the centerpunchedrivet locations. Remove the splicemembers and burr all the rivet holes. ApplyFigure 86~-Typical Repair of Buckled Rib Websness as the rib and of a length sufficient topick up four rivets at each side of the break.Bend up the sheet to fit the rib flange on theoutside. Drill out the affected rivets at eachside of the damage and fit the reinforcing angleto the rib cap. Where the rib flange is rivetedto the skin, drill the angle through, usingthe same size dri II as originally usedthrough the skin. Where the rib flange is notriveted to the skin, use a No. 40 (.098), a No.30 {.'1285), and a No. 21 (.159) dri II to dri IIthe rivet holes through .020-, .025-, and .051-inch rib material, respectively (see Figure 88).Remove the angle and burr all the rivet holes.Apply one coat of zinc chromate primer to alloverlapping surfaces and resecure the angle tothe rib flange. Where the rib flange is rivetedto the skin, replace the skin rivets; andthrough the rib web and angle, insert and driveAN442-AD rivets of the same size as the existingskin rivets. Where the rib flange is notriveted to the skin, use AN442-AD3 rivets through.020 inch thick ale lad, AN442-AD4 rivets through.025 inch thick material, and AN442-AD5 rivetsthrough .051 inch thick alcl ad. Use this repaironly on bent-up ribs where there are nodoublers or other reinforcements within 8 inchesof the dzimaged flange.REPLACE EXISTINGBI227-AD RIVETSTHROUGH SKIN«7SECTIONHA-A-TYPICALLIGHTENINGHOLE (REF) //)RILL NO. 40HOLE AT ENDOF CRACKRIBr- SKIN (REF)WEB (REF.)AN442-ADRIVETS. SAMESIZE ASEXISTINGSKIN RIVETSFigure 87— Typical Rib Flange RepairWhere Flange Is Riveted to SkinRESTRICTED[ 154 J

.RESTRICTED FIXED SURFACESPAR. 6H TO 66one coat of zinc chromate primer to a M overlappingsurfaces. Again clamp the splice membersto the rib, and insert and drive twentyeightAN442-AD4 rivets into the rivet holes.EXISTING LIGHTENING HOLE (REF.)^-]-^N442AD4 RIVETS (28 REQ.)--kl/X. TYPICAL BEAD (REF)O65. REMOVING AND REPLACING DAMAGED RIBS.Local investigation will determine the procedureto follow in gaining access to a damagedrib. With a No. 40 (.098) drill, drillout the rivets securing the rib flange to theskin and the stringer clips. Replacement ofthe ribs should be accomplished by utilizingNorth American Aviation, Inc., spare parts (seeFigures i and 2). Where spare parts are notavailable and immediate repair is necessary,use the damaged rib as a template and fabricatea replacement locally. Lay out the shape ofthe affected rib upon a panel of 1/2- inch wood.In several places, temporarily nail this panelto a second panel of 1/2- inch wood; and witha keyhole saw, cut around the marked contour.(See Figure go). Pull out the nails and separatethe two equivalent panels cut to thecontour of the affected rib. Measure thedeveloped width of the rib flange and cut anequivalent sheet of 24ST ale lad equal to the© /to eAN442-AD3 RIVETS(.020)AN442-AD4 RIVETS{.025AN442-AD5 RIVETS(.05I)(8 REQUIRED)24ST ALCLAD ANGLETHICKNESS AS RIB.DRILL N0.40 HOLEAT END OF BREAKSAMESECTION AAFigure 88— Typical Rib Flange Repair WhereFlange Is Not Riveted to Skin-.040 X Z^y. (4|^^'PLUS LENGTHOF DAMAGE) 24ST ALCLADSHEET (2 REQ.)Figure 89-Typical Rib Splice16 ZSECTION A-Acontour of the wood panels plus the developedwidth of the rib flange all around. Clamp thesheet of a'lclad between the 1wo wood form blocksin a vise; and with a mallet, pound the flangesof the replacement rib over the edge of one ofthe wood form blocks (see Figure gi ) . Usingthe damaged rib as a guide, bend up the sametype of flanges as were originally employed.In place of the flanged cutouts and stiffeningbeads employed on the original rib, bendup 1-1/4 inch wide strips of .040 inch thick24ST ale lad to a right angle, and secure themon the rib web perpendicular to the rib flanges.Taper cut the upstanding leg of the stiffeneras shown (see Figure 86). Install the replacementrib in exactly the same manner as theor ig na i I66. WING JOINT BOLTING ANGLES.No repairs to the wing joint bolting anglesare permissible. Replacement of damaged boltingangles must be accomplished. The centersectionbolting angle consists of four sections(see Figure 92J . The outer wing bolting anglealso consists of four sections (see Figureg^). Disassemble the wing at the joint,and drill out the existing rivets through theL 155]RESTRICTED

ineI otI ofIFIXED SURFACESPARAGRAPH 66RESTRICTEDCENTERSECTIONFRONT SPARHRH)-66-13104 (-IRH)Figure 90^'Cutt ing Replacement Rib Form Blocksbolting anqle section and the skin, using extremecare to prevent the elongation of anyof the existing rivet holes. Use a dri I I onesize smaller than the rivet diameter. Applythe drill to the depression in the rivet headand drill in short bursts until the rivet headtwists from the shank. Punch out the remainingportion of the rivet. Use this procedureto remove all the rivets from the bolting angles.Remove the screws from the remainingholes and take off the damaged angle. Thefuel tank cover bolting angle and the lowerrear bolting angle of the centersect ion areFigure 92— Centersect ion Bolting Angle Fart Nunberssuppi ied as spare parts with the pi holesdrilled in the proper locations. The top boltingangle of the centersect ion and a I thebolt i ng ang les of the outer wi ng are suppI iedas spans parts without pilot holes. In eithercase, Iup the new bolting angle flush withthe skin edge, and securely clamp the anglein position with heavy "C" clamps. On the newbolting angles without drilled pilot holes,drill through the bolting angle with a No. 40(.098) pilot drill, using the existing rivetholes in the skin as a guide. Then dri Ithe pilot holes through with a drill equal-55-14027-266-14027-2 (RH)CLAMPALCLAD SHEETBETWEEN TWOWOOD FORMBLOCKS54-14030 (-IRH)MAIN SPAR (REF.)/(AT-6C){AT-6B)(AT-6A)88-14029 (HRH)84-14029 (-IRH)55-14029 (-IRH)55-14028 (-IRH)Figure 93—Outer Wing Bolting Angle Part NuntersFigure 91 — Forming Replacement Ribto the size of the existing holes in the skin.As the dri I I ing progresses, slip a thin hookshapedhack saw blade, with the teeth removed,between the bolting angle and the skin andscrape out the drill chips. Using an undartBgedwing for reference, countersink and spot-faceRESTRICTEDL1^6 J

.RESTRICTED FIXED SURFACESPAR. 66 TO 67HOOK COVER ONTRAILING EDGEBRACKETFigure 94— Installing Lower Portionof Wing Joint Coverthe various holes. Again using an undamagedwing as a guide, insert and drive the requiredsize and type rivets and install the requiredsc rews67. WING JOINT COVERS.Serious l.y damaged wing joint covers mustbe replaced rather than repaired because thealignment of the covers is critical. A splicewould affect the al ignment of the cover andprevent proper installation. For replace-Figure 96^-Locating Bracket Holein Wing Joint Covermake a grease pencil mark at the centerlineof the screw hole in the wing cover attachingbracket (see Figure g^). Cut off the headof a BI286-8-7 screw, file the end down to apoint, and thread the screw into the wing coverattaching bracket hole, leaving the sharp pointprotruding about 1/8- inch. This serves as aback center punch to locate the hole throughthe cover. Slip the cover assembly over theleading edge of the wing joint. Grasp the laverportion of the cover, and hook it over thesmall bracket at the lower trailing edge ofthe wing bolting angles (see Figure c^). ContinueBI286-8-7SCREW(2 REQ)Figure 95— Installing Upper Portionof Wing Joint Coverriients, ^lorth Anerican Aviation, Inc., spareparts must be used. Obtain the NA 54-F0020lower cover and the NA 77-10023 upper cover,and assemble them by means of two B\ 28^-8-1screws at the leading edge portion (see FigureQ7J. On the lower outboard surface skin.Figure 97 — Assembling Wing Joint CoverL1&7]RESTRICTED

.FIXED SURFACESPAR. 67 TO 70RESTRICTEDhold ing the lower t rai li ng edge of the coverto keep the cover hooked; then grasp the uppertrailing edge of the cover and lift up thehandle. Pull aft on che upper cover; andwhen tne cover is in position, push the handledown over the hook to complete the installation(see Figure 95J . Along the lower surfaceskin where the centerline of the screwhole in the cover attaching bracket is marked,measure 7/8- inch down on the contour of thecover from the edge of the cover, and tap thecover lightly with a mallet (see Figure g6)The previously installed sharp screw protrudingfrom the bracket screw hole will centerpunch the hole location on the inside of thecover. Remove the cover; and in the cover,drill the center-punched rivet location witha 3/8-inch d iameter dri I I . Again install thecover as outlined above. Through the brackethole that was just drilled in the cover, instalI a BI286-8-7 screw through a B985-I 1-20-32washer. This completes the installation.68. EXTRUDED ALUMINUM REPAIR MATERIAL.The various extruded cross-sectional shapesthat are required in the repair of the structuralmembers of the fixed surfaces are listedbelow. The material used in forming theseextrusions is 24ST aluminum alloy, conformingto Federal Specification QQ-A-354. Numberswith a C prefix are North American Aviation,Inc., standard parts. Following the NorthAmerican Aviation, Inc., part number, theAluminum Company of America (Alcoa) equivalentdie number is listed in parentheses.C148TC180TC203TC204TC245TC250TEXTRUSION(K73U-JJ)(K14280)(K14653)(K14654)(K16862)(K16869)DIE NUMBERSC265TC266TC274TC366TK77AK77C69. ALCLAD SHEET REPAIR MATERIAL.(L2'3557)(L23558)(L23966)(L24910)K77a)(K77C)The aluminum sheet material used in therepair of the wing should be 24ST ale lad conformingto Federal Specification QQ-A-362.The various thicknesses of sheet that may berequired are listed below. The lengths andwidths of the sheets must be determined locallyfor the extent of repairs to be undertaken.with reference to the repair procedures out-I i ned in this Sect ion.THICKNESS OF 2ilSTALCUD (inches).020.025.032.040.051.064REMARKSRIBS, SKIN (C107LT-20)RIBS. SKIN (C373LT)RIBS, SKIN, SPARSRIBS, SKIN, SPARS (c123Lt1RIBS, SKIN, SPARSRIBS, SKIN, SPARS70. RIVETS, BOLTS, SCREWS, AND NUTS REQUIREDFOR REPAIR.The following types of rivets, bolts, scnews,and nuts may be required for the repair ofthe fixed surfaces. Types of rivets, bolts,screws, and nuts are described in Section I.PART NO.B1227-AD3B1227-AD4B1227-AD5B1227-AD6AN426-AD3AN426-AD4AN426-AD5AN426-AD6AN442-AU3AN442-AD4AN442-AD5AN442-AD6LS- 1127-4-2LS-1 127-4-4LS-1 127-4-6LS-1 12 7-4-8LS-1127-5-2LS-1127-5-4LS- 11 2 7-5-6LS-1127-5-8LS-1127-6-4LS-1127-6-6LS-1 127-6-8LS-1126-4-4BI25I-IO32-I4BI25I-IO32-I7B1251-428-14B1248-1032-17B1083AC365-103222GCH-04822G 1-048RIVET, BRAZIER HEADRIVET, BRAZIER HEADRIVET, BRAZIERDESCRIPTIONHEADRIVET, BRAZIER HEADRIVET, 100° C'SUNKRIVET, 100° C'SUNKRIVET, 100° C'SUNKRIVET, 100° C'SUNKRIVET, FLAT HEADRIVET, FLATRIVET, FLATHEADHEADRIVET, FLAT HEADRIVET, CHERRY BLINDRIVET, CHERRY BLINDRIVET, CHERRY BLINDRIVET, CHERRY BLINDRIVET, CHERRY BLINDRIVET, CHERRY BLINDRIVET, CHERRY BLINDRIVET, CHERRY BLINDRIVET, CHERRY BLINDRIVET, CHERRYRIVET, CHERRYRIVET, CHERRYSCREW, NO. 10SCREW, NO. 10SCREW,SCREW, NO. 10NUT, ELASTICNUT, NO. 10CHANNEL, BOOTSNUTS, BOOTSBLINDBLINDBLINDSTOPNUTBRAZIERBRAZIERBRAZIERBRAZIERBRAZIERBRAZIERBRAZIERBRAZIERBRAZIERBRAZIERBRAZIER100°C'SUNK3/32 DIA.1/8 DIA.5/32 DIA.3/16 DIA3/32 Dl A.1/8 DIA.5/32 DIA.3/16 DIA.3/32 DIA.1/8 DIA.5/32 DIA.3/16 DIA.1/8 DIA.1/8 DIA.1/8 DIA.1/8 DIA.5/32 DIA.5/32 DIA.5/32 DIA.5/32 DIA.3/16 DIA.3/16 DIA.3/16 DIA.1/8 DIA.1/4 DIA.1/4 DIA.RESTRICTEDL 158]

FORRESTRICTED FIXED SURFACESPARAGRAPH 7171. REPAIR TOOLS.The following list is a brief summation ofthe tools that are required in repairing thefixed surfaces:BARS,TOOLBUCKINGRIVETSEE SECTION IREMARKSTYPES OF BARS.TOOLFASTENERS, SKINFILES, HANDREMARKSUSED TO SECURE METAL WHEN RIVET-ING, DRILLING, ETC.USED FOR FINISHING WORK ON ALLTYPES OF METAL, PLASTIC, AND FIBER,AND ENLARGING HOLES, BURRING EDGES,SMOOTHING CURVES, REMOVING METAL,AND SHARPENING TOOLS.BRAKE, HANDBURNISHING TOOLBURRING TOOLCALIPERS, SLIDEUSED FOR BENDING ALCLAD SHEET.USED TO SMOOTH MINOR SCRATCHES INMETAL.USED TO SMOOTH TRIMMED EDGES OFMETAL.INSIDE OR OUTSIDE DIAMETER MEAS-UREMENT.FILEHOLDERGUN, CHERRYRIVETGUN, CLECOSAFETYGUN,PNEUMATICUSED TO HOLD FILES IN CURVED POSI-TION.MODELS GIO OR G15 (SEE SECTIOI IX).USED TO INSERT SKIN FASTENERS.USED TO DRIVE STANDARD RIVETS.CHISELCLAMP,"C"USED FOR METAL CUTTING AND REMOVALOF RIVET HEADS WHERE DRILLING ISIMPRACTICAL.USED FOR HOLDING TWO OR MORE PIECESOF METAL UNTIL DRILLING AND/OR RIV-ETING IS ACCOMPLISHED.HACK SAW,KEYHOLEHACK SAW, LARGE,ADJUSTABLEUSED FOR CUTTING HOLES IN RESTRICT-EDPLACES.USED FOR METAL, FIBER, ANDPLASTIC CUTTING. BLADES VARYIN LENGTH AND IN NUMBER OF TEETHPERINCH.CLAMP,SKINUSED TO SECURE METAL WHEN RIVETING,DRILLING, ETC.HACK SAW, SMALLUSED FOR LIGHT CUTTING OF METALFIBER, OR PLASTIC.COUNTERSINKCOUNTERSINK,BACKCOUNTERSINKS HOLES IN HEAVYSTOCK FOR INSERTION OF COUNTERSUNKRIVETS AND SCREWS.USED FOR COUNTERSINKING IN LOCA-TIONS WHERE MOTIVE POWER CANNOTBE APPLI ED ON SI DE DES IRED.HAMMER,INGMALLET,FINISH-LEADMALLET, PARALYN,RAWHIDEUSED TO FORM AND FINISH METAL.USED FOR SHCET METAL FORMING.USED TO FORM AND BEND METAL.DIVIDERSUSED FOR SHEET METAL LAYOUT.MALLET,RUBBERUSED FOR SHEET METAL LAYOUT.DRILL, HANDDRILLS, TWISTUSED FOR RESTRICTED OR SLOW DRILL-ING, AND ALSO IN THE PRESENCE OFGAS FUMES, IF AN AIR MOTOR IS NOTAVAILABLE.THESE DRILLS USED WITH HAND ORPOWER DRILLS ARE STANDARD.PILOT, HOLE SAWPLIERS, CLAMPINGPROTRACTOR, FLATUSED TO PILOT HOLE SAWS.USED FOR HOLDING TWO OR MOREPIECES OF METAL UNTIL DRILLINGAND/OR RIVETING IS ACCOMPLISHED.USED FOR SHEET METAL LAYOUT.EXTRACTOR,EZY-OUTUSED FOR REMOVING THE REMAININGPORTIONS OF BROKEN SCREWS, BOLTS,AND PINS. DRILL HOLE IN OBJECTAPPROXIMATELY ONE-HALF THE DIAM-ETER, INSERT EZY-OUT, AND WITH-DRAW OBJECT BY COUNTERCLOCKWISEROTATION.PUNCHES, CENTERPUNCH, DRIVEPIN "DRIFTS'USED FOR CENTERING HOLES AND AS APUNCH BEFORE DRILLING TO ENSURE ANACCURATE DRILL START.USED FOR DRIVING OUT PINS, RIVETS,AND BOLTS. USE NEXT SIZE SMALLERTHANHOLE.[ 159]RESTRICTED

FIXED SURFACESPARAGRAPH 71RESTRICTEDTOOL REMARKS TOOL REMARKSREAMER, HAND, USED FOR ENLARGING DRILLED HOLES SKIN FASTENER,SQUARE SHANK, TO DESIRED SIZE. VARIOUS SIZES CLECOSPIRAL FLUTED AVAILABLE.INSERTED IN DRILLED HOLES TO HOLDTWO OR MORE PIECES OF METAL UNTILRIVETING IS ACCOMPLISHED.REAMER, SQUARESHANK, TAPERPIN, SPIRALFLUTEDRULE, 6-FOOT,FLEXIBLERULE,RIGID6-INCH,RULE, SQUARE,CENTER AND PRO-TRACTOR COMBI-NATIONSAW,SAW,BANDHOLEUSED FOR ENLARGING DRILLED HOLESIN SHEET STOCK AND FOR REAMINGSHAFT HOLES FOR TAPER PINS. INELECTRIC AND AIR MOTORS, REAMERSARE EXCELLENT FOR MAKING CUTOUTSIN ALUMINUM MATERIAL.USED FOR SHEET METAL LAYOUT.USED FOR SHEET METAL LAYOUT,USED FOR SHEET METAL LAYOUT,USED FOR CUTTING FUSELAGE FRAMEFORM BLOCKS.USED WITH PILOT IN HAND AND POWERDRILLS, THIS TOOL IS EXCELLENT FORCUTTING CIRCULAR HOLES IN DAMAGEDSHEET METAL PRIOR TO APPLICATIONOF STANDARD BUTTON PATCH. SIZESRANGE FROM 5/8 TO 2 INCHES.SNIPS, COMBINA-TION CIRCLE"DUCK BILL"SNIPS, LEFT-HAND, DOUBLE-ACTION"DUTCHMANS"SNIPS, TINNERSSPOT-FACER,BACKTAPVISE, DRILLPRESS, SPEEDWHEELS, EMERYUSED FOR STRAIGHT LINE AND CURVEDCUTTING.USED FOR CUTTING TO LEFT AND/ORCUTTING- ENDS OF TUBING.USED FOR STRAIGHT LINE CUTTING OFSHEET METAL UP TO .064 INCH THICK.USED WITH VARIOUS SIZES OF PILOTSWHERE ACCURACY OF HOLE LOCATION ISDESIRED. ALSO USED FOR FLUSHINGBOLT HEADS IN CASTINGS.USED FOR THREADING VARIOUS SIZESOF HOLES FOR ALL STANDARD THREADS.AVAILABLE IN DIFFERENT DIAMETERS,THREADS PER INCH, AND PITCH.USED TO HOLD OBJECTS WHILE DRILL-ING.SMOOTH AND ENLARGEOF LARGE HOLES.NSIDEEDGESSCRIBESETS,RIVETUSED FOR SHEET METAL LAYOUT.USED FOR BRAZIER, FLAT, AND COUNT-ERSUNK RIVET HEADS.WRENCH, TAP USED IN CONJUNCTION WITH TAPS,EZY-OUTS, HAND AND TAPERED REAMERS.RESTRICTEDL 160]

ingRESTRICTED MOVABLE SURFACESPAR. I TO3SECTIONUMOVABLE SURFACES1. GENERAL.The movable surfaces consist of hydraul ical \yoperated, all-metal landing flaps and cablecontrolled,fabric-covered, metal frame ailerons,elevators, and rudders.2. LANDING FLAP CONSTRUCTION.The centersection and outer wing landing flapsextending from left-hand aileron to right-handaileron pass beneath the fuselage and form thelower half of the trailing edge of the wing(see Figure i). The three flap sections areconstructed entirely from 24ST alclad sheet andeach section consists of an .040 inch thickhat-section channel spar, an .032 inch thickZ-section leading edge, an .025 inch thickclosed V-section trailing edge, and .020 inchthick ribs and skin. Where the hat-section sparis riveted to the flap skin, a torsion box isformed which reacts the main bending loads. Themaximum loads on the flaps occur when the flapsare in the ful \y extended position. The structuralpart numbers of centersection and outerwing flap are illustrated (see Figure 2)' Eachof the outer wing flaps has a length of 89-3/8 inches, a constant chord of 14-11/32 inches,and a maximum movement of 45 degrees. The centersectionflap has a length of 56-5/8 inches,a constant chord of 14-11/32 inches, and a maximummovement of 45 degrees. The flaps are notstatically balanced.3. RUDDER CONSTRUCTION.The 24ST alclad aluminum alloy rudder frameis assembled about a 1-3/4 inch diameter 24STaluminum al loy torque tube which forms the primaryload-bearing member of the structure. Thecomponent part numbers of the rudder structureare illustrated and their numbers given (seeFigure q) The ribs are formed from subsequentlyheat-treated .020 inch thick 24S0 alcladsheet. In original rib forming, powerpressing is utilized to provide the integralcapstrips, stiffening beads, and flanged cutouts.The ribs extend from the leading edgeto the trailing edge and are riveted to thetorque tube by means of 24ST aluminum alloyflanged collars. The leading edge forward ofthe torque tube is formed of several sheet sectionsof .020- inch 24ST alclad. A 3-pound castlead counterweight attached at the top of therudder provides adequate static balance. Thetrai I edge consists of a Type CI44LT rol ledsection which is omitted for a short lengthto allow for the installation of a controllabletrim tab. The curved top of the rudderis formed of two shal low half sections of 52S-I/2H material welded together about the commoncircumference and riveted to the uppermostrib of the structure. A cap, pressed from asingle sheet of .032 inch thick 3S0 aluminumalloy, is affixed to the lower extremity ofthe rudder.CENTERSECTIONFLAPFigure 1—Landing Flaps Installed on WingA control horn, forged from I4ST aluminum alloy,is attached to the torque tube at the lowerhinge point. Sealed-type ball-bearing hingepoints are provided by means of Type BI226 eyeboltsinto which are pressed NA 200-K4 ball bearings.The bearings are secured in the eyeboltsby annu lar crimping. An all-metal controllabletrim tab, a trim tab operating mechanism andcable, three hinge fittings, and a control hornare assembled on the frame. The rudder frameis covered with doped fabric as outlined in afollowing paragraph. The complete assembly isattached to and supported by the rear spar ofthe vertical stabilizer (see Figure 4). Therudder has an area of 13.21 square feet including2.26 square feet of balance and .60 squarefeet of tab area. The rudder has an angularmovement of 35 degrees to the right and 35 degreesto the left.[I6i]RESTRICTED

MOVABLE SURFACESFIGURE 2RESTRICTEDS) ® © ®NORTH AMERICA// AVIATION PART NUMBERS1'66'18002 FLAP - CENTER SECTION *55-18001 FLAP - OUTER WING^L-H.55-18001-1, . .FLAP - OUTER WING^R.H.TEM PART NO1. 78-1800112. 78-180093. 78-18009^. 78-1800115. 66-180076. 55-180227. 66-180188. 55-180229. 66-18008TITLE ITEM PART NO. TITLERIB - R.H. ,-, r78-l8004-3 RIB - L.H.'"•RIB - R.H.178-18004-2 RIB - R.H.RIB - L.H. M. 55-18009-1 RIB - R.H.RIB - L.H. 12. 55-I8001I-I....RIB - R.H.EDGE - TRAILING 13. 55-18030 EDGE - TRAILINGHINGE - R.H. .^. fbb-\B022 HINGE - L.H,CHANNEL - CENTER SECTION FLAP'^'1^5- 18022- I.. .. H INGE - R.H.HINGE - L.H. 15. 55-18029 CHANNEL - OUTER WING FLAPEDGE - LEADING 16. 55-18031 EDGE - LEADINGFigure l—Landing Flap Part NumbersRESTRICTED[ 162]

RESTRICTED MOVABLE SURFACESPARAGRAPH 177-24002-654-2401455-24017Figure 3— Rudder Frame Part NumbersI*.ELEVATOR CONSTRUCTION.The elevator assembly consists of two interchangeablesections bolted together at the centerto a short torque tube to which are attachedthe control horn and the center hinge points forthe complete a,ssembly. When the two sectionsare bolted together, the entire structure issupported by hinge brackets which are boltedto the rear spar of the horizontal stabilizer(see Figure 6). The general construction andthe fabric covering of each section is substantially the same as that of the rudder. The followingdifferences from the rudder constructionwill be noted: The cast lead counterweightlocated inside the leading edge at theoutboard end of each section weighs 5.25 pounds.The inboard end of each elevator frame sectionL 163]RESTRICTED

ewerbearings,ingikeMOVABLE SURFACESPAR . 4 TO 6RESTRICTEDterminates at a rib and is therefore not providedwith a cap. Type BI226 eyebolts, intowhich are pressed NA 200-K4 bal I comprisetwo of the hinge points for each section-The center hinge points for the complete installationconsist of sealed-type ball bearingspressed into I4ST forgings located on the interconnectingtorque tube (see Figure 6). Thetrim tab actuating cables extend through theinboard rib forward of the main torque tube ofeach elevator section. Each of the elevatorsections has an area of 10.88 square feet including1.75 square feet balance and .73 squarefoot of tab area. The entire elevator assemblyUPPER HINGE BALANCE WEIGHTHOUSINGFigure 4—RudderVERTICAL STABILIZERREAR SPARLOWER HINGECONTROL HORNInstalled on Vertical Stabilizeri 1has an angular movement of 30 degrees up and20 degrees down. All component part numbersof the rudder structure are lustrated andtheir numbers given (see Figure 5J.Note: If either of the elevator sections isinterchanged, care must be taken toprovide proper drainage for the moisture causedby condensation inside the elevator. Draing rommets are located at the trailing edge ineach bay on both sides of the elevator fabriccovering (see Figure 6). When the elevatorsections are reversed, the centers of thesemust be cut out on the I side and the oneson the upper side resealed by the applicationof a doped fabric patch.5. AILERON CONSTRUCTION.The ai lerons are located at the trai ling edgeof each wing outboard of the flaps (see Figure8) and extend fraiiwing station 147.1 to 243.4.The aileron frame structure is made entirelyfrom 24ST ale lad and is covered with dopedfabric. The 24ST ale lad spar and leading edgeskin form a torsional ly rigid box. Balancingcounterweights, secured inside the leadingedge skin, provide adequate static balance.A metal booster tab is located on the trailingedge to provide for ease of control. A totalof three hinges supports each aileron. Theflanged leading edge ribs are riveted to theleading edge skin and to the trai ling edge ribs.The trailing edge ribs are formed from 24STale lad and are provided with integral flangedcapstrips in which dimpled holes are spaced toprovide for the insertion of the countersunkfabric attaching screws and washers. The GradeA mercerized cotton fabric covering is attachedto the ribs and at the trim tab and hinge cutoutsas outlined in a subsequent paragraph.The structural parts of each ai leron are i I-lustrated and their numbers given (see Figurey). Each aileron has an area of 11.40square feet including 2,42 square feet balanceand .44 square foot tab area. Each aileron hasan angular movement of 30 degrees up and 15 degreesdown, utilizing a differential motion of2 to I.6. RUDDER, ELEVATOR, AND AILERON ORIGINALFABRIC COVERING.The fabric covering of these surfaces is madefrom Grade A mercerized cotton fabric, dopedand finished as outlined in subsequent paragraphs.The cover is machine-stitched as faras possible along the trailing edge; then thecover is si ipped over the structure I an envelopeand hand-sewn at the open areas as required.The fabric covering is attached to thedimpled holes in the trai I edge rib capstripsby countersunk sheet metal screws insertedthrough dimpled washers. Reinforcing tape isplaced along the ribs under the washers andscrews and finishing tape is placed over themto provide a smooth finished surface. Two metalriveting strips are used in conjunction withAN426-AD3 countersunk rivets to secure the fabricto the trim tab channel. A row of holesaround the edge of each hinge cutout permitsthe fabric to be hand-stitched to the metalRESTRICTEDLIb4]

VRESWICTEDMOVABLESURFACESFIGURE 555-22023'23-2200466-22002-266-22002-666-2200555-2203450-2200655-2201555-22038—66-2202554-2200366-22013REF. DWG. NO.66-2200250-22014Figure 5— Elevator Frame PartNumbers[ 165]RESTRICTED

inedI overlappingMOVABLE SURFACESPAR. 6 TO 9mTRIMTAB -7TRIM TAB OPERATINGMECHANISMDRAINGROMMETSRESTRICTEDOUTBOARDHINGEBALANCE WEIGHTHOUSINGiAHORIZONTALSTABILIZERREAR SPARFigure 6— Elevators Installed on Complete Horizontal Stabilizerleading edge skin at these points. Finishingtape and fabric and canvas patches are appliedover all seams or rough areas to provide asmooth finished surface.7. CONTROL SURFACE TRIM AND BOOSTER TABS.All trim and booster tabs of the control surfaceson the earl ier AT-6C Series and previousairplanes are of a riveted ale lad construction,while those on the later AT-6C Airplanes areconstructed of wood and phenolic fiber and areassembled with casein glue. Since neither typeof these trim and booster tabs is balanced, itis very important that the tab hinges and controlshave as little play as possible to preventvibration. Severely damaged trim or boostertabs of wood or metal construction should berep laced.8. FLAP REPAIR - GENERAL.The flap spar, leading edge, trailing edge,and skin may be repaired as outlined in thefol lowing paragraphs. Replacement of damagedribs is recomnended; but if repair is warranted,fol low the spl ice procedure out I for trai I-ing edge ribs of the wing (see Section iii).The landing flaps are not statically balancedin original construction and no balance checkis necessary after repair. AM parts of theflap are easily accessible for repair.9. FLAP SKIN REPAIR.Damage to the flap skin may be repaired exactly as out I i ned for the repa i r of the sk i nof the fixed surfaces. (See Section iii.) Ifthe skin directly under the flap channel sparis damaged and the damage is less than one inchin diameter, the repair rray be made as illustrated(see Figure g) . With a hole saw, trimout the hole. To serve as a reinforcement, cuta sheet of .032 inch thick, 24ST ale lad havingan approximate width and length of 1-3/4inches. Drill out the affected rivets at eachside of the damage and slip the reinforcementunder the flap channel spar. With a No. 40(.098) drill, drill the reinforcement throughthe existing holes in the skin and add a rivethole between each two existing holes. Thendrill two rivet holes through the reinforcementand the skin that lies between the spar channellegs. Remove the reinforcement, burr therivet holes, and apply one coat of zinc chrormteprimer to a I surfaces. Again insertthe reinforcement beneath the channel sparand drive BI227-AD3 rivets through the channelspar leg flanges and the reinforcement plate.Insert and head an LSII27-4-4 Cherry rivet intoeach of the two blind holes, and trim thestems with a pair of flush-ground cutters- Aspreviously noted, damage to other portions ofthe flap skin may be repaired as outlined inSect ion Ml.RESTRICTEDL 166]

RESTRICTED MOVABLE SURFACESPARAGRAPH 1010. FLAP CHANNEL. SPAR.The channel spar of the centersect ion and outerwin^ flaps may be repaired as illustrated(see Figure lo). The repai r out I i ned hereincovers damage to the flap channel spar only.If the skin beneath the spar is damaged, referto applicable paragraph for skin repairs. Ifthe damage to the spar is extensive, cut outthe damaged material and locate the cut betweentwo skin rivets at each side of the damage. Ateach side of the damage, measure 4-1/2 inchesand drill out the affected skin rivets in thisarea. To serve as spl ice members, prepare twosheets of 24ST alclad having a width of 2-5/4inches and having a length equal to the damagelength plus 9 inches. Bend up the splice membersto a Z shape as shown, observing a minimum54-16057-54-16055-54-1605352-16030-54-1605166-1605655-1604966-1605455-1608866-1605255-1602754-1604555-16005 54-1604366-1605054-1604!66-1604855-16076-255-1603966-1604666-1604466-16042/fEF DWG. 66-16002 66-1604055-16076^ 55-16038Figure 7 — Aileron Frame Part Numbers[167]RESTRICTED

ad'theI theingI theMOVABLE SURFACESPAR. 10 TO 12RESTRICTEDbend radius of 1/8-inch. Prepare the fillersheets as required. Apply the splice merribersand fi ller sheets to the spar, dri I I throughseveral existing skin holes, and temporarilyfasten the members in position. Drill the thirtysixholes through the cap and the splice memberswith a No. 30 ( . 1285) drill. Dri I spl icemembers through the existirig rivet holes in theskin with a No. 40 (.098) drill. Remove thesplice members, and burr all the rivet holes.To burr the rivet holes on the inside of thehat, use a swiss needle-point file and filethrough the hole with a circular motion. Applya coat of zinc ch romate primer to overlappingsurfaces. Hesecure the spl ice members tothe spar with skin fasteners. Replace theBI227-AD3 rivets through the skin. In the blindportion of the hat, insert thirty-six LSI 127-4-4Cherry blind rivets and pul I stems of therivets with a GIO or GI5 rivet gun. With apair of nippers, trim the rivet stems flush withthe rivet heads.I I. FLAP LEADING EDGE.Damage to the leading edge of both the centersectionand the outer wing flap may be repairedas shown (see Figure ii)- If the damageto the leading edge is extensive, cut outthe damaged material between two skin rivetsat each side of the damage. From each side ofthe damage, measure 3-3/4 inches and drill outthe affected skin rivets in that area. To serveas a splice member, cut a sheet of .U32 inchthick ale I having a width of 3-1/4 inches anda length equal to the length of the damage plus7-1/2 inches. Bend the splice members up tothe Z shape shown, observing a minimum bendradius of 1/8- inch. Prepare a filler sheet ofI.052 inch thick alclad having a width of 3/4-inch and a length equal to the length of thedamage. Clamp the splice member and the fillerto the flap leading edge, and center punch therequired rivet locations at each side of thedamage. Drill the twenty rivet holes throughthe spl ice member and the leading edge with aNo. 30 ( . 1285) dri I spl ice member. Dri Ithrough the existing rivet holes in the skinwith a No. 40 (.098) drill. F^emove the splicemembers, and burr all the rivet holes. Applyone coat of zinc ch romate primer to all overlappingsurfaces of the spl ice member and f i I I er.Reel amp the spl ice member to the leading edge.Through the .I285^inch diameter holes, insertand drive twenty AN442-AD4 rivets. Replace theskin rivets previously drilled out with AN426-AD3 rivets in the quantity required.12. FLAP TRAILING EDGE.The trai I edge of the outer wing and centersectionflap may be spliced as illustrated(see Figure 13J . If the damage is in the formof a crack, drill a No. 40 (.098) hole at theends of the crack. To serve as spl ice members,bend up two sheets of .032 inch thick 24ST alcladto the shapes shown on the diagram. Bothof the splice members must be equal in lengthto 6 inches plus the length of the damage; oneof the sheets should be 2 inches wide while theother should be 3 inches wide. Drill out theFLAP SPARCHANNEL032Xl|xi| 24ST ALCLADREINFORCEMENTr-BI227 AD3 RIVETS (6 REQ.)ADD RIVET BETWEENEXISTING RIVETSTOP VIEW, L H. WING (REFBOOSTER TAB —L. H. AILERONFigure 8-^Left Aileron Installed onWingSPOTWELDOR RIVETFILLER TOREINFORCEMENTLSI 127-4-4CHERRY (BLIND) RIVET{2 REQ.)Figure 9 — Repair of One- inch Diame te r Holein Skin Under Flap Channel SparRESTRICTEDL 1B8]

RESTRICTED MOVABLE SURFACESPAR. 12 TO m-4^9 PLUS LENGTH OF DAMAGEu o o5^ T "i^io>Q> Q Q/J o o.040X2^X0 PLUSDAMAGELENGTH) SPLICE SHEETS(2 REQ.)REPLACE EXISTING BI227-AD3 RIVET THROUGH SKINLSII27-4-4 CHERRYBLIND RIVETS-^^(36 REG)Figure 10— Flap Channel Spar Spliceaffected skin rivets at each side of the damagewith a No. 40 (.098) drill. Apply thesplice members to the trailing edge, and drillthe splice members through the existing skinrivet holes. Double the number of the skinrivet holes at the splice location by drillingan additional hole between each of the existingrivet holes. Drill the eight rivet locationsthrough the overlap of the splice memberswith a No. 30 (.1285) drill. Remove thesplice members, and burr all the rivet holes.Apply one coat of zinc chromate primer to a I Ioverlapping surfaces and temporari ly secure thesplice members to the trailing edge with skinfasteners. Replace the skin rivets with BI227-AD3 rivets. Through the rivet holes in theoverlap of the splice members, insert LSI 127-4-2 Cherry blind rivets, and expand the rivetswith a GIO or a GI5 Cherry rivet gun. Trimthe Cherry rivet stems with a pair of nippers.13. STRUCTURAL REPAIR OF RUDDER, ELEVATORS,AND AILERONS - GENERAL.Because of the static unbalance created by anyweight added to these surfaces, replacementrather than repair of damaged structural partsis recommended. However* where the repairmaterial is added near the hinge I i ne of thesurface, the static balance of the surface is notseriously affected. Structural repairs to theailerons vary from those of the elevators andrudder, as each aileron has a channel spar inplace of the torque tube employed on the elevatorsand rudder. Obviously, access for any structuralrepair must be gained by cutting awaythe damaged fabric adjacent to the proposedrepair. For fabric removal and replacement,refer to the appi icable paragraphs of this Sec -tion. After a major repair, the static balanceof these surfaces must be checked as outlinedin a fol loA/i ng paragraph.I«+. ELEVATOR AND RUDDER RIB REMOVAL.It is recommended that severely darmged ribsbe replaced rather than repaired. Removal ofan entire rib as a unit is impractical, whereasthe rib may be cut and removed entirely or inpart, as the damage warrants (see Figure 12).To gain access to the structure, remove the coveringadjacent to the damaged rib as outlinedin the appropriate paragraphs of this Section.If this section of the rib aft of the torquetube is damaged, drill out the several rivetssecuring the end of the rib to the trailing edgeL 169]RESTRICTED

MOVABLE SURFACESPAR. m TO 16RESTRICTED.032 X 3-1/4(7-1/2 PLUS DAM-AGE LENGTH )SPLICE SHEET,24 ST ALCLADI ( REQ'D.)( 3/4 X DAMAGENGTH, FILLER 24 STALCLAD (I REQ'D.)PLACE EXISTING AN 442-THROUGH SKINAN442-AD4 RIVETS (20 REQ'D.)Figure 11'- Flap Leading Edge Splicestrip. If the section of the rib forward ofthe torque tube is damaged, drill out the rivetssecuring the rib to the nose skin. Witha hack saw, cut the rib web at the torque tube.Do not cut into the rib col lar. At the collar,use a chisel to cut the remaining portionof the rib web from the edge of the flange tothe torque tube. Cut careful ly to prevent damageto the col iar and torque tube. With a No.40 (.098) drill, drill out the rivets securingthe 61160 collar to the rib web. Lift the damagedrib section free from the structure. Forthe rib replacement, a North American Aviation,Inc., spare part should be utilized (see Figures'^and 5J. If a spare part is not available,and immediate repair is necessary, fabricatea replacement rib section, using the damagedrib as a template. In order to prevent disturbanceto the mass balance of the structure,it is essential that the weight of the new ribmatch the weight of the original rib..032 inch thick 24ST ale lad having a maximumwidth of 3-1/2 inches and a maximum length of3-1/2 inches. On each of the splice members,bend up an angle to match the existing rib capsand trim the splice members around the torquetube. Lightly center punch the rivet locationsin the pattern shown (see Figure 12). With aNo. 30 (.1285) drill, drill the center- punchedrivet locations and drill the splice sheetsthrough the existing holes in the rib collar-Remove the splice sheets, and burr the rivetholes. Apply one coat of zinc chromate primerto all overlapping surfaces. Secure thespl ice members to the rib, and insert and driveAN430-AD4 rivets through the rib collar. Inthe remaining holes, insert and drive AN442-AD4 rivets in the quantity required (see Figure12). Depending upon the extent of coveringremoved, re-cover the surface either whollyor in part as outlined in the appropriate paragraphsof this Section.16. ELEVATOR AND RUDDER LEADING EDGE SKINREPAIR.After the doped fabric covering on the leadingedge of the surface has been properly removedfrom the damaged area, small holes in the032X3-1/2X3-1/2SPLICE SHEETS, 24STALCLAD (2 REO.)TORQUE TUBE (REF)AN442-AD4 RIVETS (20 REQ.)BII60 COLLAR (REF.)15. ELEVATOR AND RUDDER RIB REPLACEMENT.After the rib replacement section is preparedto match the damaged section, place the replacementsection in position and drill No. 40( .098) holes through the t ra i I ing edge stripor the leading edge skin as required, using theexisting holes in the structure as a guide. Securethe rib at the trailing edge or the leadingedge with BI227-AD3 rivets in the quantityrequired. In order to splice the rib cut atthe torque tube, prepare two splice sheets ofSAW CUTDRIVE NEW AN430-AD4RIVETS THROUGH COLLARTYPICALAND SPLICE SHEETSLIGHTENING HOLEFigure 12—Rtxider and Elevator Rib ReplacementRESTRICTEDLl70]

RESTRICTED MOVABLE SURFACESPAR. 16 TO 18leading edge skin may be repaired as set forthin the applicable paragraphs in Section III.If the damage is extensive, remove the entireaffected leading edge section by drilling outthe rivets securing the skin to the ribs. Rollan equivalent leading edge skin section from.020 inch thick 24ST ale lad sheet. Apply thereplacement skin to the proper position on theleading edge ribs; and with a curved scribe, rmrkthe rivet locations on the inside of the skin.Remove the skin, and drill the skin rivet locationswith a No. 40 (.098) drill. Apply onecoat of zinc chromate primer to all overlappingsurfaces of the replacement skin. Again applythe replacement skin to the proper position onthe leading edge ribs and fasten in place withskin fasteners. Through the rivet holes in theskin replacement section, insert and drive AN426-AD3 or BI227-AD3 rivets as originally usedthrough the skin. Rivet first at the absoluteleading edge and then around the periphery ofthe skin. Remove the skin fasteners as thedrilling progresses. Apply fabric covering tothe exposed structure as outlined in the applicableparagraphs of this Section.17. RUDDER, ELEVATOR, AND AILERON TRAILINGEDGE STRIP REPLACEMENT.As all additional weight in the form of repairsshould be avoided at the trailing edgeof a structure, replacement rather than repairof trailing edge strips is recommended.After the doped fabric covering has been removedas set forth in the corresponding paragraphsof this Section, drill out all rivetsthat secure the tral ling edge strip to the ribsand to any longitudinal stiffening tubes. ANorth American Aviation, Inc., spare part shouldbe used for replacing the damaged strip. (SeeFigures 3, 5, and 7. J However, if none is aval I-able, fabricate a replacement strip of 24S0ale lad. Bend up a V-formed section to matchthe length, thickness, weight, and cross sectionof the removed damaged strip. In order to observethe minimum bend radius allowable, use24S0 ale lad and subsequently heat treat to requiredhardness. (See Section I.) Place the stripin position; and with a blind hole locatingtool (see Section III ) . mark the locations ofthe holes for the rib attaching rivets. Removethe strip, drill and burr the holes, and applya coat of zinc chromate primer to all fayingsurfaces. After replacing the strip inposition. Insert two or more skin fastenersto hold it rigidly in place. Insert and upsetall attaching rivets. With Grade Amercerlzedcotton fabric, re-cover the partiallyLS1127-4-2 CHERRY ._(BLIND) RIVETS (8 REQ.)Figure 13—Flap Trailing Edge SpliceSKIN(REF.)exposed structure as set forth for partial fabricre-covering in a following paragraph.18. AILERON SPAR SPLICE.If the aileron spar Is damaged, it cannot bereplaced satisfactorily, but the spar can berepaired without seriously affecting the massbalance of the aileron as a whole. It is obviousthat the entire fabric covering or a sectionof the covering must be removed in order togain access to the spar for repair. For theremoval of covering, refer to the appropriateparagraph of this Section. After the coveringhas been removed, examine the damage to thespar. If the damage is not too extensive, cutout the damaged material between two skin rivetsat each side of the damage. On both the upperand the lower cap of the spar, measure 3-3/4inches on each side of the damage, and drillout the affected skin rivets securing the capIn this area (see Figure 14). To serve assplice members, two sheets of .032 inch thick24ST ale lad are required, each having a maximumwidth of 3 inches and a maximum lengthof 7-1/2 inches plus the length of the damage.Along the length of one of the splice members,bend up an 80-degree, 3/4-inch flange. Alongthe length of the other splice member, bendup an 87-degree, 3/4- inch flange. Observe aminimum bend radius of 1/8-Inch. Trim thesplice members to fit the spar, and trim a-round the splice members to match any existinglightening holes In the affected area. Clamp[171]RESTRICTED

1 thelyI outMOVABLE SURFACESPAR. 18 TO 19RESTRICTEDthe spl ice members to the spar and center punchthe required twenty rivet locations at each sideof the damage. Center punch other rivet locationsto suit the spar depth. Drill the center-punchedrivet locations with a No. 30 {.1283)dri II . Dri I spl ice members through theexisting skin rivet holes. Remove the splicemembers^ and burr all the rivet holes. Applyone coat of zinc chromate primer to all overlappingsurfaces. Reel amp the splice membersto the spar; and through the skin rivet holes,insert and drive AN426-AD4 rivets in the quantityrequired. (See Figure 14.) Through the remainingholes, insert and drive AN442-AD4 rivets.Depending upon the extent of covering removed,re-cover the surface either who I or inpart as outlined in the appropriate paragraphof this Sect ion.19. AILERON TRAILING EDGE RIB REPLACEMENT.The simplicity of the construction of theaileron trailing edge ribs and the ease of theirattachment to the ai leron spar warrants replacementrather than repair of damaged ribs.Rib replacement maintains the static balance ofthe aileron, as no resultant weight from repairmembers is added to the trailing edge portionof the structure. Most of the ai leron trai lingedge ribs are secured to the aileron trailingedge by two BI227-AD3 rivets, to the aileronspar web and to the extending flange on theleading edge ribs by four AN442-AD3 rivets,and to the spar flange and overlapping leadingedge skin by two AN426-AD3 rivets. Obviously,if the trailing edge rib is damaged enough towarrant replacement, the surrounding fabric willalso be damaged and must be removed and repairedas outlined in the applicable paragraphs of thisSection. Remove all fabric necessary to obtainaccess to the damaged rib area. Dri I theeight rivets attaching the rib to the structure.If a stiffening tube extends through the damagedribs, drill out the attaching rivets onone end of the tube to permit the damaged ribto be lifted free and the new rib to be insertedinto position. When removing any of the threeribs which support the booster tab supportingchannel, drill out the two additional rivetssecuring this channel to the rib. For replacementof the damaged rib, a North American Aviation,Inc., spare part should be utilized. (SeeFigure J.) However, if a spare part is notavailable, fabricate a replacement rib from asheet of 24S0 ale lad of the required thicknessand heat treat to required hardness afterSECTIONAA-SPLICESHEET (REF.)• .fm.,.....nM .rTT tfTM.SPAR (REF.)r-UZPLUS LENGTH OF DAMAGE^rxEXISTING TYPELIGHTENING HOLEAN442-AD4RIVETS ASREQUIRED TOSUIT SPARDEPTHAN442-AD4 RIVETS(2.0 REQ.)TRIM SPLICE SHEETSTO MATCH EXISTINGLIGHTENING HOLES,IF NECESSARY.032 X3X (7-1/2 PLUS DAMAGELENGTH)SPLICE SHEETS, 24ST ALCLADFigure 14—Aileron Spar SpliceSECTIONBBRESTRICTED[172]

I rivetI outingRESTRICTEDMOVABLE SURFACESPAR. 19 TO 20Figure15—Sewing Small Fabric Tearforming. (See Section I.) Observe a minimum bendradius of 1/16- inch and round all corners to a1/4- inch minimum radius. After the replacementrib has been obtained or formed, insert it intoposition; and wherever possible, mark and drillall rivet holes through the existing holes inthe structure. Use a curved scribe or a blindhole locating tool (see Section III) for markingthe hole locations of the two rivet holesI i ng progresses.in the spar web. After marking these hole locati ons, remove the rib and drill these twoholes. Burr a I holes and apply one coatof zinc chromate primer to all faying surfacesof the replacement rib. Again place the ribin position. Insert skin fasteners in the rivetholes to hold the rib rigidly in place;then insert and drive all attaching rivets, removingthe skin fasteners as the dri IBy the use of special bucking bars,manipulated through the lightening holes in thespar, buck the rivets securing the rib to theweb of the ai leron spar. After the rib has beenriveted in position, re-cover with fabric theexposed area of the structure by following thedirections set forth for fabric repairs in theappropriate paragraphs of this Section.20. AILERON NOSE RIB REPLACEMENT.In order to remove and replace a damaged aileronnose rib, it will first be necessary toremove the leading edge fabric and skin fromthe damaged rib area. To gain access to theFigure l6~'Cutting Out Damaged Fabricinterior of the leading edge skin through thelightening holes in the aileron spar, removeenough fabric from the adjacent trai I edgebay to permit proper tool manipulation. Witha hack saw or a tapered reamer, cut away thedamaged skin as set forth in a following paragraph.Dri I the attaching rivets of thenose rib and remove the rib from the structure.A North American Aviation, Inc., sparepart should be used for replacing the damagedrib (see Figure j) . However, if such a sparepart is not available, fabricate a replacementrib from 24S0 ale lad sheet of the required thicknessand heat treat to the required hardnessafter forming (see Section I). Observe a minimumbend radius of 1/16- inch and round all cornersto 1/4- inch minimum radius. After the replacementrib has been obtained or forned, insertit into position; and wherever possible,mark and dri M all rivet holes through the existingholes in the trailing edge rib and noseskin. By the use of special bucking bars (seeSection I), manipulated through the lighteningholes in the spar, buck the rivets securing therib to the leading edge skin. After the ribhas been securely riveted in position, repairthe leading edge skin and replace the removed[173]RESTRICTED

MOVABLE SURFACESFIGURE 17BESTRICTEDSEW UP THE TEARAPPLY DOPE TO SEWED AREASCRAPE OFF OLD DOPEAPPLY NEW DOPE>PRE-DOPED PINKED TAPEDOPE LAP AREAPRESS DOWN PATCHING TAPESiDOPE ENTIRE PATCHFigure 17— Patching V-shaped Fabric TearRESTRICTED[174]

ivetRESTRICTED MOVABLE SURFACESPAR. 20 TO 23fabric by following the instructions set forthin the applicable paragraphs of this Section.21. AILERON LEADING EDGE SKIN REPAIR.After the doped fabric covering on the leadingedge of the surface has been properly removedfrom the damaged area, snail holes in the leadingedge skin may be repaired as set forth inthe appropriate paragraphs in Section II f. However,if the damage is extensive, the leadingedge skin of the aileron rrRy be repaired by removingthe entire damaged skin between the twoadjacent ribs and by splicing in a new sectionof skin cut and formed from 24ST ale lad sheet.To gain access to the damaged skin area and toprovide for proper tool manipulation, first removethe doped fabric covering from the leadingedge and adjacent trailing edge bay or bays byfoMcwing the directions set forth in the applicableparagraphs, in this Section, on removingfabric. With a hack saw, metal snips, or a taperedreamer, cut away the damaged skin within7/8-inch of an adjacent rib flange rivetline. D''! M out the affected skin rivets aroundthe circumference and also drill out the rivetsthat attach the skin to the spar flanges. Toserve as a splice section for the leading edgeskin, cut and form a sheet of .025 inch thick24ST ale lad to match the width of the removeddamaged portion. The length of the splicesection should be 2-1/2 inches longer than theremoved portion to permit the ends to overlapthe rib flanges at least 1-1/4 inches, permittingthe new skin to be lap riveted to therib flanges. After the splice member has beencut and formed to fit the contour of the leadingedge, lay it in position over the nose ribs,and with a curved scribe, mark al I holelocations through the existing holes in theribs and skin. At each side of the damage,mark an additional rivet line on the lap areaadjacent to the rib flange. Along this rivetline, center punch and drill rivet locationsat one inch on centers. After the holes arethus marked, remove the skin splice member anddrill all existing rivet holes. Burr the rivetholes and apply one coat of zinc chronate primerto all overlapping surfaces. Reapply the splicemember and secure it in place by the use ofskin fasteners in alternate rivet holes. Insertand drive all attaching rivets, startingat the tip of the leading edge to preventthe skin from buckling as it is riveted aroundthe contour of the leading edge. After rivetingis accomplished, r^^over the exposed structurewith fabric by following the directionsset forth in the applicable paragraphs in thisSect ion.Figure 18--Damage Requiring FabricSection Replacement22. GENERAL FABRIC PATCHING.Minor patching of small tears, V-shaped tears,and small holes may be accomplished in the conventionalmanner. Such repairs, and all fabricre-covering procedures from replacement of onesection to re-covering of the entire controlsurface, are set forth in the follcwing paragraphs.When repairing tears and holes or replacingentire sections, the repairs must developthe full strength of the surrounding material.Patches should match the adjacent areas in appearance.This may be accomplished by applyingthe same number of coats of dope tc the patchas were applied to the surface originally. Sandlightly with No. 7/0 sandpaper between each coat.It is to be noted that care should be taken whenapplying additional coats of dope or finish tothe COTplete surface, as such applications addweight to the surface, causing disturbance tothe static balance. The amount of materialadded in the form of patches is limited bythe resultant fabric surface tension and bythe static balance limitations. Patches mustnever overlap another patch. If a large patchis not feasible or if the patch area must exceedhalf the bay section area, the completeaffected fabric section must be replaced.23. TESTING FABRIC FLEXIBILITY.The necessity for re-covering fabric-covered[175]RESTRICTED

I beMOVABLE SURFACESPAR. 23 TO 2 5RESTRICTEDCUT FABRICY'INCH —FROM RIBSREMOVEFINISHINGTAPES—Figure 19— Damaged Fabric Section Removedcontrol surfaces is made evident by cracked finish,loss of tautness and flexibility, excessivepatching, extensive damage, insecurity ofattachment, etc. The flexibility of the materialis an accurate index to its service expectancy.At regular intervals, therefore, theflexibility of fabric covering should be tested.This may readily be accomplished by pressing heavilyan open area with the thumb. If the dent thuscreated disappears within a few minutes, thefabric is quite new. If the dent requires anhour or more to disappear, service expectancyof the fabric is short. If the fabric crackswhere it has been dented, the finish is deadand will require re-covering. Rejuvenation inthis case, or in any other case, is not recommendedand should not be attempted. Also, partialre-covering will be required where the summed-upareas of existing patches between ribs exceed 100square inches.the edge of the tear (see Figure i^) . As theseam is made, draw the edges tightly togetherbut do not pull the thread too taut or rippingof the fabric will result. After the stitchingis completed, proceed with cleaning, doping,and patching as outlined in the following paragraphfor V-shaped fabric tears. One strip ofpinked finishing tape of the required lengthadequately serves as a patch for this type of tear.This strip of finishing tape should extend at least1-1/4 inches beyond each end of the tear.25. V-SHAPED FABRIC TEARS.If the opening of the tear is V-shaped, asis often the case, the tear may be patched bysewing the edges of the tear together and applyingdoped pinked tape over the stitches. Tackstitch the torn corner of the tear before startingthe sewing so that the edges of the fabricwi I held in place wh i le the sewing is beingaccomplished. Use a single strand of No. 8,4-ply hand-sewing cotton thread and a curvedneedle. Knot the end of the thread and startthe stitching so the knot will be left on theinside of the fabric. Start sewing at one endof the tear and continue through the tack stitchingat the corner until the complete tear hasbeen sewed. Using a baseball stitch, space thestitches not more than 1/4- inch apart and stitch1/4-inch back into the cover away from the edgeof the tear (see Figure ij, Detail A). As theseam is made, draw the edges of the tear t ight lytogether to maintain the original tautness ofthe fabric section. Do not pull the thread tootaut or ripping of the fabric will result, especiallyif the fabric has been in service andthus exposed to the elements. After the tear24. SMALL FABRIC TEARS.A satisfactory repair for small straight fabrictears may be accomplished by sewing the tornedges tightly together, using a baseball stitch.Sewing the tear prevents further ripping of thefabric during flight and should not be omittedunless the tear is less than one inch in length.Merely applying a doped fabric patch over thetear is not sufficient and will result in a lossof tautness of the fabric covering, especiallywhen two or more tears appear in one section.Use a single strand of No. P, 4'ply hand-sewingcotton thread and a curved needle. Knot theend of the thread and start sewing at one endof the tear. Using a baseball stitch, spacethe stitches not more than 1/4- inch apart andstitch 1/4- inch back into the cover away fromHAND 8COTTONFigure 20— Final Folding, Pinning,Sewing of Fabric SectionandRESTRICTED[176 J

sides.RESTRICTED MOVABLE SURFACESPAR. 25 TO 27ihas been satisfactorily sewed, clean the surfaceover wh ich the repai r tape wi II be appi ied.Remove any dirt, grease, or paint by rubbingthe surface with a rag dipped in dope, dopesolvent, or acetone. Wipe the surface witha clean rag. Care must be taken to prevent dopesolvent, thinned dope, or acetone from drippingthrough to the inside of the opposite surface,thus causing the dope to blister. To avoidthis, tip the surface straight up, or turn theunit over and apply from the lower side. Toremove satisfactorily the clean and semipigmenteddope of the original finish, apply awet coat of thinned dope to the patch area(see Figure m , detail B). Allow the thinneddope to soak into the fabric for 10 minutes.After the softening of the old dope of the originalsurface has taken place, scrape off allapplied and original dope with a dull knifeFigure 22— Rubbing Down Reinforcing Tapiric section, proceed as follows: Remove theFigure 21— Detail of Fabric Section Stitching finishing tape from the ribs on both sides of(see Figure 27, Detail C). Cut a patch of p redopedat Fabric Section Seamspinked tape or fabric of sufficient sizeto cover the tear to a point at least one inch be repaired by removing the damaged fabric andbeyond the tear in all directions. Predoped sewing in a patch. In all cases of badly frayed2-1/4 nch pi nked f in i sh ing tape will often tears, large tears, or large holes, cut out theserve satisfactori ly for patching V-shaped tears. damaged portion of the fabric to a triangularorApply one brush coat of clear dope to the cleanedrectangular-shaped opening (see Figure 16).surface and press the predoped patch into positionBetter seams, binding the new fabric to the old,(see Figure 27, Details D and E) . Make can be made along the straight sides of thecertain that all patch edges are firmly stuck resultant opening. After the fabric has beendown. Apply one brush coat of dope to the entirecut out as directed, measure the opening and cutpatch (see Figure 27, Detail F) . After a piece of new fabric to extend at least 1/2- inchthis coat of dope is dry, proceed with the appiication of the second brush coat of clear of the new fabric under 1/2 — inch and sew to thebeyond the opening on al I Fold the edgesdope, the spray coat of clear dope, and the two edges of the opening. Temporary stitches at eachspray coats of semi pi gmented aluminized dope corner facilitate seam stitching. Use a singleas directed in a subsequent paragraph.strand of No. 8, 4-ply cotton hand-sewing threadand space the stitches close together. Use an26. URGE TEARS AND HOLES IN FABRIC.overcast or a herringbone stitch. The tautnessof the patch fabric should be such that,Large tears and holes in fabric coverings may after doping, the surface tension will equalthat of the original surface. Proceed withsewing, cleaning, tape application, and dopingas outlined in the preceding paragraph. Nomore than one large hole should be repaired inany one section.27. FABRIC SECTION REPUCEMENT.Occasions may arise when an entire fabric baysection between ribs, from leading edge to trailingedge, must be replaced because patches overlapo r exceed one-half the bay section area,resulting in static unbalance or loss of surfacetautness. Fabric sections should also bereplaced if the damage is large and the patcharea would exceed 100 square inches (see Figure18). To remove and replace a damaged fab-L 177]RESTRICTED

I ingI extendeadI belingI theII edgesingMOVABLE SURFACESPAR. 27 TO 28RESTRICTEDFABRIC CUT 1/2FROM RIBIREINFORCINGTAPE LEFT ONLAST RIBFigure 23"' Applying Finishing Tape on FabricSection Seamsi i Ithe damage by pu I away from the dam.agedarea. Pu I toward the damaged area wi Iresult in darrage to adjacent sections. Removethe trailing edge tape in a similar manner.Draw a line 1/2-inch in from each adjacent riband cut out the damaged section along thesenes from the i ng edge to the t ra i I i ngedge (see Figure ig) . Pry up and turn backthe reinforcing rib tape. Cut a piece of newfabric which wi I from a point 2 inchesbeyond the centerline of the leading edge toa point one inch beyond the centerline of thetrailing edge. The fabric must be wide enoughon the leading edge section to match the pinkededges of the finishing tape for that section.Allow for a 1/2- inch fabric turn-under for sewingat the rib section. Apply a brush coatof dope to that part of the nose section to bere-covered; align the new fabric and press intoposit ion. Al low to dry. Pu Irib sectionof the fabric taut and pin in position. Spacethe pins evenly around the edge of the sectionso as to maintain a uniform tightness. Startingat the spar^ or just aft of the leadingedge skin, turn under the remaining strip oforiginal fabric along the rib; turn under theedge of the new fabric as mentioned above andstart sewing at the edge of the leading edgeskin. This wi Ijust aft of the spar inthe case of the ai leron. Use a herringbone ora baseball stitch and sew toward the trai lingedge, removing pins as the sewing progresses(see Figure 20). The seam must be close enoughto the rib to come under the edge of the ribfinishing tape or, more preferably, under theedge of the reinforcing tape (see Figure 21).Follow the same process for the opposite rib.Remove the pins along the trai ling edge; applySTITCHINGTAKEN OUTFigure 24— Typical Removal of FabricFrom Two or More Sectionsa coat of dope to the trai I edge and stickdown the remaining edge of the section, a I lowingthe fabric to extend beyond the centerlineof the trai ling edge at least one inch. Aftera I of the new fabric section have beeneither sewed or doped in place, apply one brushcoat of dope over the stitching along the ribsand rub down the reinforcing tape (see Figure22). Allow to dry. Apply a brush coat of dopeover the entire new panel; and while the dopeis still wet, lay the finishing tape on theribs and on the leading and trailing edges(see Figure 23). Apply drain grommets if necessary.Follow routine doping and paintingprocedure set forth in a subsequent paragraph.Allow to dry and cut out the centers of alldrain grommets.28. PARTIAL FABRIC RE-COVERING.In the event that two or more adjacent fabricbay sections are damaged, or if the damagingprojectile or object has gone through both sidesof the fabric covering, the static balance ofthe control surface miay be more easi ly retainedby removing the entire damaged portion of theold fabric surface and by re-covering the exposedstructure. When re-covering, utilizethe presewed^ partial envelope method. Cuttingthe old fabric may be accomplished by two differentmethods, and the method used will be dependentupon the location and nature of theRESTRICTED[178]

PhillipsI IngingRESTRICTED MOVABLE SURFACESPAR. 28 TO 29damage. When partial re-covering is practicaland when the damage is near a rib, remove thefabric from both sides of all sections thatare to be re-covered (see Figure 24). Removethe finishing tape from the area concerned,taking care to pull towards the undamaged portionof the surface when removing the tape adjacentto the remaining portion. This preventsdamage to the adjacent fabric sections. Afterthe finishing tape is removed, screw out allfabric attaching rib screws (see Figure 21^).With a sharp knife, cut the stitching aroundany hinge or access cutouts present in the sections.Do not remove the screws and reinforcingtape from the rib enclosing the adjacent undamagedsection of the fabric covering. Witha sharp knife or a pair of scissors, cut thefabric to within 1/2-Inch from this rib (seeFigure 24}. Fold this fabric under In a straightline along the rib. Fabric attachment alongthis ri b wi II follow the method set forth andillustrated in the preceding paragraph. Preparea partial envelope cover from new fabriclarge enough to cover the stripped area of thestructure and long enough to lap over the undamagedarea at least 2 Inches. This extrafabric is needed for pinning, folding, and sewingoperations. Slip the partial envelope coverover the structure and pull taut (see Figure 26).Pin the edges in place and proceed with thehand sewing as set forth and Illustrated inthe corresponding paragraphs in this Section.Attach the reinforcing tape to all ribs underthe new fabric, using the screws and dimpledwashers that were previously removed. Applyall finishing tape, canvas and fabric patcheswith the second coat of dope. For the properlocation and dimensions of tape, patches, anddrain grorrmets, refer to the corresponding fabricre-covering diagram (see Figures 21, 2^,and 2g) . Follow routine doping and paintingprocedure set forth in a subsequent paragraph.If the damage is not near a rib, the fabricmay be cut further from the rib and the upperand lower fabrics sewed together (see Figure2^). Pinning, sewing, and doping proceduresare the same as previously outlined,29. REMOVAL OF ENTIRE FABRIC COVERING.To remove the old fabric covering f rem theentire structure, first tear the trailing edgetape free. Then lift the rib finishing tapeand tear it free. Remove all other finishingtape around cutouts and over seams. Belowthe rib finishing tape are located Phillipshead, countersunk, self-tapping screws and dimpledwashers sunk Into recesses in the rib capstrip.Extract these screws by unscrewing themwith a No. Iscrewdriver. Care mustbe taken to screw out and not pull out all ribscrews. Pu I the screws out enlarges theholes In the rib, and a larger screw must beused when replacing the fabric. These largerscrews add weight aft of the hinge line and theiruse should therefore be avoided. Pass a sharpknife about the trai I edge to cut a I I machineand hand stitching. Drill out the AD3 countersunkrivets In the two metal riveting stripsat the trim tab cutout. Peel the covering carefully forward, cutting attaching threads aboutthe circumference of access holes and hingecutouts as necessary. After the covering is removed,clean the frame thoroughly with dopethinner and a clean cloth. Inspect the entirestructure carefully for damage.Figure 26—Pulling on Partial FabricFigure 25— Removing Fabric Retaining Screws Envelope Cover/[179]RESTRICTED

MOVABLE SURFACESFIGURE 27RESTRICTEDSECT/ONc-cMETAL RIVETING STRIP(REF.)2-1/4" PINKED TAPE (REF)2-1/4" DIA. FABRIC PATCHBI2I9 AD3 RIVET 25 REQ.WASHER (REF.)LING EDGESEWN WITH NO. 8-4THREAD, 21 YDS. REQ."PINKED TAPE8 YDS. REQ2-1/4 PINKED TAPE (REF.)46-3 WASHER55 REQ./8" DIA. DRAIN HOLE1/4" PINKED TAPE 39YDS. REQCOTTON FABRIC GRADE42" WIDE X 2 YDS,68" WIDE X 4 YDS7- 1/2" X 8" FABRIC PATCHPARKER KALONSCREW TYPE "a"155 REQ.'*^ASECTIONBBFABRIC (REFJ7-1/2" X 8" FABRIC PATCH(REF.)2-1/4" PINKED TAPE (REF.)6" X 8-1/2 FABRIC PATCH 2 REQ55-24044REINF ASSEM.MACHINE THREAD NO. 20-4 LEFTBIlOO WASHER (REF) 30 YDS. REQ.^CUT 3/8" DIA. HOLE THRU FABRICDETAIL A ROTATED 180°=%IIOO WASHER 4 REQ.Figure 77—Rudder Fabric Covering RequirementsRESmiCIEDL 180]

RESTRICTED MOVABLE SURFACESFIGURE 28PINKED TAPE2 5-" WIDE X 16 YDSFABRIC PATCH4f-X 4:i"SEW OPENINGSAND CUTOUTS WITHHAND SEWING TH'D.NO. 8-4 LEFT (20YDS. REQ'D.)DRAIN WASHERS(30 REQ'D.)PINKED TAPE2t"X 18" (I REQ'D.)FABRIC PATCH 6" X Ij"PINKED TAPE zV X 26t"(I REQ'D.)PINKED TAPE 2i^"X 7" (I REQ'D.)DETAILPINKEDTAPEDETAILFABRIC(REF.)Figure 28— Elevator Fabric Covering Requirements[181]restr:cted

MOVABLE SURFACESFIGURE 29RESTRICTEDMERCERIZED COTTON FABRIC42" X 3 YDS. (I REQ.)SECURE TO RIBS PER DETAIL "BBIlOO DRAIN WASHER(10 REQ.) LOWER SURFACE ONLYDETAILFigure 29— Aileron Fabric Covering RequirementsRESTRICTEDL 182]

i neingI thengRESTRICTED MOVABLE SURFACESPAR. 30 TO 31ITCHING10. 20-4MACHINE THREADX^^liifRCERIZ^DI^ COTTONFABRICthe trailing edge, spacing the pins close. Pullthe fabric more in a spanwise direction thanin a chordwise direction. With a soft pencil,mark the fabric between and upon the pins. SI ipthe cover off the small end of the structureand sew two seams along the penciled marks, removingthe pins as required, bnd the machinestitching where indicated on the diagram or atsuch a place as to permit the covering to beslipped over the narrower end of the frame.Trim the fabric covering to within |/4-inch ofthe machine-stitched seam and turn inside out.The envelope cover is now ready to be pu I ledover the structure.FLATSEAMFELLFELL SEAMFOLDING ATTACHMENT31. ATTACHMENT OF THE NEW FABRICENVELOPE COVER.Figure 30 — Fabric Double's ti tchin g Machine30. PREPARING NEW FABRIC ENVELOPE.The preparation of a new fabric envelope coveringfor the aileron, rudder, or elevator willdiffer in some respects, as each structure isof a different shape. These differences willvary the amount of possible machine stitchingfor each cover. AM machine-sewed seams arepointed out on the corresponding illustrations(see bigures 27, 28, and 2g)- Materials requiredand their respective quantities are alsospecified in these diagrams and in the closingparagraphs of this Section. The cover shouldbe prepared and the seams so made that warpthreads (parallel to the selvage) lie parallelto the I of f li ght of the a i rpl ane. Fromthe appropriate diagram and the available fabric,determine the number of fabric widths requiredto cover the span of the surface. Sewthem together by means of a double-stitchingmachine equipped with a doubl e-fo Id ing attachment(see Figure 30J. If this type of equipmentis not available, the folding can be doneby hand and the stitching done on a regularsingle-stitching sewing machine. Care shouldbe taken to ensure that the stitching is uniformand goes through all four layers of thefold. The rows of stitches must be at least1/ 16-inch from the edge of the fold and the rows1/4- to 3/8-inch apart. The cover should bemachine-sewed in the form of an envelope. Trailingedge seams that may be machine-sewed areindicated on the re-covering diagrams (seeFigures 27, 28, and 29J. Pass the fabric aboutthe frame trailing edge, up and over the leadingedge, then back on the opposite side to thetrailing edge. Pull the fabric evenly to requiredtautness and pin the edges together atSI ip the presewed envelope cover over thesmall end of the structure and pull taut. Theappi ication of powdered purnice to the frame wi IIfacilitate pulling the fabric along the leadingand trailing edges of the structure where thefriction is greatest. A small rough rubbersheet may be used as a further aid in pullingor pushing the fabric covering along the leadingand trailing edges of the frame (see Figure32J. Press the sheet firmly upon the leadingedge and rub the fabric in the required direction.Keep the pull on the leading and trailingedge fabric as nearly equal as possible andwork toward the large end of the structure.Pull the fabric taut around the large end ofthe structure and pin the fabric together withstraight pins spaced not more than 2 inchesapart (see Figure 2^^' ^'^ around the entireunsewed end of the cover and along thet rai I edge unt i mach i ne st i tch i isreached. Care must be taken to prevent theROUGH PIECEOF RUBBERFABRIC SPLICESEAM^, ^^m MACHINE STITCHING~^^ STARTS HEREFigure 31 — Pulling New Fabric CoverOver Control Surface S t rue ture[ 183]RESTRICTED

I curvedstitch,I hingePhi.MOVABl'^ SURFACESPAR. 3 TO 32RESTRICTEDIavailable, reinforcing tape, finishing tape,canvas and fabric patches may be cut from dopedfabric by the use of a hand pinking machine(see Figure 57J . Pinking may also be accomplishedwith pinking shears if a machine ofthis type is not available. Apply a coat ofdope to the fabric on each rib; and while thisdope is still wet,- center and pin a length ofone-inch predoped reinforcing tape over each ribof the trailing edge. Attach the fabric coveringto all ribs by inserting the Parker-Kalonself-tapping screws into the dimpled washersand then into the holes in the recesses providedin each rib capstrip (see Figure ^S)Use a pneumatic screwdriver with a No. I I ipshead bit if available (see Figure 40). Forthe proper clutch setting, experiment withidentical screws in the same size holes drilledin scrap metal of the same type and thickness.Figure 32—Pinning Unsewed End of Fabric Envelopefabric from si ipping whi le being pinned and thuscausing an uneven pull. The pull or tensionon the fabric along the open end and the trai I-ing edge should be uniform. After the unsewedend of the fabric envelope has been pinned inthis manner, trim the excess fabric with a pairof regular or pinking shears. Cut the fabricon a line one inch from and parallel to the rowof pins (see Figure 53J . Apply a coat of dopearound all hinge and access cutouts and alongthe trim tab channel. Allow this dope to drybefore removing the pins and stitching the unsewedend. This action retains fabric tautnesswhen the fabric is cut around all such cutouts.Pull the fabric taut around the trim tab opening,and pin the fabric (see Figure ^4). Cutthe fabric diagonal ly into each corner of thetrim tab opening (see Figure ^^J . Fold thefabric under the attaching flange of the trimtab channel spar and place the metal rivetingstrip in alignment with the rivet holes in theflange. Punch holes through the two layers ofthe fabric at each rivet hole, and insert anddrive an AN426-AD5 rivet in each hole (seeFigure ^6). Trim excess fabric to within 1-1/2inches of the metal riveting strip; apply onecoat of dope to the remaining fabric strip;turn the fabric back around over the rivetingstrip and up on the fabric surface that coversthe trim tab channel spar. This detail is setforth on the re-covering diagram for each surface(see Figures 27, 28, and 2q) . Obtainor cut all reinforcing and finishing tape requiredfrom predoped fabric. Quantities requiredand dimensions are listed on the re-coveringdiagrams. If rolls of pinked tape are not32. HAND SEWING FABRIC ENVELOPE AFTER AHACH-MENT.Hand sewing is required along the trailingedge where the machine stitching of the envelopecover terminates, at both ends of the trim tabcutout, and around a I and access cutouts.To hand-sew the trailing edge and thepreviously pinned end of the envelope cover,use a sma I sack needle and a singlestrand of No. 8, 4-ply cotton hand-sewing thread.Use a basebal I providing a minimum offour stitches per inch. Remove two or more pinsas necessary and fold under the excess materialremaining after the trimming operation. Startat the point where the machine stitching stops;fold under the material in such a way that, whenCUT FABRICI" FROM PINS-Figure 33— Trimming the Unsewed End ofFabric Envelope After PinningRESTRICTEDL184 J

theizingI formtherowI clear.RESTRICTEDthe stitching is made, the edges wi M be pulledtightly together. Knot the sewing every 6inches. Proceed with the hand sewing untilthe leading edge is reached (see Figure 43JAt both ends of the trim tab cutout, use anoverthrow or roll stitch; and starting at thetrailing edge, fold the excess fabric under andsew along the edge of the rib unti I trimtab channel is reached (see Figure gpj . Atthe hinge bracket cutouts and the access cutouts,small holes in the metal edges are providedfor the stitches and a different methodof stitching must be utilized. Cut the fabricaround the hinge cutout to within 1/2-inch ofthe metal edge, and fold this fabric under untiI rol led edge is even with the metal edgeof the cutout. This wi I a double thicknessaround the cutout and a smooth rol led edgefor sewing. Sew the fabric edges to the metaledge of the cutout, uti I the smal I ofholes provided for the stitching (see Figure41). Pass the needle through the fabric andthe hole in one direction and back through theadjacent hole in the opposite direction for onecomplete circuit around the edge; then reversethe direction of travel and make another completecircuit so that the sewing appears tobe a continuous thread. Simi lar stitching isutilized at the inboard hinge cutout on theaileron. Here, however, the fabric is foldedand this double thickness doubled under themetal edge, forming four thicknesses of fabricto sew through. See the correspondingdetail on the aileron re-covering diagram.33. DOPING FABRIC - GENERAL.Whenever possible, all doping should be accomplishedin accordance with accepted prac-HAND SEW FROM M TO NRIVETINGSTRIP—"MACHINEFigure 35--Cutt ing Fabric atMOVABLE SURFACESPAR. 32 TO 33STITC^Trim Tab Cutouttice at room temperatures of from 20° to 23°C(68° to 73°F). Relative humidity should notexceed 50 percent. When small fabric tears arerepaired while the control surface is stillmounted on the airplane, accomplish all dopingoperations during the day when the temperatureis near 70°F and when there is very littlemoisture in the air. This is usually between10 a.m. and 4 p.m. in most localities. If dopingis done during the hottest part of the dayor if wind is prevalent, the dope wi II dry toofast and blushing will result. If it is necessaryto apply dope under these conditions,add a relative amount of retarder to the dopebefore application. Retarder slows the dryingaction of the dope and permits the chemicalcomponents of the dope to penetrate the fabricbefore drying. If blushing is noted after appMeat ion, remedy this condition by sprayinga thin "dry" coat of dope thinner over the areaThe same effect may be obtained by wiping thearea lightly with a clean cloth saturated withdope thinner. All major doping operations suchas fabric section replacement, partial fabric recovering,and re-covering of the entire surfaceshould be accomplished in a doping room to ensuresatisfactory results and a smooth, tautsurface. Th i n a I and aluminized dopewith lacquer thinner (du Pont formula No. 3^P5)until the specific gravities, measured in Baumedegrees, are in accordance with the followingtable.DOPEDEGREESAT 2I°CBAUME(70°F)Figure 34^- Fabr ic Pinned at Trim Tab CutoutALUMINIZEDCLEAR (brush)CLEAR(spray)2932-1/4THIN BRUSH DOPE 1:2L 185]RESTRICTED

MOVABLE SURFACESPAR. 33 TO 35RESTRICTEDA tolerance of 1/2 Baunife degree is permissible.For the two final spray coats, n)ix the a Inized dope as follows:INGREDI ENTurn i-

utteI necessitateigibleingI adequatelyRESTRICTED MOVABLE SURFACESPAR. 35 TO 36hinge centerline. In the original constructionof the control surfaces for this airplane,an unbalance condition exists; that is, theweight aft of the hinge line exceeds that forwardof the hinge line. Because of the relativelew speed of this type airplane, this conditionis not critical. However, definite unbalancelimits are set up which must not beexceeded. (See Figures ^9, ^0, and 52 . J Theaddition of weight to the surfaces in the formof repairs may cause the unbalance limits tobe exceeded, and is therefore a possible causeof f I ri ng or dest ruct i ve osc i I lat ion ofthe surface in flight. Minor patching of fabricsurfaces, which adds neg I weight tothe control surface, will not be cons ide redcause for checking the unbalance. It is to beremembered that the nearer the patch is to thehinge line of the surface, the less will beits effect on the unbalance. Large, heavy,FOLDUNDERRUNNING STITCHFigure 38—Attaching Fabricto Ribsor numerous patches, especially those on ornear the trailing edge, should therefore beavoided by section replacement or partial recovering.After any sizeable repairs, mostof which wi I removing the surfacefrom the airplane in order to perform properlyall re-covering and doping operations, checkthe unbalance before remounting the surface onthe airplane. It is necessary that the unbalancebe held within the specified limits toavoid possible fluttering of control surfacesat high speeds. Prior to the accomplishmentof any balance check, the surfaces must be completelyassembled and finished, with balanceweights, hinge fittings, trim or booster tabs,and trim or booster tab operating rods allin place. After the surface is thus prepared,use the method outlined in the fo Mewing paragraphto check the unbalance.36. DETERMINING CONTROL SURFACE UNBALANCE.I I necessitateTo ascertain whether or not the unbalanceof the control surface is within the specifiedlimitations, the surface must be mounted sothat it lies in a level, horizontal positionand moves freely on its cwn hinges. Inasmuchas the unbalance need not be checked exceptafter s izeab le repa i rs wh ich w iremoval of the surface from the airplane,checking of the unbalance while the surfaceis mounted on the airplane should not be attempted.A balancing Jig, consisting of twowelded angle iron standards clamped or boltedto a level table, wi Iserve for thispurpose (see Figure 48). The angle iron standardsmust be so made that the hinge points will belevel and will not bind against the structureat any point. The surface when mounted in thismanner must be free to nove of its own inertiaon its own bearings. Place a balance scalethat is calibrated in ounces under the trailingedge of the structure at the inboard endof the trim tab cutout on the elevators andai lerons, and at the lower end of the trim tabcutout on the rudder (see Figures 49, ^0, and 52 J.It must be remembered that the weight of anyblocks or device placed on the scale to levelthe structure must be subtracted from the totalweight reading on the scale in order to ascertaincorrectly the unbalance weight of thetrai I edge. Refer to the corresponding diagramand to the fol lowing table for max imum al lowableunderba lance conditions (see Figures 49,50, and 52 J.STRUCTUREELEVATORRUDDERAILERONFigure 39~'Hand Sewing atMAXIMUM SCALE READING2 LBS. 2 OZ.2 LBS. 4 OZ.13 OZ.Trim Tab Cutout[IB7]RESTRICTED

MOVABLE SURFACESPAR. 36 TO 37RESTRICTEDFigure 42-^ Apply ing Brush Coat of Dopeto FabricDOPED AND PINNEDREINFORCING TAPEFigure 40 — Use of Pneumatic Screwdriver toAttach Fabric to RibsIf the weight exceeds these limits, the balanceis not satisfactory and the surface mustnot be used until corrective action has beentaken. Methods of correcting static unbalanceare set forth in the following paragraphs.37. CORRECTING UNBALANCE WHEN ABOVE LIMIT.centerline. Addition of weight in the formof long, heavily doped patches applied to theleading edge of the aileron or to the balanceweight housing of the rudder or elevator wi I Isometimes bring the unbalance weight within thelimit. To apply such a balance patch to theaileron, apply a heavy coat of unthinned dopeto the leading edge of the aileron. Place thefabric patch in position over the wet dope andrub down firmly until the dope has penetratedthe fabric (see Figure 41)-. Allow to dry beforeapplying any additional coats of dope. Ifnecessary, a total of six coats of dope may beIf, after following the instructions in thepreceding paragraph, the unbalance weight exceedsthe maximum allowable limit, this conditionmay often be corrected by the additionof weight forward of the control surface hingeHAND SEWINCOTTON THRFigure 41'^Hand Sewing Fabric at Hinge CutoutFigure 43'^Hand Stitching the Unsewed End ofFabric Enve lopeRESTRICTEDL 1B8]

I imitredi nedRESTJUCIED MOVABLE SURFACESPAR. 37 TO 38mFigure 44— Applying Finishing Tape to Fabricapplied to each patch, and a total of threelong patches may be applied to the same area.Continue adding patches and dope to the leadingedge until the reading on the balance scaleis within the specified limitations. If, afterthe application of three patches to the leadingedge, the unbalance' is still beyond thebecause of numerous patches aft of the spar,replace the entire affected fabric section orsections if feasible; otherwise, re-cover theentire surface and recheck the static balance.The procedure for the rudder and elevator fol-RETAININGSTRIPRIVETEDTHANDSEWNFINISHINGTAPESCANVASa FABRICPATCHESFigure 46— Applying Spray CoatsDope to Fabriciof Aluminizedlows that out I above with the exceptionthat the balance patches are applied to thebalance weight housing instead of the leadingedge (see Figure 52).38. METAL AND FABRIC REPAIR TOOLS.In addition to the standard sheet metal cutting,forming, bending, and riveting tools,all or part of the fol lowing tool s will be requi for structural repairs to a I I membersand for repair to the fabric covering of tneelevator, rudder, and ailerons.HEAVY COATOF WET DOPEHANDSEWNREINFORCING TAPES ATTACHEDTO ALL RIBS/^EAOr FOR FINISHINGTAPES a PATCHESFINISHING TAPES 3PATCHES APPLIEDFigure 45—'Fabric Envelope Prior to First andSecond Coats of DopeLONG AND WIDEFABRIC PATCH —Figure 47—-Apply ing Leading Edge Balance Patchto Aileron[ IB9]RESTRICTED

MOVABLE SURFACESFIGURE 1*8RESTRICTEDrHINGE POINTSMUST BE LEVELnBALANCE WEIGHT HOUSINGUSE CLAMPS TO SECURESTANDARDS SO THEY CANBE MOVED TO FIT OTHERSURFACESLEVEL TABLEHINGEPOINTMOt/AfTCONTROL SURFACE AS SHOWNPLACE TIP OF BLOCKSADJACENT TO TRIMTABWELDED ANGLE IRONDETAIL ATOP OF BLOCKS MUSTBE LEVEL WITH HINGESBALANCE SCALETRIM TAB a OPERATINGROD INSTALLEDNOTE'SUBTRACT WEIGHT OFLEVELING BLOCKS FROMTOTAL WEIGHT READINGCHECK BALANCE AT TRAILING EDGEFigure 48— Determining Control Surface UnbalanceRESTRICTEDI 190]

istJIRESTRICTED MOVABLE SURFACESPAR. 38 TO i|0TOOLREMARKSMATERIALAMOUNTREQUIREDAAF SPEC.EQUIPMENTSHEARSMACHINE,MACHINE,KNIFEBRUSH,SEWINGPINKINGDOPENEEDLESSCREWDRIVERSDOPESPRAYINGPLAIN AND PINKING(See Figure 30,)(See Figure 37.)WIDTH TOSUITCURVED AND STRAIGHT TO SUITNO. 1 AND NO. 2 PHILLIPSTAPE, PINKEDTAPE, PINKEDCANVAS, 10 OZ.PATCH, FABRIC2-1/4 IN. X 40YDS.1 IN. X 8 YDS.3 IN. X 3-1/2IN. (2 REQ.)2 IN. X 9-1/26-6 2 B6-62BCCC-0-771TYPE IAN-CCC-C-399I N.PATCH,FABRIC4 IN. X 18 IN.AN-CCC-C-399(2 REX?.)PATCH,FABRIC7-1/2 IN. X 8AN-CCC-C-399IN.TRIM TAB (REF)PATCH,FABRIC8 IN. X 10 IN.AN-CCC-C-399PLACE BALANCE SCALEEXACTLY AT THIS POINTPATCH,FABRIC6 IN. X 8-1/2IN. (2 REO-)AN-CCC-C-399REACTION ON BALANCE SCALEMUST NOT EXCEED 2 LBS. 40Z.Figure 49'^Rudder Unbalance LimitsPATCH,PATCH,FABRIC12 IN. X 12 IN.2-1/4 IN. DIAM.(2 REQ.)AN-CCC-C-399AN-CCC-C-39939. METAL STRUCTURE REPAIR MATERIALS.The fol lowing is aI of materials neededto repair properly the metal structures of elevator,rudder, ailerons, and landing flaps.MATERIALSSHEET, 24ST ALCLADRIVETS, AN4U2-ADURIVETS, ANU26-AD4RIVETS, AN426-AD3RIVETS, B1227-AD4RIVETS, LS1127-4-4RIBS, REPLACEMENTW. RUDDER COVERING MATERIALS.REMARKS.020 TO .040 INCH THICKFLAT HEADCOUNTERSUNKCOUNTERSUNKBRAZIER HEADCHERRY BLIND RIVETS(See Figures 2, 3 , 5 , and 7The following materials are required for thecomplete re-covering of the rudder. (See Figure2J.) Accomplish all patching or partialre-covering by the use of the same rraterials insmaller quant it ies. All patch materials shouldconform to the same specifications as requiredfor the original covering of the structure.MATERIALAMOUNT REQUIRED42 IN. X 2 YDS.68 IN. X 4 YDS.AAF SPEC.AN-CCC-C-399RE INF. ASSEM.55-24044WASHER,BllOOWASHER, B1246-3SCREW,KALON, NO. 4,PHILLIPS HEAD,TYPE A, C'SUNKSCREW,SCREW,B1286-832-7B1287-832-6THREAD, MACHINE,N0.20,4-PLY,LEFTTHREAD,SEWING, NO. 8,4-PLY,DOPE,NITRATE,LEFTCELLULOSECLEARFABRIC, MERCER-IZED COTTONPARKER-HAND-PASTE, ALUMINUM-PIGMENT, AIR-CRAFT1 REQ.4 REQ.155 REQ.12 REQ.2 REQ.30 YDS.21 YDS.V-T-276b,TYPEIBIV-T-2 76b,TYPE II IB155 REQ-AS REQUIRED AN-TT-D-514[See correspondingPar.)AS REQUIRED AN-TT-A-461(See correspond"ing par.)Ll9l]RESTRICTED

)MOVABLE SURFACESPARAGRAPH 11RESTRICTEDm. ELEVATOR COVERING MATERIALS.tvbterials required for the complete re-coveringof each elevator section are listed below.For location and dimensions of tapes and patches,refer to the elevator re-covering diagram. (SeeFigure 28.MATERIALAMOUNT REQUIRED AAF SPEC.FABRIC, MERCER-IZED COTTONTAPE, PINKED42 IN. X 1.4 YDS. AN-CCC-C-39968 IN. X 2.6 YDS.2-1/4 IN. X6-62B17-1/2 YDS.PLACE BALANCE SCALEEXACTLY ATTHIS POINTREACTION ON BALANCE SCALEMUST NOT EXCEED 2LBS. 20Z.TAPE, PINKEDPATCH, FABRIC1 IN. X 8 YDS.4-5/8 IN. X4-5/8 IN.6-62BAN-CCC-C-399Figure Sl'-^Elevator Unbalance LimitsMATERIALAMOUNT REQUIREDAAF SPEC.PATCH,PATCH,PATCH,FABRICFABRICFABRIC4-1/4 IN. X4-3/4 IN.(2 REOO1-3/4 IN. X9-1/8 IN.5-1/8 IN. X9-1/8 IN.AN-CCC-C-399AN-CCC-C-399AN-CCC-C-399PATCH,PATCH,FABRICFABRICRE INF. ASSEM.55-22020WASHER,BllOO6 IN. X 7-1/2 IN. AN-CCC-C-3992-1/4 IN. DIA. AN-CCC-C-3991 REQ.30 REQ.PATCH,FABRIC8 IN. X9-1/2IN.AN-CCC-C-399WASHER, B1246-3162 REQ.PATCH,FABRIC4-3/4 IN. XAN-CCC-C-399SCREW,PARKER-162 REQ.PATCH,FABRIC2-5/8 IN.2-1/2 IN. X2-3/8 IN.KALON, NO. 4,PHILLIPSHEAD, TYPE A,C'SUNKPATCH,FABRIC7-1/2 IN. X6-1/4 IN.SCREW, B1286-832-78 REQ.R.H. AILERON SHOWNSCREW, 81 287-832-62 REQ.TRIM(REF.)TABTHREAD,MACHINE,NO. 20, 4-PLY,LEFT25 YDS.V-T-276b,TYPE IB ITHREAD,HAND-20 YDS.V-T-276b,PLACE BALANCESCALE EXACTLYAT THIS POINTREACTION ON BALANCESCALE MUST NOT EXCEEDFigure SO-^A Heron Unbalance LimitsSEWING, NO. 84-PLY,DOPE,NITRATE,PASTE,LEFTCELLULOSEALUMI-CLEARNUM-PIGMENT,AIRCRAFTAS REQUIRED (Seecorres. Par.)AS REQUIRED (Seecorres. Par.)TYPE II IBAN-TT-D-514AN-TT-A-461RESTRICTEDi 192]

RESTRICTED MOVABLE SURFACESPARAGRAPH ^2DOPEMATERIALFABRIC, MERCER-IZED COTTONAMOUNT REQUIREDAAF SPEC.42 IN. X 3 YDS. AH-CCC-C-399TAPE,PINKED2-1/4 IN. X46-1/2 YDS.6-62BTAPE,PINKED1 IN. X 7 YDS.6-62BWASHER,BllOO10 REQ.BALANCE WEIGHTHOUSINGLARGE FABRICPATCHWASHER, B1246-3SCREW,PARKER-KALON, NO. 4,PHILLIPS HEADTYPE A, C'SUNK141 REQ.141 REQ.Figure 52'^Applying Balance Patch toElevator and RudderTHREAD,MACHIKE,NO. 20, 4-PLY,LEFT30 YDS.V-T-276b,TYPE IB I^2. AILERON COVERING REQUIREMENTS.The materials listed in the fol Icwving tableare required for the complete re-covering ofeach aileron. For location and dimensions oftape and patches, refer to ai teron re-coveringdiagram. (See Figure 29. J Accomplish allpatching or partial re-covering by the use ofthe same materials in the required smallerquantities. All patch materials should conformto the same specifications as required forthe original covering of the structure.THREAD, HAND-SEWfNG,NO. 20, 4-PLY,DOPE,LEFTNITRATE,PASTE,PIGMENT,CELLULOSECLEARAIRCRAFT15 YDS.ALUMlNU^f-AS REQUIRED(See correspondingPar,)AS REQUIRED[See corresPond, ingPar JV-T-276b,TYPE I I IBAN-TT-0-514AK-TT-A-461L 193]RESTRICTED

ikeumi.tankRESTRICTED FUEL i OILPAR. I TO 3SECTION 5FUEL AND OIL SYSTEM1. GENERAL.The two removable fuel tanks are housed withinthe wing centersect ion structure and areheld securely in position by three felt-paddedmetal straps around each tank. (See Figure1.) The tanks are installed between the frontand rear spars in the wing centersect ion structure,and access to them is gained by removingthe fuel tank doors which are bolted to thelower side of the centersect i on structure.The oil tank, located at the top of the engineaccessory compartment, is supported on theupper tubes of the engine mount structure nearthe firewall. (See Figure 5.J Two feltpaddedmetal straps support the tank betweentwo upper cross tubes attached to the enginemount structure.2. STRUCTURAL DESCRIPTION,The fuel and oil tanks are each constructedof two power-formed aluminum alloy sections,or shells, welded together. The fuel tankshave also a number of baffles of 52S0 material,spot-welded across their inner widths. A ventedbox I compartment of 52S-I/2H material is formedover the fuel sump between two centrally locatedbaffles to retain a volume of fuel overthe sump. (See Figure 2.) The oi I has nointernal baffles but depends upon the shapeof its outer shell for its strength. (See Figure//.j A tubular, vertical compartment throughthe center of the tank serves as an oil warmingcompartment. (See Figure ^. j This seamlesstube is fabricated of 52S0 aluminum alloy andis welded into position. All fittings onthe fuel and oil tanks are cast from I95-T6a I num a I loyFigure 1 — Fuel Tank Supports3. GENERAL REPAIR.Repairs to the fuel and oil tanks consist®SUMP FUEL RETAININGBAFFLESFigure 2— FuelTank Construct ionL 194]RESTRICTED

'nstaFUEL & OILPAR. 3 TO 4RESTRICTEDof smoothing out negligible damage, sealingsmall cracks, welding split seams, and weldinginserts into holes or large cracks. Priorto welding or applying an open flame of anycharacter, steam the inside of the tank thoroughlyfor at least three hours, or carefullyflush with fire extinguisher fluid (carbontetrachloride). A one-hour period is sufficientfor the oil tank. Prior to weldingfuel tank areas located between the outermostbaffles, remove the synthetic rubber flappervalves located within the central chamber.These valves are subject to distortionwhen in the proximity of heat. Access tothe valves may be gained by removing the fuelsump.FILLER NECKCASTINGWARMINGCOMPARTMENT^. LOCATING SMALL CRACKS.The presence of minor seepage and vaipor leaksin the fuel and oil tanks may be ascertainedby air pressure testing. Support the tank inSUMP CASTING—'FORWARD SHELLFigure 3'^Oil Tank Cons true t iona manner equivalent to the original i I lationbefore starting the test. AccessibilityFILLER CAP VENT LINEADAPTEROIL INLETVENT LINEADAPTERSIDE VIEW FRONT VIEWFigure 4— OilTank AssemblyRESTRICTEDL'9^]

RESTRICTED FUEL & OILFIGURE 5I. OIL TANK 4. OIL COOLER SUPPORTSZ FRONT SUPPORT 5. OIL COOLER3. PADDED METAL STRAPFigure 5— Oil Tank and Oil Cooler Installation[ 196]RESTRICTED

FUEL & OILPAR. i| TO 5RESTRICTEDof the oil tank seams and fittings while Installedpermits the use of the original installationsupporting straps for this purpose.Supporting the fuel tanks may be accomplishedby sawing out pieces of lumber (2 inches thick,8 inches wide, 42 inches long) to the contourof the tank, padding the edges which will bein contact with the tank with felt, and securingthe ends together with strap hingesand hasps as illustrated. (See Figure 6.)Apply and maintain an air pressure of nox morethan 5 Ibs./sq.in. to the oil tank or 3-1/2Ibs./sq.in. to each fuel tank. Connect theair supply hose to the tank sump or filleropening after plugging tightly any remainingoutlets. Mix one cup of water with l/2-cupof liquid castile soap; and using a softbristledbrush, apply this solution to allseams or suspected areas. The formation ofa soap bubble indicates a leak. If a soapbubble is found, mark the location and proceedwith the examination until the entiretank has been covered. After all leaks arelocated and marked, release the air pressureand proceed with repairs as outlined in thefollowing paragraphs.5. SEALING SMALL CRACKS.Minor seepage and vapor leaks, caused byminute cracks, may be repaired by sealing witha suitable synthetic rubber mixture. Beforeapplying sealant, clean all dirt and greasefrom the area to be treated and brighten withsteel wool. The sealant used must be unaffectedby fuel and possess good adhesive qualities.LOW AIR PRESSURE SUPPLYSUMP OPENINGNOTEJIG IS OF PINEBOARD CONSTRUCTIONCUT OUT TO MAKE ALLOF SEAM ACCESSIBLE.Figure 6— FuelTank Testing JigRESTRICTEDLl97]

I toRESTRICTEDFUEL & OILPAR. 5 TO 7Dolphin No. 1625 water and gasoline resistingcement, made by the Dolphin Paint andVarnish Co., Toledo, Ohio, is recommended.Dolphin cement is supplied in a paste formand is applied by pressing it into the crackswith the fingers or thumb. (See Figure j.)Fi Met out we I a feather edge and removeall excess cement. The tank may be used immediatelyafter repair. This type of repairwill last a considerable length of time; butif the crack has a tendency to enlarge, itis-advisable to replace the synthetic compoundby means of a welded insertion-type patchas set forth in the following paragraphs. IfDolphin cement is not immediately available,du Pont Cavalprene Cement, manufactured byE.I . du Pont de Nemours ^ Co., Fairfield,Connecticut, may be used; but sufficient timemust be allowed for drying before the tankmay be used. Test for sufficiency of repairas outlined in a subsequent paragraph. Ifinsufficiency of repair is indicated, applysealant to the inner surface of the tank ina similar manner, but do not use steel woolon the tank interior. Access to the tank interioris gained by removing the tank sump.If access cannot be gained by this method,the leak should be repaired in accordance withthe instructions in the following paragraphs.SYNTHETICRUBBERCEMENTdented area from the inside. This method maybe used only when it is possible to gain accessto the interior of the tank. If thismethod is not practical, dents may sometimesbe pulled back into shape by welding a 2S0welding wire to th6 lowest point of the dent andby exerting a steady pull on this wire while applyinga torch flame evenly over the dented area.If this method is successful, no further repairis necessary. Snip off the welding wire "and filethe sharp point, homver, if this " pu II i ng-w i re"method fails, or if residual cracks or abrasionsare attendant after re-forming, cut out thedamaged area and treat as set forth in the correspondingsubsequent paragraphs.7. SPLIT SEAM REPAIR.Dripping or running leaks caused by crackedwelds along the seams may be remedied by fusingthe crack together by means of a welding torch.(See Figure 8.) Hydrogen welding facilitiesare desirable, but oxyacetylene equipment maybe used with the application of special technique.Use a tip a size larger than wouldordinarily be used with steel of the same gageand adjust the torch to produce a slightlycarbonizing flame. The torch should producea low velocity bluish flame and not be "blowy."Use 2S0 welding wire in conjunction with suitablefluxes. Concluding welding, neutralize allwe Id ingt ion ofuxes by flushing with a warm solupercentsulphuric acid and water.V /Figure7 — Sealing Small Cracks6. REMOVING DENTS.All dented areas in the fuel and oil tanksshould be restored to the original contourof the tank shell. This may be accomplishedby using a leather mallet and striking theFigure 8 — We Iding SplitSeams[ 19b]RESTRICTED

I NSERTi quidALcupatFUEL & OILPAR. 8 TO IIRESTRICTED8. LARGE CRACK AND HOLE REPAIR.Holes and large cracks which may be roundedinto holes may be repaired by welding a square,a circular, or an ovaloid insert into a flangedopening. Such repairs are illustrated in thissection. (See Figures lo and ii.) Proceed asfollows: If the hole is somewhat regular tobegin with, file it out to form a clean circleor square, whichever is the most suitable.To prevent filings from entering the tank interior,turn the tank upside down, reverse theends of the file, and file downward. Next,turn the tank to its original position andflange the edge of the hole slowly upward, asillustrated, by means of a metal bar or pipeand a heavy hammer. (See Figure lo.) Form aninsert of 2S0, 3S0, or 52S0 material over awooden block. The material should be cut tosuch size and shape as to pennit the insert,upon completion of forming, to be pressed easilyinto the flanged opening in the tank. Pressthe insert into position and apply a weld aroundthe top. Completed welded insert repairs generallyassume one of the three common shapesillustrated. (See Figure ii- ) Concludingwelding, neutralize all welding fluxes by flushingwith a warm solution of 5 percent sulphuricacid and water. Large holes, long tears, orlengthy cracks may be repaired in a similarmanner, the insert being shaped to suit thecondition. The repair of holes by welding aflat patch over the damage is not recommended,as tests have proven that such repairs failunder prolonged service. However, a flat weldedpatch may be used in the repair of holes inthe oil tank hopper, as this tube is merelya partition between the warm and cold oil.The hopper is easily removed by filing theweld around the hopper inlet plate and pullingthe assembly outward.9. HYDRAULIC OIL TANK REPAIRS.Inasmuch as the construction of the hydraulicoil tank is similar to that of the lubricatingoil tank, all repairs set forth for the latterwill be applicable. (See Figure gJ The endsof the tank may be removed by filing the weldsaround the periphery of the tank. Drain andflush the tank with carbon tetrachloride beforeattempting any repairs.IISteam the interior of the tank with live steam,and allow to cool and drain. Support the tankin a manner simi lar to the original instal I ionas set forth in a foregoing paragraph. Attachan air pressure supply hose to one tank outletand plug tightly all remaining outlets.Apply and maintain an internal air pressureof 5 Ibs./sq. in. to the oil tank or 3-1/2 lbs./sq. in. to the fuel tank. The tank shouldwithstand the specified pressure with no evidenceof distortion. Thoroughly mix l/2-cupof casti le soap with of warm water,and with a soft-bristled brush, apply thissolution to all seams, welded or riveted areas,and to all fitting edges. Carefully scrutinizethe applied solution for evidence ofsoap bubble formation. If soap bubbles arefound, indicating a leak, mark the locationand continue the examination until the entiretank has been covered. If leaks are found,repair as outlined in the preceding paragraphs.If no leaks are found after a careful scrutinyof all possible areas, the tank is ready foruse. Relieve the air pressure before removingthe supporting jig. Wash all soap solutionfrom the tank with a rag dipped in clean warmwater.55-58112058 52SO ALUM.AL55-58114064 2S0 ALUM ALREF. DWG 55-58116Figure 9—Hydraulic Oil TankII. REPAIR TOOLS AND MATERIALS.To successfully accomplish the repairs outlinedin this section, all or part of the followingtools and materials are required:10. TESTING.TOOL OR MATERIALREMARKSupon completion of repairs, thoroughly cleanand flush the inside of the tank, using a warmsolution of 5 percent sulphuric acid and water.52S0 ALUMINUM ALLOY SIZE ANDMATER I TO SU I T52S0 WELDING ROD DIAMETERTH ITOCKNESSSUITRESTRICTEDL 199]

RESTRICTED FUEL & OILFIGURE 10CIRCULARINSERTSTEP I.>^lSTEP 2.PREPARE INSERT OF 2S0, 3S0, OR 52 SOMATERIAL- FORM ABOUT WOODEN BLOCK.FORM INSERT AS SHOWN.INSERTVISEBACKING BLOCK "A" FORMING BLOCK "B" ^ASTEP 3.CLEAN OUT HOLEAND FLANGE ASSHOWN.STEP 4.WELD INSERT IN POSITION .caution! OBSERVE NECESSARYFIRE PRECAUTIONS.NSERTMETALBAROVALOIDINSERTSTEP I.SAME PROCEDURE AND MATERIALSAS ABOVE. USE OVALOID INSERTFOR LARGE HOLES, LENGTHYCRACKS, ETC.INSERTFigure10— Forming and Welding Insert[ 200 ]RESTRICTED

FUEL & OILFIGURE IIRESTRICTEDWmimmmimmmmmim§WELDED SQUARE INSERTWELDED OVAL INSERTFigure 11—-Completed Fuel and Oil Tank RepairsRESTRICTEDL2OI ]

OUNUMDDiefI NGiveRESTRICTED FUEL & OILPAR. II TO 15TOOL OR MATERIALREMARKSTOOL OR MATERIALREMARKSIALUM I ALLOY WELDI NGFLUXHYDROGEN WELDING TORCHAND TANKSSULPHURI C AC ICARBON TETRACHLORI DEL I CASTI LE SOAP2-INCH SOFT BRUSHPROCURE LOCALLYOXYACETYLENE MAYBE SUBSTITUTEDPROCURE LOCALLYPROCURE LOCALLYPROCURE LOCALLYPROCURE LOCALLYSYNTHET IC RUBBER CE-MENT, WATER ANDGASOL INE RESISTI NGRUBBER AIR HOSE ANDSHUT-OFF VALVEMETAL BAR, HAMMER,AND WOODEN BLOCKSI NSERTS.DOLPH IN OR CAVALPRENEOR EQUIVALENTUSED Wl TH TANK TEST-JIGUSED TO FORMOILCOOLERS12. GENERAL.A Type C-5 vertical oil cooler, on whichis mounted a Type 0-5 thermostatic re I valve,is located on the bottom of the engine accessorycompartment and provides automatic oiltemperature regulation for the engine. Theoil cooler assembly is supported by means ofa bracket extending from the top of the coolerand attached to the two lower engine mounttubes with rubber hose connections. (See Figure5-^ This Type C-5 oii cooler is manufacturedby United Aircraft Products, Inc., Dayton,Ohio. Repair of these coolers may be necessitatedby collapsed tubes, tube leaks, coresurface . I eaks, surface leaks between the coreand shell, assembly leaks through the silversolder bond, dents in the shell, and bulletholes in the shell or core.13. OIL COOLER CONSTRUCTION.The oil cooler is cylindrical in shape andhas a diameter of 8 inches, a depth of 9-1/2inches, a frontal area of 50.3 square inches,and a cool ing surface of 29 square feet. Enclosedby a brass shell, the 9 inches long coppercore tubes have .323-inch hexagonal ends,a .268-inch diameter, and a .006-inch wallthickness. By means of tin-lead solder, thehexagonal ends of the core tubes are held together,one tube to all its adjacent tubes,and the core assembly to the shell. When thecooler is installed, the oil seeps down betweenthe round portion of the cool ing coppertubes, the hexagonal ends of the tubes beingsealed with solder. (See Figure 12-)m. CLEANING OIL COOLERS.Prior to making any repairs, remove thecooler and drain all oil in the cooler. Heata quantity of carbon tetrachloride to the boilingpoint, pump the solution through the cooler,and strain the solution as it leaves the cooler.If bearing metal particles are found, the coolershould be replaced. After thoroughly flushingthe cooler with the organic solvent, thoroughlyclean the cooler with Isteam forone-half hour. Allow the steam to pass downwardinto the internal passages, and positionthe cooler so that the condensed steam willdrain freely from the bottom.15. TESTING OIL COOLER FOR LEAKS.If it is necessary to locate leaks in thecooler, seal all openings in the cooler exceptone, and to this opening apply an air pressurenot to exceed 75 Ibs./sq. in. Slowly submergethe cooler in warm water; and where bubbles appear,mark the location with a wire clip. Continue theexamination until the entire cooler has beencovered. Remove the cooler, disconnect thepressure line, and allow the cooler to drain.TYPE D-5THERMO-STAT! CRELIEFVALVEAIR FLOW REGULATORSLOVi/ERMOUNTING BRACESUPPERMOUNTINGBRACKETNLETOUTLETCORE FACEAIR SCOOP OUTER SHELLFigure 12—-Oil Cooler AssemblyI 202 ]RESTRICTED

ngI themayeadI assemblyFUEL i OILPAR. 16 TO 23RESTRICTED16. REMOVING SINGLE TUBES.The best method for removing leaking or collapsedtubes is as follows: (See Figure 25. iObtain at least two pieces of 5/16-inch hexagonalmetal bar stock 4 inches long, and bevelthe sharp edges from the ends. Heat thesebars in a gas furnace; and while the ironsare heating, push a 12-inch wire through themarked end of the defective tube in order toidentify the corresponding tube ends. Fluxboth ends of the tube to be removed. With asmal I pair of tongs, remove each of the heatedmetal bars from the furnace, withdraw the identifyingwire from the defective tube, and insertthe tip of one iron into the tube end.As the wire is withdrawn from the opposite end,insert the t ip of the other iron into that end.If the irons are hot enough, the solder bondshould be broken immediately. When this occurs,gently tap one of the hexagonal bars untilthe tube is pushed out the opposite end. Removethe bars from both ends of the tube; andus i a pai r of p lie rs, pu I protrud i ngend of the tube out through the core. Cleanthe adjacent tube surfaces. If any lumps ofsolder adhere to the adjacent tubes, shapethe opening in the core with the hexagonali ron bar.17. REPLACING TUBES.Replacement tubes should conform to AAF Og.065948-9. (See Figure iq. ) Flux each end ofthe replacement tube and dip the hexagonalends into molten tin-lead solder. Flux theopenings in the core where the old tube hasbeen removed and insert the new tube. Witha pair of sharp-nosed pliers, re-form the hexagonalsides of the tubes adjacent to the newtube. After refluxing, use the standard solderingiron and melt tin-lead soldering wireover the face of the core with an oscillatingmot ion of the i ron.18. REMOVING LARGE SECTIONS OF CORE.The procedure outlined for the removal ofsingle tubes should be used when removing onetube or several tubes. However, if a largesection of the core must be removed, use anoxyacetylene flame on both ends of the core.Care must be taken to avoid the use of excessiveheat when melting the solder bond fromlarge sections of the core. Use only enoughheat to melt the solder, then swiftly applypressure to remove the core section.19. LEAKS ON CORE SURFACE.Clean the area around the core surface leak,and flux the surface with zinc chloride flux.Using a t i n- 1 solder wire and a hot iron,solder over the point of the leak with an osciMating motion of the iron. Leaks betweenthe core and she I may also be stoppedby the application of solder.2C. LEAKS AROUND SHELL SEAMS.To stop leaks around the silver-soldered seamsof the shell assembly, thoroughly clean andflux the area. Using a soldering iron, melta tin-lead soldering wire over the leak withan oscillating movement of the iron.21. DENTS IN SHELL.Large dents in the shell not indicating sharpradii may be corrected by the following procedure:Apply an air pressure of from 30 to40 Ibs./sq. in. to the inside of the cooler.Utilizing an acetylene flame, carefully heatthe area affected by the dent. Under the airpressure from within, the material should returnto its normal contour. Sharp dents sometimesmay be corrected by soldering the endof the silver solder wire to the center of thedent. After this is done, gradually pull theshell back into position.22. HOLES IN SHELL.Small holes not exceeding 1/4-inch in diametermay be patched by utilizing a piece of.040 inch thick or .050 inch thick brass. Thoroughlyclean the area with steel wool, and softsolder the patch over the hole. Large holesin the she! I be repaired by silver soldering,provided the core is properly protectedfrom excessive heat by utilizing wet cloths.Holes in the inside shell are extremely difficultto repair, and the cooler should be replacedrather than repaired.23. TESTING OIL COOLERS.Upon completion of repairs, test the cooleras follows: Plug all outlets except one, andto this outlet apply an air pressure of 75Ibs./sq. in. Submerge the cooler in warm waterand check for leaks. If a satisfactoryrepair is indicated, thoroughly flush the coolerwith hot water and then apply live steam tothe cooler interior for one-half hour. If theRESTRICTEDI 203 ]

RESTRICmDFUEL & OILFIGURE 13HEXAGONAL BAR FORTUBE REMOVAL

FUELPAR.& OIL23 TO 25RESTRICTEDcooler is to be installed immediately, pumphot light engine oil through the cooler priorto installation. If the cooler is to be stored,prior to storage, heat a tank of SAE 20 engineoil to I2rc (250°F) and immerse the coolerin the oil. Vigorously agitate the cooler untilthe bubbling ceases. Remove the cooler, allowit to drain, and plug the fittjng opening.24. OIL COOLER REPAIR MATERIALS.MATERIALFLUX, ZINC CHLORIDEWIRE, SILVER SOLDERWIRE,DERTIN-LEAD SOL-TUBES, CORE REPLACE-MENTSHEET,BRASSREMARKSFOR BOTH SI LVER ANDTl N-LEAD SOLDERAAF SPEC OO-S-5611/16- IN. Dl AMETERAAF SPEC. QO-S-5711/8- IN. DIAMETERAAF PART NO. 065948-9AAF SPEC. OO-B-611.025, .040, .050 IN.THICK, HALF HARD.25. OIL COOLER REPAIR TOOLS.Depending upon the extent of repairs to beundertaken, all or part of the following toolsshou Id be aval I able.TORCH,TANK,TOOLOXYACETYLENEOXYGENTANK, ACETYLENEREGULATOR, PRESSUREFURNACE, GAS-FIREDIRONS, SOLDERINGBAR, 5/16-1 N. HEX.METALSMALLREMARKSTWO LARGE ELECTRICOR GAS-HEATED IRONSFOR REMOVING TUBES,TWO BARS 4 IN. LONGREQUIRED.RESTRICTEDI 205]

nfiner,of.icaI -cantRESTRICTEDLANDING GEARPAR. I TO 3SECTION 6LANDING GEARI. GENERAL.A fully retractable, hydrau I My operatedmain landing gear and a nonret ractab le, fullpivotingtail wheel are provided. All componentsare of the single-leg, half-fork type and areequipped with air-oil -type shock struts. (SeeFigure 1.) Inasmuch as most repairs to thelanding gear are accomplished by part replacement,procedures outlined in this Sectiondeal mainly with the removal, repair, and reassemblyof the tires, tubes, and wheels.Broken or cracked components of the landinggear assembly, or bent or chafed axles andpivot pins must be replaced. Aii special toolsand materials mentioned are tabulated in theclosing paragraphs.AIR-OIL SHOCK STRUT AXLEFigureBRAKE PISTONBRAKE SHOE1 — Main Landing Gear Shock Strut2. MAIN LANDING GEAR.The main landing gear consists of two identicalleft-hand and right-hand componentswhich retract inboard into the wing centersectionjust forward of the main spar onwhich they are supported. (See Figure 2.)On each component is mounted a one-piece, castaluminum al loy, drop-center wheel assemblywhich is machined to permit installation ofthe steel drum I the bearing cups, andthe 27-inch tire. The wheel access fairingis of stamped aluminum al loy and is eas i lyremoved to inflate the tire and inspect thebearings. Small raised inflation seams ordeflection marks are provided on the peripheryof each sidewal I the tire as an aid toprope r i lat i on3. TAIL WHEEL.In likeness to the main landing gear, the1. STRUT FAIRING2. SHOCK STRUTFigure3. WHEEL JACKING POINT4. 27 "SMOOTH CONTOUR WHEEL2— Main Landing Geartail wheel assembly is also a single-leg,half-fork, fu I i lever structure. Thetail wheel support assembly consists of analuminum alloy casting and a chrome molybdenumsteel forging knuckle assembly mounted onroller bearings within the rear portion ofthe support. (See Figure g.j The assembly isattached to the fuselage aft section at threepoints. Two of these points are the maintrunnions at the forward end of the cast aluminumalloy support to which the fork andthe steering mechanism are mounted. The thirdpoint of attachment is located at the top ofthe pneudraulic shock strut. The axle andfork are of heat-treated chrome molybdenumsteel. The fork is so heat treated as tomake it the weakest item in the assembly andthereby prevent failure of more importantunits. Unlike the main landing gear, thetail wheel hub casting is made up of two partsas an aid to easy mounting of the tire andtube. (See Figure 4.) A 10-inch, smoothcontourtire is mounted on the tail wheel,the casing of which is also provided withsmall raised inflation seams or deflection[206]RESTRICTED

LANDING GEARPAR. 3 TO 5RESTRICTEDmarks on the periphery of each sidewall asan aid to proper inflation.^. REMOVAL OF WHEEL ASSEMBLY.To remove the main landing gear wheels,place an outer wing jacking stand under theouter wing jacking points and extend the jackuntil the wheel is free to rotate. The tankdoors must be installed before jacking theairplane at the outer wing jacking points.An alternate method of jacking is accomplishedby placing a small hydraulic jack in a mainlanding gear wheel jack cradle and placingthe center of the jack under each landinggear strut at the jacking point provided.(See Figure 5. J Extend the jack until thewheel is free to rotate. This latter methodshockstrut:trunnTonFigure 4 — TailWheel Componentstools. Inspect inside and outside of tirecasing for damages. Classify the damage asrepairable or not repairable, and treat accordingly.External defects of the casingor damage which does not affect the fabricimy be repaired. All damaged casings shallbe repaired except those having one or moreof the following defects: internal breaks,cuts or tread wear which exposes the fabriccarcass; diagonal or X breaks; injuries tothe bead or bead area, except in cases wherethe chaffer fabric is damaged or loose, providedthe damage is not caused by working ofthe beads; evidence of body ply separationor severe sidewall flexing; injuries throughat least 50 percent of the cord body andlarger than 2 inches; sidewall injuries requiringpatch repairs; two sidewall injuriesFigure 3 — Tail Wheel Assemblymust necessarily be used if the wings are notattached. Remove the dzus-fastened accessfairing plate. Remove the hub nut cap retainingspring and the hub nut cap. Extract the cotter keyand unscrew the large hex nut on the axle. (SeeFigure 6.) To remove the tail wheel, place a jackunder the tall jacking point located at station215-5/8 and extend the jack until the wheel isfree to rotate. Remove the screw-fastened wheelfairing plate; extract the cotter key and remove theretaining nut and washer from the axle. (SeeFigure 7.jTJ3II4JACK CRADLEHYDRAULICJACK5. REMOVAL AND INSPECTION OF TIRE AND TUBE.Remove the tire from the wheel with anystandard or improvised tire tools which arefree from sharp edges. Care must be takennot to injure tire, tube, or wheel with theseFigure 5 — Hydraulic Jack CradleRESTRICTED[207]

I weakenediesI reparsI cord1 beRESTRICTEDLANDING GEARPAR. 5 TO 8within 60 degrees of each other; and injuriesof any kind that cannot be repaired in asatisfactory manner. Worn casings retainingmore than 25 percent of their estimated servicelife and having no irreparable injuriesshall be treaded except when the casing hasbeen treaded twice or where the tread is unserviceablebecause of cactus spines. Inspectthe tube for the following defects: physicaldamage to the valve or faulty attachment tothe tube; tube body defects, or evidence ofwrinkles or creases; tube body thin spots,cuts, punctures, chafing, or damaged areasresulting from casing breaks. To expediterepair, mark all injuries as they are found,ii ndentat ons^ caused by foreign substances,when such indentations are smooth and shallow;spread cords when the space between cordsdoes not exceed the width of one cord andwhen the cord count over a 15-degree arc ofthe casing is not reduced more than 15 percent.6. WHEEL INSPECTION AND REPAIR.After the tire and tube have been removedfrom the wheel, inspect the rim for evidenceof corrosion, cracks resulting from carelessuse of tire tools or impact of any object,and worn places in the protective coating ofpaint on the rim. If the wheel becomes damagedto the point where excessive distortion isnoted or where cracks appear in the casting,no attempt should be made to repai r the wheelor rim. Portions of the wheel where theoriginal finish has become worn or has peeledoff should be retouched with lacquer. Oneor two coats should be applied, dependingupon the condition of the old paint. If itis necessary to ref inish the surface completely,one coat of primer fol lowed by two coatsof lacquer should be applied. It is essentialthat the surface be thoroughly cleaned beforethe application of the primer. The old finish.should be removed with acetone.7. NEW CASING REPAIR.New casings having defects of a minor oriTBJor nature may be repaired before use. Repairall defects of a new casing in the samerranner as set forth in the following paragraphsfor used casings. The follcwing defectsappearing in new casings shall be considerednegligible: lettering i 1 leg ib le on one s ideonly; portions of deflection markers missingprov'ided that at least 75 percent of therrarker is visible on each side of the casing;rounded edges of tread design due to lack ofpressure when curing, when the radius doesnot exceed one half of the nonskid depth;depressions in the tread or sidewalls causedby foreign material or air traps when thepresence of such depressions does not affectthe serviceability of the casing; insideFigure 6— Removing Main Landing Gear Wheel8. CASING REPAIRS.A I i to cas i ngs sha I made withstandard commercial type uncured or semicuredcord fabric. The number of cords and thequality shall be the same as those in theoriginal casing. If any of the repairabledefects or damages listed in a foregoing paragraphare present, repair as foiloA/s: Removea I or weathered cords and roughenall skived tread surfaces. Thoroughly cleanand buff a I surfaces to which new materialis to be applied and coat all exposed cordswith cement immediately after they are exposed.Repair injuries through the tread or sidewallonly and those smaller than 1/2-inch extendinginto not more than 25 percent of the cord bodyp I ing the ent i re hole with gum andby f i I Iapplying a small piece of sheeted gum stockover the hole on the inside cord ply. Reinforcewith an inside patch all injuries largerthan 1/2- inch or extending through more than 25percent of the cord body plies. Clean, cement, andreplace all chaffers damaged by tire tools inremoving the casing from the rim. Cure allrepairs by any suitable means. If originaltype molds are used, register the tread designon the casing with the tread design in themold. Regulate the time and temperature ofthe cure so that it will be appropriate forthe rubber stock being cured. Tread cutswhich do not penetrate the fabric are to beI 208]RESTRICTED

lowI damagedLANDING GEARPAR. 8 TO 12RESTRICTEDcleaned, filled with commercial tire cutfiller, and cemented in place. If sidewallblisters are present that can be cleaned andrepaired with no resultant damage to the fabriccarcass, or if the rubber fairing immediatelyabove the rim flange separates from the fabriccarcass with no apparent injury to the carcass,the damaged areas may be repaired bycleaning with wash thinner or gasoline, buffingthe areas to be cemented, and cementingseparated areas with airship rubber cement.A I cement to become tacky, and press rubbersidewall firmly against the fabric carcass.Some means should be provided to maintaina constant pressure on the repaired areauntil cement has become thoroughly dry.surface with vulcanizing cement and apply thenew stock* Care must be taken when rollingdown the new stock so as to exclude all airfrom between the new stock and the originalcarcass. Cure the treaded casing in fullcircle or factory-type molds, using appropriatetime and temperature for the rubber stockbeing cured. Clean and paint the treadedcasing after removing it from the mold.10. MARKINGS.All repaired or treaded casings shall beproperly marked upon completion of repair.Identify a repaired casing by painting orrubber stamping an arrow and the letter Ron the repaired area. Use suitable red oryellow paint or enamel for this purpose. Burnthe brand REP on each repaired casing immediatelyfollowing the original serial number.Apply the above brand each time a casing is repaired,in addition to all former brands.Burn the brand TR on each treaded casing immediatelyfollowing the original serial number.If the original serial number was removedin the treading operation, replace it by someapproved method. Apply the TR brand each timea casing is treaded, in addition to all formerbrands. All casings repaired or treadedby other than the original manufacturer musthave the contractor's name or trademark burnedupon them.11. TUBE REPAIR.9. TREADING.Figure 7 — Removing Tail WheelAll treading materials and workmanshipshall be in accordance with high-grade commercialpractice covering this type of repairwork. Repair a I casings except thosehaving one or more of the defects listed ina following paragraph. Normally, the treadingoperation shall be confined to the entire portionof the wearing surface of the casing that isbetween and including the inflation deflectionmarkers. Replace the deflection markers. Whentreading a casing having the tread or breakerplies separated from the carcass or havingdeep cuts in the tread, include as much of thesidewall as is necessary to produce a serviceablecasing. Remove the old tread and portionsof the sidewall where necessary and thoroughlyroughen the working surface. Cover the workingThin spots, cuts, and punctures may be repairedby patching or vulcanizing by anyapproved method. The patch size and the numberof patches should be limited, as the weightof the patch or patches may cause tube unbalance.If large holes or extensive thin orchafed areas are present, the tube must be discarded.12. TIRE AND TUBE BALANCE.In order to prevent dynamic unbalance of thelanding gear tires at high speeds, it is necessaryto balance the tires and tubes statically.The maximum static unbalance for a 27-inchsmooth-contour tire as set forth in the AAFSpec. AN-C-5 is 12 inch-ounces. Inasmuch asattaching balance weights on the outsideperiphery of the tire is difficult, balancingmethods set forth in the following paragraphswill designate all weights at the rim line.Twelve inch-ounces on the outside peripheryof the casing is equal to approximately 2-1/2RESTRICTED[209]

I seekRESTRICTED LANDING GEARPAR. 12 TO mounces at the rim line. Hence, all main landinggear tires should not exceed a maximum staticunbalance point of 2-1/2 ounces at the rimline. The tail wheel need not be balanced.13. TUBE BALANCING.In order to balance statically a patchedtube, a balance wheel is needed. (See Figure8.) One may be fabricated from wood or froma discarded wheel. The balance wheel must bestatical ly balanced before any attempt ismade to balance the tire or tube. Inflatethe tube until the true shape is reached butno stretching occurs. Mount the tube on thebalance wheel and allow the wheel to rotate.The heavy portion of the tube wi I thelowest point on the wheel. Mark this point,rotate the wheel 90 degrees, and add a balanceweight at the rim line. Experiment with differentweight values until the accurate amountof weight required to balance the tube isfound. If this weight exceeds 2-1/2 ounces,patches may be added 180 degrees from theheavy spot for a counterbalance. Apply allbalance patches with the tube fully inflated.When the weight required is less that thatmentioned, mark the light area of the tubeto facilitate wheel balancing upon insertionof the tube into the casing.m. TIRE BALANCING.fvbunt tire on balancing wheel and determinethe balance weight as directed for tube balancingin the preceding paragraph. (See Figure H.)As soon as the correct amount of balance weightis found, rmrk the heavy section. If the staticunbalance exceeds 2- 1/2 ounces at the rim line, removethe tire from the balance wheel and spreadthe beading apart with wooden blocks directlyopposite the heavy section. The inside areaof the casing to which balancing dough is tobe applied is indicated by a dotted line. (SeeFigure H.) Clean the inside of the casingfor an area of at least 1-1/2 square feet,using a wi re brush and a rag slightly moistenedwith wash thinner. Apply one coat of Goodyearrubber cement No. 865 with a 2- inch stiffbrush. (See Figure g.) Allow to dry forADD BALANCING DOUGHOPPOSITE HEAVY MARK>2" STIFF- BRISTLEDBRUSH-HEAVYSECTIONMARKFigure S^Balance WheelFigure 9—Applying Cement[210]RESTRICTED

I atLANDING GEARPAR. m TO 15RESTRICTEDCEMENTFigure 10— Applying Balancing DoughFigure 11 -'Applying Tire TalcBalancing Doughto30 minutes and apply Goodyear No. 865C balancingdough cement with a metal spreader. (SeeFigure 10.) A I Ida/ to dry for 30 minutes. Theamount of dough applied will depend upon theweight required to balance the tire. Recheckthe tire balance on the balance wheel; addor remove balancing dough until balancingweight is within the minimum of 2-1/2 ounces.After the balancing dough is dry, apply powderedtire talc to the balancing dough. (See Figure11.) This prevents the tube from stickingto the balancing dough. The tire is now readyfor i nsta I i on.15. ASSEMBLING TIRE AND TUBE ON WHEEL.Lay the wheel on a flat smooth surface withthe drop-center rim portion toward the top.Utilizing standard tire tools and a rubbermallet, mount one bead of the casing on therim. (See Figure i6.) Before inserting thetube, dust the entire tube surface with pcMderedtire talc. This may be best accomplished byplacing a quantity of tire talc in a shallowbox, laying the tube on the bottom, and applyingthe tire talc with a cheesecloth. (See Figurei6, Detail B.) The tube rmy also be dusted whenfully inflated if a dusting box is not available.(See Figure 12.) Tie a 2-foot cord securelyonto the valve stem, or screw on a valve extension.(See Figure 29. j Insert the tube intothe casing, carefully aligning the rmnufacturer'stire and tube heavy and light section marks.These appear in the form of a red dot or mark onthe side of the tube and the casing. (SeeFigure 14.) Al ignment of these two red dotsis essential to whe^^l balance upon installationof the tire and the tube on the wheel. Placethe valve stem opposite the valve hole. Threadthe valve stem string or extension through thevalve stem hole in the rim, screw on the valvelocknut a few turns, and add just enough airto shape the tube properly. If too much airis added, difficulty will be encountered whenmounting the remaining bead of the casing.Mount the second bead of the casing, startingat the valve. The use of tire-holding clipswill expedite the mounting of the second beadRESTRICTEDL211]

,RESTRICTED LANDING GEARPAR. 15 TO 1616. MOUNTING TAIL WHEEL TIRE AND TUBE.CHEESE CLOTHInasmuch as the hub of the tail wheel is ofa two-piece construction, the mounting of thetire and tube of the tail wheel is not similarto that of the main gear. Apply tire talc tothe tube by placing the tube in a dusting boxand applying tire talc with a cheesec loth, f^egFigure i8.) Insert the tube into the casingCASINGMARKHEAVY SECTIONFigure 12-^Applying Tire Talcto Inflated Tubeof the casing. (See Figure i^.) After the casingis properly mounted, add air slowly until thecasing beads seat. If the beads of the casingdo not seat properly, deflate the tube, realignthe casing, and re inflate. Remove the valvecore, and fully deflate the tube to relievethe pressure on any folds or buckles and topermit the tube to assume its proper contourwithin the casing. Be sure both casing beadsTUBE LIGHT SECTIONMARKFigure 14— Tire and Tube Markingsand partially inflate. (See Figure ij, DetailsA and B.) Remove the demountable rim portionof the hub and slip the tireand tubeover the remaining portion. (See Figure lyDetails C and D.) Thread the valve stemthrough the valve stem hole and place the demountableportion of the hub in position. Placethe retaining spring in position and inflatethe tube. (See Figure ij , Details E and F.)T J329i-2TIREINSTALLATIONCLIPSFigure 13— Valve Extensionremain properly seated on the rim. Replacethe valve core, inflate the tire to the requiredpressure, and tighten the valve locknut.Inflate the tire until the deflection markson the side of the tire just contact the groundline when the airplane is in a three-pointposition on hard level ground. Under normalconditions, this inf lat ion wi I I be equivalentto approximately 30 Ibs./sq.in.Figure 15-^pecial Tire Tools[212]RESTRICTED

LANDING GEARFIGURE 16RESTRICTEDFigure 16— Mounting Main Gear Tire and TubeRESTRICTED[213]

RESTRICTED LANDING GEARFIGURE 17EIEFigure 17— Mounting Tail Viheel Tire and Tube[214]RESTRICTED

MER,INCHESALLANDING GEARPAR. 17 TO 18RESTRICTEDTOOLSCRADLE - JACKSTAND - WING JACK! NGSTAND - TAIL JACKINGWRENCH - MAIN AND AUXILIARYLANDING GEAR HUB NUTTOOLS, TIREMALLETS, RUBBEREXTENSION, VALVE,FLEXIBLEJACK, SMALL, HYDRAULICPARTPARTPARTREMARKSNO.NO.NO.TJ 3114TJ 4612TJ 2164IPART NO. TJ 44291-1/2 BY IfINCHES, OR LARGERLARGE OR MEDIUMAT LEAST 6 I NCHESLONGUSE UNDER WHEELJACK ING POI NT18. MATERIALS.Figure 18—Apply ing Tire Talcto Tail Wheel TubeThe following materials are needed for repairsoutlined in this Section:MATERIALSREMARKS17. TOOLS.Jacking stands and special tools are obtainableto expedite wheel jacking, removal, andreassembly. All special tools listed are manufacturedby North American Aviation, Inc., foruse with this series of airplanes. To remove,repair, and reinstall landing gearwheels andtires properly, all or part of the followingtool s are requ i red:CEMENT, A! RSH t P RU BBERTALC, POWDERED, TIRELACQUER, CELLULOSEN TRATEPR I METAL Z I NCCHROMATEACETONEFILLER, RUBBER T I RECUTSPEC. 20-37CLASS 04-BSPEC. 4-33CLASS 29SPEC. 3-158SPEC. 14080SPEC 0-A-51ACOMMERC ! TYPE,PROCURE LOCALLYRESTRICTED[215]

controlandRESTRICTEDCONTROL CABLESPAR. I TO 7SECTION 7CONTROLCABLESI. GENERALIThe control cables are fabricated mostly ofextra-flexible preformed corrosion-resistantsteel; however, tinned steel cable is also usedinterchangeably. Cables vary from 1/16 to 5/32-inch in diameter. Cables of 1/8-inch and largerare composed of seven strands of 19 wires each.Cables 1/16 and 3/32- inch in diameter are composedof seven strands of seven wires each.The main control cable terminals are of thedie-swaged friction type. The remaining cablesare either sweat-soldered into tinned terminalsor woven spliced. For information pertinentto the type cables and cable terminals usedin original fabrication see Figures and 2.It is to be noted that these figures list onlythe terminals directly swaged, soldered, orspliced to the cable. Miscellaneous fittingssubsequently screwed to the end terminals arenot I isted.in accordance with the data contained in FiguresI 2. However, if facilities and suppliesare limited and ininediate replacement isimperative, replacements sometimes may beprepared in the customary rmnner, using thimbles,bushings, and turnbuckles in place of originalterminals. When this is done, cables havinga diameter of 3/32- inch or over may be wovenspliced by means of the five-tuck method, andcables less than 5/32- inch in diameter may bewrap-soldered. The following chart specifiesthe particular terminals recommended for usewith an improvised woven spl iced and wrap-solderedcable only.CABLE2. NEGLIGIBLE DAMAGEDuring overhaul periods, or more frequentlyif circumstances permit, test the control cablesfor broken wires by passing a cloth alongthe length of the cable. Broken wires willbe indicated where the cloth is snagged. Brokenwires should not De permitted in the cablesof the trim tab control system^ nor of the rraincontrol system where such breakages occur inthat length of a cable norrmlly passing overa pul ley or through a fa i riead . A max imum ofseven broken wires along a straight uninterruptedlength rmy be permitted without replacementof the cable. Broken wires, when found,should be cut off short and served or solderedto the cable. It is recommended that recordsdesignating tne status of the damaged cablesbe maintained as an aid in determining the futuredisposition should additional damage occur.3. CABLE REPLACEMENTWherever possible,duplicate spare cables shouldbe utilized for replacements. See Figures Iand 2 for the part numoers of al I cables.^, CABLE FABRICATIONIf spare cables are not available, exact duplicatesof damaged cables should be prepared

CONTROL CABLESFIGURE IRESTRICTEDWr'ii'O-kiIwB%^'mmvL3C212n27InwF-^iSOLDER GABLETWO (2) PLACESINSIDE DRUM BE-FORE SPLICINGfi^ ^END-8L"TTTTIXALL FITTINGS ARE SWAGEDEXCEPT AS NOTEDFigure 1 — Typesof CablesHIA WRAP SOLDERED SPLICE4> SOLDERED FITTINGRESTRICTED[217]

I inch,ightubricatItoIRESTRICTED CONTROL CABLESPAR. 7 TO 8CABLELOCATIONNO.REQD.N. A. DWG.NO.TYPEFITTINGACABLEDIA.FITTINGBLENGTHIN INCHESADDITIONALFITTINGSAS NOTEDPARKING BRAKEREAR BRAKE - REAR RUDDER PEDAL TOMASTER BRAKE CYLINDERFRONT BRAKE - FRONT RUDDER PEDAL TOrtASTER BRAKE CYLINDERLANDING GEAR POSITION INDICATORLANDING GEAR POSITION INDICATORTAIL WHEEL CONTROLMANUAL STARTER ENGAGING CONTROLELEVATOR FRONT LOWERELEVATOR FRONT UPPERELEVATOR REARAILERON CONTROL FRONT - WING C. S.AILERON CONTROL REAR - WING C. S.AILERON CONTROL FRONT - OUTER WINGAILERON CONTROL REAR OUTER WING-RUDDER CONTROL FRONT LHRUDDER CONTROL FRONT RHRUDDER CONTROL REAR LHRUDDER CONTROL REAR RHRUDDER - BALANCEELEVATOR TRIM REARELEVATOR TRIM FRONTRUDDER TRIM - FRONT277-33t5066-33t7666-331*7736-33562-36-33562-66-3H0I277-4501166-5221766-5221877-5221966-5231065-5231166-5231266-5231366-5215977-5215966-5216077-5216066-5216166-5252766-5252866-52537IVIVVIIIIVIVIVIIIIIVIVIVVVIVIXIIXIXNAF3I062IB-2NAF3I062IB-5NAF3I062IB-5AC066I35AC066I35NAF3I062IB-577-15039NAF3I062IB-1NAF3I062IB-1NAF3I062IB-1BI277-1B 1 277-1NAF3I062IB-1NAF3I062IB-1NAF3I062IB-1NAF3I062IB-1NAF3I062IB-1NAF3I062IB-1NAF3I062IB-3BI281-LNAF3I062IC-2BI281-R1/ 165/325/321/ 161/165/321/161/81/81/81/81/81/81/81/81/81/81/83/321/161/161/16AC066I35NAF3I062I-S-32D-5RHNAF3I062IB-525-3352625-33526NAF3I062I-S32D-5LH77-15012NAF3I062I-S2ID-1RHNAF3I062I-S2ID-1RHNAF31062I-S2ID-1LHNAF3I062I-S2I0-1RHNAF3I062I-S2ID-1RHNAF3I062I-S2ID-1LHNAF3I062I-S2ID-1LHNAF3I062IB-1NAF3I062IB-1NAF3I062I-S2ID-1LHNAF3I062I-S2ID-1LHNAF3I062I-SI6D-3LHAN 1 1 1-3NAF31062IC-2Bi281-R25-15/1572-11/3212-5/85388-3/833-5/3225-1/16215-7/32173-11/3215-19/3292-9/1590-5/3261-21/3269-15/32165-1/8165-5/1663-27/3263-5/32111-9/ 16231-1/8306-15/16315-3/823- 52153(C)23-52161(0)19-52510 C)19-52505(0)RUDDER TRIM - REARFLAP POSITION INDICATORINSTRUMENT FLYING HOOD RELEASELATCH - FRONTINSTRUMENT FLYING HOOD RELEASELATCH - REAR66-5253855-5262958-7301259-73031XIIIVVIIIBI281-LAC066I3555-7301555-730151/161/161/161/16ANIII-325-3352625-3352625-33526191-5/877-7/1618-3/185-3/8I9-525I0(C)19-52505 (D)52-73010(055-73011(0)52-73016 E)55-73005 (F)52-73010(055-73011(0)52-730 16(E)59-73005 (F)58-73035(6)Figure 2—Cable DataWipe off all excess oil. It is to be noteathat corrosion-resistant steel cable does notrequire this treatment for rust prevention.8. SV^AGED TERMINALScable ends,prior to the second swaging operation.Swaged terminals should be utilized in fabricationonly if they were original ly employed.The specific term-inals noted on Figures I and2 must be used in their original function. Afterpreparing the necessary caole length with allowancemade for the fitting elongation underswaging and proof loading, coat the end of thecable with No. 10 SAE Iing oi I . Insertthe cable into the terminal approximatelyDend the cable toward the terminal,straighten the cable back to normal position,then push the cable end entirely into the terminal.Apply a drop of I oi Ithe terminaland insert the terminal about 1/4- inch intothe die when the die is moving at a slew speed.Then insert the fitting entirely into the dieand adjust the die for regular speed (see Figure3J . If swaged terminals are used on both Figure 3— Cable Terminal Swaging[216]RESTRICTED

CONTROL CABLESFIGURE 4RESTRICTEDBUSHING (REF.)4 6 7FREE ENDSERVING CORDWRAP TIGHTLYAND VARNISHFigure 4-^Preparation of aWoven Cable SpliceRESTRICTED[219]

amount.lowlewandnargateRESTRICTED CONTROL CABLESPAR. 8 TO 10accurately measure the over-all caDle length.If, after swaging, the terminals have more thanthe allowable l/2-clegree bend, secure them ina vise and straighten with as few appi icationsof pressure as possible.9. SWEAT-SOLDERED TERMINALSSweat-soldered terminals should be employedin fabrication only where they were originallyized. See Figures 2 for specific typesut i I Iof t inned term! na Is ut i I ized in or igi I fabrication.Stearic acid or a suitable mixtureof stearic acid and resin or resin dissolvedin alcohol may be used as the soldering flux.f'Ajriatic acitl should not be used as a flux. Thetenninals should be prepared as follavs: Applythe soldering flux to the end of the cable andinsert the end of the cable into the barrel ofthe terminal, allowing the cable to extend throughthe barrel a smal I Free the end strandsof the cable and a I them to fray. Pul I thecable back into the barrel until it is flush,then thoroughly sweat solder into the cableand barrel until solder appears at both endsof the barrel. Avoid overheating the solder;Sweat-soldered terminals can be eas i l-y distinguishedfrom swaged type terminals by the airholes provided in the barrel of the terminal.These air holes a I the molten solder to perrreatethe strands of the cable with no a i bubblesattendant. (See Figure ^.)10. WOVEN SPLICED TERMINALSThe five-tuck woven spl iced terminals may beutilized on cables of 3/32-inch diameter orgreater in place of swaged terminals where facilitiesare limited and immediate replacementis imperative. In some cases it will be necessaryto splice one end of the cable on assembly.For this reason, invest i the or ig ina Iinstallation for pulleys and fairleads thatmight restrict the passage of the splice. Theprocedure for the fabrication of a woven spliceis as fol lows: See Figure 4 for the designationof numbers and letters referred to in thissequence of operations, and see Paragraph 4for fittings required.NAF 3I062IC(SWAGED TYPE)NAF 3I062IB(SWAGED TYPE)miuumiuuiH"•tmmmmr'Figure 6 — Cable Clamp for Woven SpliceNAF 3I062I-S2ID (SWAGED TYPE)1. Secure the cable around a bushing or thimbleby means of a splicing clamp, leaving 8inches or more of free end. Secure the splicingclamp in a vise with the free end to theleft of the standing wire and away from theAG066I35 (SOLDERED TYPE)operator. (See Figure 6.) If a thimble isused as the end fitting, turn the points outwardapproximately 45 degrees.juimimmDNA25-33526 ( S LD E R E D TYPE )2. Select the free strand (I) nearest tne standinglength at the end of the fitting and freethis strand from the rest of the free ends.Figure 5— Types of Cable Terminals Next, insert a marlinspike under the first three[ 220]RESTRICTED

CONTROL CABLESPAR. 10 TO IIRESTRICTEDstrands (A, B, C) of the standing length nearestthe separated strand of thQ free end and separatethem momentarily Dy twisting the marlinspike.Irvsert the free strar^d ( I) under thethree separated strands through the openingcreated Dy the marlinspike. Pull the free endtaut by means of pliers.3. Unlay a second strand (2) located to theleft of the first strand tucked, and insertthis second strand under the first two standingstrands (A,B). Loosen the third free lengthstrand (3) located to the left of the firsttwo, and insert it under the first standingstrand (A) of the original three (Detail A).4. Remove the center or core strand (7) fromthe free end, and insert it under the same standingstrands (A,B). Temporarily secure the corestrand to the oody of the standing caole (DetailB). Loosen the last free strand (6) locatedjust to the right cf the first (I) andtuck in under the last iwo strands (E,F) ofthe standing caole. Tuck the fifth free end(5) around the fifth standing strand (E). Tuckthe fourth free end (4) around the sixth standingstrand (F) (Details B,E) . Pull all strandssnug toward the end fitting with the pliers.This completes the first tuck.5. Begin with the first free sirand (I) andwork in a counterclockwise direction, tuckingfree strands under e\/ery other standing strand.After the completion of e'^ery tuck, pull thestrands tight with pliers. Pull toward theend fitting (Detail C) . After the completionof the third complete tuck, cut in half thenumoer of wires in each free strand. Uake anothercomplete tuck with the wires remaining.At the completion of the fourth tuck, againhalve the number of wires in the free strandsand make one final tuck with the wires remaining.Cut off all protruding strands and poundthe splice with a wooden or rawhide mallet torelieve the strains in the wires. Serve thesplice with waxed linen cord. Start l/4-inehfrom the end of the spl ice and carry the wrappingover the loose end of the cord and along thetapered splice to a point between the secondand third tucks. Insert the end of the cordback through the last five wrappings and pullsnug. Cut off the end; and if a thimble isused as an end fitting, bend down the points.Apply two poats of she 1 lac to the cord, a I lowingtwo hours between coats (Detail D) .Carefully inspect the cable strands and splicesfor local failure. Weakness in a woven spliceis made evident oy a separation of the strandsof serving cord.II.WRAP-SOLDERED TERMINALSOn cables of 1/16-inch diameter only, the wrapsolderedsplice may be employed to fabricateend fittings. See Figure 7 and Paragraph 4for method and materials. Stearic acid ora suitable mixture of stearic acid and resinor resin dissolved in alcohol may be used asthe soldering flux, fvuriatic acid should notbe used as a flux. Arrange the cable and thefittings as required, allowing approximately2-1/4 inches of free end. Place the assemblyin a splicing clamp and secure in a vise. Startingas close as practicable to the end fitting,press the free end standing lengths of the cabletogether tightly and wrap with a singlelayer of No. 20 brass or copper wire, leavinga space of approximately 1/8-inch between es/ery1/2- inch of wrapping. Care must be exercisedto prevent the standing length from twistingduring this operation. Al low the wrapping toextend approximately 1/4-inch beyond the freeend. Dip the wrapping in tin-lead solder andcarefully sweat the solder into the cable andabout the wrapping. Apply the solder untilthe wrapping wire is barely discernible andmake certain that the open spaces between thewrap sections are thoroughly impregnated withL-A STEP 4SOLDERAATitle 7— V/rapSoldered SpliceRESTRICTED[221 ]

:RESIRICTEDCONTROL CABLESPAR. II TO msolder. After the splice has cooled, thoroughlywipe clean and positively remove all solderingflux Dy washing in hot water. Wipe the cableand impregnate the spliced section with CorelNo. 95 or equivalent. Carefully inspect thesplice. A wrap-soldered splice, easily oentwith the fingers, is unsatisfactory because of alack of solder penetration. Cracks in thesolder located between the wrapping wire andthe short space provided between wraps is apositive indication of slippage in the wrapsolderedspl ice.12. TESTINGAll cables and splices should be tested forproper strength prior to installation. Arrangecables to simulate installation, includingpulleys where required, and gradually applya load to one end of the cable for a periodof 3 minutes. A suitable guard shouldbe placed over the cable while it is being tested,to prevent personal injury in the eventof cable failure. Apply test load to cablesas fo I I owsCable Size1/163/321/85/3213. COLOR BANDINGLoad in Pounds300550I 1501550All cables fabricated locally should oe colorcoded in accordance with the instruction giveni n Sect i onm. TOOLSIIn connection with the preparation of fabricatedcables, the following tools may oe requinsp i ke ofred: ( ) a sma II poi nted mar I ioval cross section, (2) a pair of side cuttingpliers, (3) a splicing clamp (see Figure6)^ (4) a rawhide or hard wood mallet, (5) adie-swaging machine. (See Figure 3.J[ 222 ]RESTRICTED

I ikeimiIfRESTRICTEDELECTRICALPAR. I TO nSECTION 8ELECTRICAL SYSTEMi. GENERAL.Refer to NA CWgs. 77-54001, 84-54001, and 88-54001 for general installation of electricalequipment and the various main assembly componentpart numbers for the AT-6A, AT-6B, andAT-6C Airplanes, respectively. For radio equipmentinstallation, refer to NA Dwg. 77-71001for the.AT-6A and AT-6B Airplanes and to NA Dwg.88-71001 for the AT-6C Airplane. For wire routing,length, and size and for terminal typesand sizes, refer to NA O/gs. 77-54002, 84-54002,and 88-54002 for their respective models. Radiowiring charts are separate from the electricalwiring diagrams and will be found on NA Dwg.77-71002 for the AT-6A and AT-6B, and on NA Dwg.88-71002 for the AT-6C Airplane. Electrical andrad io wi ring d iagrams for the part icular modelare furnished with each airplane and are storedin the data case in the cockpit. Concludinginstallation, ascertain that interiors of junctionand switch boxes are free from filings,cuttings, bits of solder, pieces of tape, andpieces of wire, the presence of which is causefor shorting of circuits and the lowering ofthe dielectric strength of the insulation. Ascertainthat terminals, terminal nuts, and conduitfittings and clamps are secure.factory fnethod of lacing a wi re group is accomplishedby the use of the marl in, or hammock,hitch. This hitch is the reverse of a commonhalf hitch. The cord passes over-and-under theloop instead of under-and-over as it does ina ha I hitch. (See Figure 6, Item ^.)3. WIRE NUMBERING.When reterminal ing or replacing a wire, caremust be taken to renumber the wi re properly.If no identification number exists on the wirebefore repair, or if the number is i legible,NUMBEREDIftfCELLULOSE TAPEil^j^ Mill2. WIRING.Figure 1 — Typical VfireRoutingAil wiring should be installed in a workmanrmnner.Wiring within conduit should beinstalled neatly and every effort made to preventcrossing of wires inside rigid or flexibleconduit. Wiring must be of gage specifiedon the electrical wiring diagram and shouldbe free of kinks, frays, and ruptures. Installationof wi ring shou Id be such that a Itension on wi res is e I nated. Wires withinjunction and switch boxes and on all items ofelectrical equipment should be routed to providesufficient wire length to facilitate reterminaling in the event such becomes necessary.(See Figure i.J Cord lacing and spottying of wires and wire groups are requiredwithin all switch and junction boxes and onswitch box covers. See Figure 6, Items 4 and5, for typical examples. Where a number ofwires enter a box, individual groups leadingfrom each conduit should be laced separatelyto facilitate wire renewal. The most satis-ascertain the correct number by careful ly checkingthe electrical wiring diagram. Originalidentification numbers are cut from MinnesotaMining Company numbered cellulose tape. (SeeFigure 1.) However, if this tape is not available,a facsimile may be made by the use ofwhite paper and clear eel lulose tape. Wrap aminimum of 1-1/2 turns of tape around the wireand apply one coat of clear lacquer over thetape.4. WIRE REPLACEMENT.iDamaged wires should be replaced with newwires. Before removing the damaged wires orwire group, tie a strong waxed cord to oneor more of the wires to be removed. By pul I-ng the wires out the opposite end, the cordis threaded through the conduit. After thedamaged wires are pulled out of the conduit,the cord may be utilized to pull the new wires[ 223 ]RESTRICTED

e,IipJELECTRICALPAR. i^ TO 6RESTRICTEDSIRCO OR IRVOLITETUBULAR INSULATIONI— TRIMAS SHOWNCONNECTORFigure 2— Slipping Insulation Over Wireback through the conduit and into position.If a conduit repair is to be made, this methodcannot be used. However, after the conduitrepair is completed, a stiff res i I lent wi re,s uch as p iano w i may be ut i I i zed to pu Ithe wires or waxed cord through the conduit.If this method fai Is or if wi re is not aval I-able, a string may be blown through the conduit.Knot the string and place the knot afew inches within the conduit. Place a compressedair nozzle at the opening and feed thestring into the conduit whi le al loving the compressedair to blow the string through the conduitand out the opposite end.5. WIRE TERMINALS.Two types of terminals are employed on theelectrical wiring, the soldered and the crimpedtype. The terminals used in repair work mustbe of the size and type specified on the electricalwiring diagram. (See Figure J.) Thetypes of terminals illustrated are used onthe AT-6A, AT-6B, and AT-6C Airplanes. Whenremoving insulation prior to terminal ing, caremust be exercised not to nick individual strandsof the conductor. In case a soldered terminalis to be used, avoid overheating the wire andFigure 3— Placing Terminal Over Bared Wireapplying excessive solder. Use a tin-lead solderand a resin base flux. Acid base flux lowersthe dielectric strength of the insulation materials.Shoulders of the terminals should betaped and coated with clear lacquer to preventshort-circuiting by contact with adjacent terminals.6. ATTACHING CRIMPED TERMINALS.To attach crimped-type terminals or connectorsto the electrical wiring, proceed as illustratedin this Section. (See Figures 2, 9, 4, and 5.Trim the insulation from the wire for a length ofI rvoliteI the1/16-inch plus the length of the terminal shoulder.Slip the appropriate length of Sirco ortubular insulation over the wire beforeattaching the terminal. Slip the terminal orconnector shoulder over the bared wire end asshovn. (See Figure 9. J tvfeke certain that a Iwire strands are inside the tubular shoulder.Using the special Sta-Kon crimping pliers, crimpthe shoulder of the terminal or connector ands I the Si rco or Irvolite insulation down overthe shoulder of the terminal as shewn. (See Figures4 and 5. j If neither of these insulators isavailable, tape the terminal with friction tape andapply a coat of lacquer. Sta-Kon terminal crimpingRESTRICTED[224]

I otherRESTRICTEDELECTRICALPAR. 6 TO 8pliers are available in three sizes: size No.WT-M2 for use on Types A and B terminals; sizeNo. WT-I 13 for use on Type C terminals; sizeNo. WT-I 14 for use on Types D, E, and F terminals.7. CONDUIT.For d imens ions, part numbers, and routing ofall rigid and flexible conduit, refer to NADwgs. 77-'^400l and 77-71001 for the AT-6A Airplane,NA [>gs. 84-54001 and 77-71001 for theAT-6B Airplane, and NA CWgs. 88-54001 and 88-71001 for the AT-6C Airplane. Sizes specifiedon these drawings must be adhered to. Conduitshould be supported by means of suitable clampsor clips, as specified on the respective drawingfor that model airplane. Under no circumstancesshould cord lacing be utilized as ameans of conduit support or attachment. Conduitshould be free from dents, scratches, and/orother blemishes. Care should be exercised wheninstalling conduit to ascertain that conduitenters box fitting or bulkhead union properlyand not at an angle. Do not attempt straighteningof conduit by tightening fitting nuts,as this action will dent or bend the conduitif the conduit is of a small diameter. Ifthe conduit is thick-walled and of a larged iameter, the s ide of the j unct ion box wl I Ibe pulled out of shape. Concluding installa-tion, ascertain that conduit fittings are secureand that all conduit has clearance atall frame cutouts and at a I sharp orprotruding members. Conduit must be bondedby means of removing the finish at all unions,box fittings, and under all supporting clamps.^4ethods of splicing conduit which has beendamaged are outlined under tubing repairs inSect ion IX.8. BONDING.Electrically isolated metal parts and shieldingshould be bonded by interconnecting theseparts to the airplane structure. Connectionsfor bonding must be of good electrical value,offering resistance not to exceed .001 ohm ateach connection. Bonding should be accomplishedby soldering, bolting, clamping, or weldingas required. Metal -to-metal joints pemianentlyjoined by welding, soldering, sweating, or rivetingrequire no further bonding. No bonding isrequired for small isolated parts such as screwsand small coverplates, or for control cables, insulatedlines or members less than 60 incheslong, or permanently isolated members subjectto unlimited travel. With the exception ofthose located forward of the firewall, flexibleand rigid conduit of the AT-6C Is bonded bythe clamps which support it, and on the AT-6ACRIMPEDCONNECTOR-PLIERSFigure 4— Cr imping the Terminal Figure 5 — Slipping Insula tion Into Position[225]RESTRICTED

ELECTRICALFIGURE 6RESTRICTEDSWITCHBOX COVER1. AN 660 TERMINALS - LIGHT TYPE2. AN 660 TERMINALS — ROLLED TYPE3. AN 660 TERMINALS— HEAVY TYPE4. SPOT TIES5. WIRE GROUP LACINGFigure 6 — Typical Switch Box WiringRESTRICTED[225 J

RESTRICTEDELECTRICALPAR. 8 TO 9and AT-6B by tinned copper flexible bondingbraids. Type BI209. (See Figure 8.) In orderingthese bonds, the first dash number isthe terminal size on one end; the second dashnumber is the terminal size on the oppositeend; the third dash number is the width ofthe bond in 1/8-inch; and the fourth and fifthdash numbers designate the length of the braidmeasured from terminal hole to terminal hole.Example: A B 1209-4-8-1 -3-'3 bonding braidhas a terminal on one end for a No. 4 screwand a terminal on the other end for a No. 8screw. The bonding braid is 1/8-inch wide andWELDED BI209 BONDINGSHERMAN-21LIGHT AMPERAGETSB STA-KONTYPE A86FLEXIBLECONDUIT BONDINGaaHH BfiAC 660-50LIGHT TYPEAC 660-7HEAVY TYPETaB STA-KONTYPE B87'^SPP'^ v«.iil -*iiwT a B STA - KONTYPE C76Figure 8 — Conduit Bonding Methods9. ELECTRICAL REPAIR TOOLS AND MATERIALS.For electrical repairs, all or part of thefollowing tools and materials are required:TOOL OR MATERIALAC 660-60ROLLED TYPEFigure 7 — TypesTaB STA-KONTYPE D7Iof Termina Isi I lust3-1/2 inches long. Attach the bonding braidto copper flexible conduit by means of sweatsoldering and to rigid conduit by means of thewelded bonding lugs provided. Attach the bondingbra id to the rigid condu it only asrated with the terminal under the nut to serveas a shakeproof lock washer. (See Figure 8.)IRON, SOLDERING, 100 OR 200 WATTPLIERS, CRIMPING, STA-KON, SIZE TO SUITTERMINALS, SOLDE RED-TY PE , AC660, SIZE TO SUITTERMINALS, STA-KON, MADE BY THOMAS i BETTS CO.,ELIZABETH, NEW JERSEY, SIZE TO SUITTAPE, FRICTION, 1/2-INCH WIDE, PROCURE LOCALLYCORD, LACING, STRONG WAXED, PROCURE LOCALLYWIRE, SOFT SOLDER, 50/50 TIN-LEADFLUX, SOLDERING, RESIN BASE, PROCURE LOCALLYTUBING, SIRCO INSULATION, MFG. BY SUPRENANTELECTRIC INSULATION CO., BOSTON, MASS.TUBING, IRVOLITE INSULATION, MFG. BY IRVINGTONVARNISH & INSULATOR CO., IRVINGTON, N.J.LACQUER, CLEAR, PROCURE LOCALLYTAPE, CELLULOSE, CLEAR OR NUMBERED, MFG. BYMINNESOTA MINING & MFG. CO., ST. PAUL, MINN.BRAID, BONDING, B1209, LENGTH & TERMINALS TO SUIT[227]RESTRICTED

ELECTRICALFIG. 9 & 10RESTRICTED%% S^^Uf'^'^1. WIRE STRIPPERS2. DIAGONAL CUHERS3. LONG NOSE PLIERSi^. CRIMPING PLIERS5. FORMING PLIERS7. MIDGET SOCKET WRENCH SET8. PHILLIPS HEAD SCREWDRIVER9. STANDARD SCREWDRIVER10. SOCKET EXTENSION HANDLE11. SOLDERING IRON1. SOLDERING FLUX 6. LINEN SERVING CORD2. SPOT-TYING CORD 7 & 8. CRIMPED TYPE TERMINALS3. LACING CORD 9, 10 & 11. SOLDERED TYPE TERMINALSU. CLEAR LACQUER 12. NUMBERED CELLULOSE TAPE5. WIRE SOLDER 13. FRICTION TAPE6. WIRE HOLDING CLAMP FOR PULLING ON INSULATIONFigure 9— Electrical Repair ToolsFigure lO—Electr ical Repair MaterialsRESTRICTED[22b]

numberindricalinesthe\IIIIRESTRICTED SECTION IXPAR. I TO ^SECTION 9MISCELLANEOUSI . GENERAL.TUBES AND TUBING REPAIRSPractically all tubing in this airplane ismade of 52S0 and 53SW aluminum alloy. A relativelysmal I of I are made of copper,corrosion-resistant steel, and 2S-I/2H aluminumalloy. The tubular lines also vary in size,shape, and end finish. Damaged lines shouldbe replaced with standard parts from North AmericanAviation, Inc., Ingle^ood, California, U.S.A.However, replacement is sometimes impractical,and therefore a repair method is outlined. Repairprocedures described in the following paragraphsare applicable to all lines, regardless of theirmetal composition. Small scratches and abrasions,or minor forms of corrosion occurringon the exterior, after having been smoothed outwith a burnishing tool or steel wool, may beconsidered as negligible.the die approximately one-half the outside diameterof the tube, and secure the assemblyin a vise. Rest the tapered end of the flaringtool in the tube and tap lightly with a hanmerunti I wal Is of the tube are forced to assumethe shape of the countersunk hole in the die.(See bigure 1. )TJ44IIFLARING TOOL GRIP DIE SHOWNIN VISE2. TUBE JOINTS.Two types of joints are used at pressure tubeconnections. The beaded or upset type of jointis used in the vacuum, fuel, and oil systemswhere hose connections are utilized. All otherjoints are of the flared type, to be used inconjunction with tube connector fittings. Gripdies and flaring or beading tools are requiredto form flared and beaded-type joints.3. FORMING FLARED TUBE JOINTS.The TJ 4411 grip die and flaring tool (seefigure 4, Detail B) is recommended for formingflared tube joints. The flared-type grip dieconsists of two steel blocks placed side by sideand held in alignment by three steel pilot pinspressed into one block and extending into correspondingholes in the other block. (See Figure1- ) A number of countersunk holes are dri I ledalong its length, the holes varying in size tocorrespond to the tube sizes. The grip die shouldbe designed to grip the tubes tightly withoutdamaging them. The flaring tools for this typeof joint consist of cy I bars tapered atone end to correspond to the angles of the countersunkholes in the die. To make a flared tubeend, insert the end of the tube into the grip diethrough the space provided when the two blocksare parted. Al low tube to extend upward throughTUBINGTJ44II GRIP DIE7^,^ n nWP VIEWSIDE%VIEW^r 7 —Figure l-'Tube Flaring Tools^, FORMING BEADED TUBE JOINTS.When making a beaded-type joint, a separategrip die should be used for each size tube.This grip die consists of two steel blocks placedside by side and held in alignment by suitablepilot pins. A hole, equal to the outside diameterof the pipe and chamfered slightly aroundthe top, is drilled down the center of the die.(See Figure 2.) The beading tool consists ofa cylindrical rod with a chamfered hole equalto the outside diameter of the tube drilledapproximately 1/4-inch into one end of the[ 229]RESTRICTED

I beled—inchl-BT\Burr-HTMISCELLANEOUSPAR. 4 TO 6RESTRICTEDcenter. A hole, equal to the inside diameterof the tube, is dri I down through the centerof this hole and the rod to a depth of I inch.A steel pilot is driven into this hole andpermitted to extend approximately I belowthe tool. To make a beaded joint, insert thetube into the die through the -space formed whenthe two blocks are parted, allowing the pipeto extend upward from the die approximately1-1/2 diameters. Secure the assembly in a vise.Place the pilot of the beading tool into theopen end of the tube and strike lightly untilthe desired bead is formed. (See Figure 2.) AnI inea fitting, union, or connection, cut thebeyond the point of damage and remove the lengthof line which contains the injury. If damageis not near a fitting, the line must be cut onboth sides of the damage. Lines must be cutin locations where cutting, burring, and otheroperations can be accomplished. The use of akeyhole hacksaw is recommended for cutting thetubing.CHAMFER DIAMETER TO SUIT TUBING SIZEPILOTSGRIP DIE (TOP VIEW)VISEINSIDE ROLLERDETAILGRIP DIE IN VISE^ ^--tubingFigure 2 — Beading Small Tubingalternate method of beading tubes of large diameter,utilizing a hand roller, is illustratedin Figure 3.b. GENERAL TUBING REPAIR.For the purposes of repair, the tubes maybe divided into three groups consisting of:the high-pressure, flared end group; the lowpressure,beaded end group; and the conduitgroup which retains no pressure and requiresno end finish. All tubing is repaired by replacingthe damaged section with a new lengthof similar material. The length of tubing removedwi I control led by the location ofthe damage, the extent of the damage, and theniost convenient location for tool manipulation.When the damaged portion of the line is nearFigure 3-^Beading Large Tubing6. REPAIR OF HIGH-PRESSURE TUBING.Cut out and remove damaged portion of tubeas directed in the preceding paragraph (seeFigure 4, Detail A) .the ends of undamagedtube portions, and clean. Slip anAC- I nut and an AC-81 l-T sleeve cf propersize and material onto the remaining tube ends.Flare the tube ends using a TJ 4411 grip dieand a flaring tool of proper size and a hammer(see Figure 4, Detail B) . Support rear sideof grip die to prevent tube bending when hammerthrusts are applied. Tap the flaring toolinto end of tube until the walls of the tubeassume the shape of the countersunk hole inthe die. Insert an AC-81 I union of properdiameter and material into flared tube ends.Apply Sea lube to threads and slip sleeve downagainst flared tube end. Thread the nut ontoRESTRICTED[230]

A,)RESTRICTEDMISCELLANEOUSPAR. 6 TO 7the union and tighten. (See Figure 4, Detail C.Measure accurately' the distance between theunion faces (Length "E" , Detail C, Figure 4J.If the length of the damage is several inchesor the damaged tubing is to be removed to thenearest fitting, a repair insert must be used.Cut a repair length of similar tubing; cleanand burr the ends; slip on two s leeves and twonuts in proper sequence; and flare the tube endsas directed in Paragraph 3. Completed repairinsert is shewn in Figure 4, Detail D. Place7. LOW-PRESSURE LINE REPAIR.Remove the damaged portion of the tube asdirected in a foregoing paragraph, burr theremaining tube ends, and remove cuttings fromtube interiors. (See Figure 6, Details A and B.)If possible, bead the remaining tube ends asoutlined in Paragraph 4. Cut a repair section1/2-inch shorter than the length of the removeddamaged portion. Burr the section and beadas prescribed in a foregoing paragraph, and^ui^ IIIDJ^CUT DAMAGED AREAMc4«ir^cur OA MA CEO AREABAC-8H-HT UNION( I REQ)OBTAIN UNION INSERT^ AC-8I1-BT NUT (2 REO) AC- 8II-T SLEEVE (2REQ)AC-8I1-HT UNION (2 REO )DEADD TUBE END UNIONS.«G\AC-8II-8T NUT (2 REQ)flC-8ll-T SLEEVE (2 REO)FLARE ENDSVPREPARETUBE ENDSFLAREDTUBE ENDSAC-8II-BT NUT (2 REG)AC-8II-T SLEEVE(2 REO)-D-cQrCteCOMPLETE REPAIR:^COMPLETE REPAIRFigure 4— High-pressure TubingRepair of Major Damagethis repair sect ion between the two unions; threadon the two nuts and tighten. (See Figure ^f,Detail E.) The use of special wrenches adaptedfor the hex nuts on high-pressure line connectionswill, in many cases, expedite repairs. Occasionsarise when restrictions prohibit the use of openend wrenches. (See Figure 8, Detail A, for tvotypes of wrenches and Detail D for their use.)If the damaged portion of the line does notexceed the length of a union, a repair sectionis not needed. Cut out the damage as directedin Figure 5, Detai I so that after the tubeends are flared as shown on Figure 4, DetailB, and a union is inserted, a tight fit willresult. (See Figure 5, Detail D.)Figure 5—High-pressure TubingRepair of Minor Damageclean. (See Figure 6, Detail C.) Cut hose connectionsof proper length and diameter fromsynthetic rubber hose and slip two hose clampsover the ends of each. Slip one hose connectionwell back over original tube before placingrepair insert into position. Slip hose connectionsmidway over the junction formed bythe original tube and the repair section asillustrated in Figure 6, Details D and E.Tighten c lamps. G lycerin appi ied to a meta Ipipe aids the sliding of rubber hose,Note: Two hose clamps on each end of a connectionare normal on some beaded oillines, as these lines maintain a relativelyhigh pressure.[231 ]RESTRICTED

I under)MISCELLANEOUSPAR. 8 TO 10RESTRICTED8. CONDUIT REPAIR.Before starting repair to conduit, removeall wires from the tube interior. Cut out thedamaged portion of the conduit in accordancewith Paragraph 5, and as illustrated in Figure7, Detail A. Burr and clean the remainingDcur DAMAGEDAREA9. MISCELLANEOUS.I iVarious lines such as instrument, drain, andoverflow nes fa I the high-pressureand the low-pressure groups. Although someof these maintain little or no pressure, theymust be treated and repaired as high-pressurelines, as no leaks are permitted. If the fittingsof the damaged line are of the flaredtype, repair as outlined in Paragraph 6. Ifthe lines are assembled by hose connections,repair as outlined in Paragraph 7."B•»-BEADBURRBEADENDSDTUBE ENDSjlHOSE CLAMP AC-745-TYPE A(4 REQ)SYNTHETIC RUBBER HOSE (2 REQ)cur DAMAGEDAREAMINUS Vi INCHPREPARE INSERT ASSEMBLYSECONDHOSE CONNECTIONREPAIR INSERT ^ ORIGINAL TUBEBURR AND CLEAN TUBE ENDSCUT ONE CONDUIT UNION IN HALF-JK. MINUS '/g INCHAC-33A-243I CONDUIT UNION (2 REQ)PREPARE INSERTCONDUIT NUTSAC- 31-1177 (4 REQ )Figure 6— Lav-pressure Tubing Repairtube ends. (See Figure 7, Detail B.) Cut anInsert of identical tubing the same length asthe removed damaged portion. Burr the endsand clean the tube interior. (See Figure 7,Detail C.) Place a conduit union on each endof the repair insert and place this assemblybetween the remaining ends of the original tube.If rigidity of the original tube ends prohibitsthis type of insertion, one or both union sleevesmay be cut in half, opposite normal slit. (SeeFigure 7, Detail C.) Place conduit nuts overthe tubing before placing repair insert intoposition. Apply a thin coating of thread lubeto the union threads. Hold union halves inposition around the junction formed by theinsert and the original tube end; screw onconduit nuts and tighten. Completed repairis shown in Figure 1, Detail D.10. TOOLS.Figure 7 — Conduit RepairAll or part of the fol lowing tools are necessaryto complete tubing repairs. All toolshaving a TJ part number are made by, and canbe purchased from. North American Aviation,Inc., Inglewood, California, U.S.A.WRENCH,TOOLSSPECIALNUT TJ 2802REt-IARKSLOOSENING OR TIGHTENINGAC-811-BT NUTS. ALL RE-QUIRED SIZES AVAILABLE.(See Figure 8.RESTRICTED[232]

TNGlowisoredutIIMUMLABLENCHES.ONSAMETERRESTRICTEDMISCELLANEOUSPAR. 10 TO 12TOOLSREMARKSTOOLSREMARKSWRENCH, OPEN-ENDTJ 601LOOSEN I OR T IGHTEN INGHIGH-PRESSURE AND/OR CON-DUIT FITTINGS.SIZES RANGE FROM l/2-INCHTO 1-1/2 INCHES.TOOL, BEADING,(SEE FIGURE 2.)WRENCH, CONDUITNUT TJ 2332HAND BEADING TUBES OF DI-AMETERS UP TO 5/8~INCH.USED TO GRIP CONDUIT NUTS.GRIP DIE OR FLARINGTOOL TJ 4411USED TO FLARE TUBE ENDS.BEADING TOOL, HANDROLLER (SEE FIG-FOR BEADING TUBING FROM 5/8-INCHUP.HACK SAW, SMALL ORKEYHOLECONVENIENT FOR CUTTING TUB-ING IN RESTRICTED PLACES.II.URE 3-)MATERIALS.WRENCH,CONDUITNUT TJ 2488LOOSENING OR TIGHTENINGCONDU I NUTS. Dl AMETERSOF l/2-INCH TO 1-1/2 INCHESAVAILABLE.The fo I09 ismay be requ i toof tubing repai r.MATERIALSa list of materials whichaccomplish the three typesREMARKS\CUT WRENCHTOTUBESLIP OVERTUBING, REPAIRLENGTHSTUBING, SYNTHETICRUBBERMUST BE IDENTICAL TO DAMAGEDTUBING IN MATERIAL, DIAMETER,AND WALL TH ICKNESS-MUST BE IDENTICAL TO OTHERCONNECT I FOUND ON THELI NES OF THE SYSTEM BE! NGREPA I RED.NUT,ACBll-BTSLEEVE, AC811-TUNION, AC811-HTHIGH-PRESSURE, FLARED-TYPEFITTING REPAIR.CLAMP,HOSEIAN745, TYPE AMINIMUM DIAMETER AVAILABLE9-1/6 MAXIMUM DI-AMETER AVAILABLE 3" 1 / '^NCH ES.CLAMP,HOSEAN745, TYPE BMINIMUM DIAMETER l/2-INCH.MAX I D I l-l/SNCHES.NUT,CONDUITAC31-1177U SED ON CONDUIT UNIONS.Figure 8— Special Wrenches forHi gh- pressure ConnectionsUNION,CONDUITCHERRY BLIND RIVETSAC33A-2431AVA IIN SI ZES FROM1/4-INCH DIAMETER TO 2 INCHES.2. GENERAL. Rivet Company, Los Angeles, California, offera pract ica I I i on to the problem of blindCherry blind rivets, manufactured by the Cherry spot riveting, as they require no Ducking.[233]RESTRICTED

owledf-f-uggI onuggI DENTICALI NALLYLOTESZEDNCHMISCELLANEOUSPAR. 12 TO mRESTRICTEDDesigned to take care of the difficult rivetingjobs in inaccessible places or in double-surfacedstructures where access to both sidesof the work is impossible, this type rivet isrecommended for repair work whenever standardriveting is impractical. Cherry rivets areavailable in two types, the ho I Ida/ and the selfpluggingtype. Being the stronger of the two,the se I p lugg ing type is the only one recommendedfor repair work. The shear strengthof se I p I ing Cherry rivets of 5/32- inchdiameter and less is 60 percent of the standardrivet shear strength specified in the ANC-5Manual (Strength of Aircraft Elements). Forrivets of 3/16-inch and larger, the shear strengthis 50 percent of that of standard rivets of thesame diameter and material.13. SELF-PLUGGING CHERRY RIVETS.The se If- pi ugg ing type has a stem or mandrelwith an expanded section. This expanded sectionis pu I into the hoi low member and stays inplace when the rivet is applied. When thisrivet, with the stem projecting from the headend, is placed in riveting position, the gunfits over the stem and exerts pressure on thehead and a pu I the stem. Thus the preformedhead on the stem is forced into thetail of the rivet. This expands the tail andtightens it down against the back of the piecesto be fastened together". The pull on the stemis continuous until the stem breaks and fallsout on either side of the rivet. The remainingend of the stem may then be nipped off if aflush surface is desired. (See Figures 12,23, 14. and 16.) The expansion of the selfpluggingtype is particularly effective inf i I I ing the hole and g ives a good t ight jobwith satisfactory shear and bearing values.Three d iameters are available: I/P-, 5/32-, and3/j6-inch. Each diameter is made in a seriesof grip lengths and, with the sizes regularlymade, it is possible to handle total thicknessesof material from .030 to .265 inch.m. TYPES OF SELF-PLUGGING CHERRY RIVETS.Four types of se I f- p I ing Cherry rivetsare available which are recommended for repairs.These types are listed and described in thef ol I ing table:LS1126TYPESERIESREMARKSTHIS IS AN AD RIVET MADEOF A17ST ALUMINUM ALLOYAND REQUIRES NO ADDITIONALHEAT TREATMENT. THE HEADBRAZIER HEADTYPE LSII27 OR LSII29LS1127LS1128INSER TEDC ' SUNK HEA DTYPE LSII26 OR LSII28D=GRIP LENGTH REQUIRED D + F=GRIP LENGTHREQUIREDrf^SHOP FORMEDHEADEXPANDEDmFINISHEDRIVETTRIMMEDD= TOTAL MATERIAL THICKNESSE = l" MINIMUM TO /^" MAXIMUM64 64F = AMOUNT OF DIMPLETYPEFigure 9-^Cherry Rivet Grip LengthSERIESSERIESREMARKSAND SHANK Dl MENS IONS ARETO THE NEW 100-DEGREE COUNTERSUNK AN426R IVET. THE PLUGGING STEMIS MADE OF 17ST ALUMINUMALLOY, ADDING TO THE SHEARVALUE OF THE RIVET.THIS RIVET IS THE SAME ASTHE LS1126 EXCEPT THAT THEHEAD AND SHANK HAVE THED IMENSIONS OF THE WESTERNSTANDARD BRAZIER.ITH IS SER I IS DES IGNEDPRIMARILY FOR USE IN DOUBLE-DIMPLED CONSTRUCTION WHERETHE P IHOLE SIZE INTHIN SHEETS HAS BEEN EN-LARGED FROM .010 TO.020 INCH OVER THE NOM-S IFRACTI ONALDIAMETERS. THE SHANK DI-AMETER HAS BEEN ENLARGEDRESTRICTEDI 234 ]

ecosRESTI?ICTEDMISCELLANEOUSPAR. m TO 19TYPEREMARKS.099 INCH TO .016 INCH OVERTHE NOMINAL FRACTIONAL DI-MENSION IN ORDER TO FIT THEENLARGED PILOT HOLES- TESTSTO DATE INDICATE THAT TH ISRIVET WILL BRING THE SHEARAND FATIGUE VALUES OF THEDIMPLED JOINT NEARLY EQUALTO THE EQUIVALENT JOINT WITHSOLID BRAZIER HEAD RIVETS.-2 FROM .030 TO .077 (I/I6-INCH)-4 FROM .078 TO .140 (1/8-INCH)-6 FROM .141 TO .203 (3/I6-INCH)-8 FROM .204 TO .265 {I/4-INCH)16. DRILLING PREPARATIONS.In dri I I ing holes for the use of Cherry bl indrivets, care should be taken to have the holethe correct size and at right angles to thesheets to be fastened together. The followingfinished hole sizes are recommended:LS1129SERIESTHIS SERIES HAS THE SAMESHANK DIAMETER AS LS1128,BUT HAS WESTERN STANDARDBRAZI ER HEADS.15c DETERMINING CHERRY RIVET GRIP LENGTH.The grip length of Cherry rivets is measuredin thirty-seconds of an inch and appears asthe last dash number following the Cherry rivettype number. Example: LSII27-5-6 refers toa rivet with a grip length_to handle materialhaving a total thickness of 3/16-inch. Theproper grip length to use is determined bymeasuring the total thickness of the sheetsto be fastened together. Select a rivet whichhas a nominal grip length as close as possibleto this measurement. There is, however, aconsiderable amount of tolerance, in that Cherryrivets will handle material which is 1/64-inch(.016 inch) th icker or 3/64-inch (.047 inch)thinner than the nominal or recommended griplength. The total length of an unexpanded rivetshank includes approximately 1/16-inch additionalfor proper heading. Therefore, to ascertainthe grip length of a Cherry rivet in the eventthis information is unknown, measure the shankand subtract 1/16-inch. Where dimpled sheetsare to be used in connection with countersunkrivets, determine the proper grip length bytaking the total thickness of the sheets plusthe amount which the dimple extends beyondthe inside surface, or add the height of thehead to the thickness of the sheets beforedimpling. Example: If the sheets are .052inch in total thickness and the dimple extends.038 inch on the inside of the work, then thetotal grip length to be handled would be .090inch. This is .035 inch less than 1/8- inch whichis within the limits covered by the 1/8-inch griplength. The required grip length would be asindicated by -4. (See Figure ii. } The rangeof application of last dash numbers is as follows:FOR 1/8-INCH RIVETFOR 5/32-INCH RIVETFOR 3/16-INCH RIVETNO. 30 DRILLNO. 20 DRILLNO. 10 DRILLRecommended finished hole sizes are shownplainly on each box of rivets as well as thelimits of grip length. In special cases, thesemay vary from the standard given above and careshould be taken to meet the dimensions givenon each box.17. DIMPLING PREPARATIONS.In general, material for the use of countersunkCherry blind rivets should be prepared inthe same way as though sol id rivets were to beused. In the use of countersunk Cherry bl indrivets in dimpled holes, the usual steps must betaken to ensure proper straight-sided holes.Chips must be carefully eliminated from betweenthe sheets, and any burrs resulting from drillingshould be removed, as they prevent properseating of the preformed head as well as thetulip-type head on the bl ind side of the work.18. SHEET CLAMPING.The use of skin fasteners and/or C I atreasonable intervals is recorrmended for holdingsheets together before drilling or riveting.If skin fasteners and Clecos are not available,sheet metal screws may be substituted.19. CHERRY RIVET INSERTION.The hole sizes recommended in a foregoingparagraph are large enough for the rivet toslip in easily (see Figure 12). It is notnecessary to have a snug fit. Careless handlingof Cherry blind rivets sometimes results innicks or dents on the thin edge of the preformedhead. As this may prevent the headfrom seating properly, such rivets should be[235]RESTRICTED

ofsizesnnlyMISCELLANEOUSPAR. 19 TO 21RESTRICTEDFigurelO—'Inserting Cherry RivetsFigure11-^ Expanding Cherry Rivetsdiscarded. Several rivets may be placed intheir respective holes and then expanded byapplying the rivet gun to the projecting stems.(See bigure lo.) In some cases insertion ofthe rivet into the gun head and then into thehoi e is p referabi e.20. TYPES OF CHERRY RIVET GUNSTwo types of Cherry rivet guns are manufacturedby the Cherry Rivet Company, a pneumaticand a hand type. These guns have interchangeableheads and can handle al I ofCherry rivets. Each head has two adapter caps,permitting the use of brazier head or countersunktype of rivets. The pulling heads onboth guns operate on the same principle. Eachhas a slotted end into which the stem of theCherry rivet is inserted. Upon the applicationof force, an inside draw bar pul Is the headof the rivet against the gun and, by continuedpull, forces the inside end of the stem intothe tai I the rivet and eventual ly breakstne stem. (See Figure 12- ) Tfie Gl^ rivet gun,a pneumatic tool, operates on 75 to 90 lbs./sq.in. air pressure and exerts a 1300-poundpull on the rivet stems. The gun weighs 4pounds and is balanced for easy handling atany angle. (See Figure ii.) The GIO gun is animproved, ratchet-type, hand-operated gun whichcan be used where air power is not available.(See Figure 14.) The ratchet feature permitsthe operator to exert an intermittent pull untilthe stem of the rivet breaks. A G6AA, 90-degreeangle adapter is available for both types ofgun. This adapter permits Cherry rivets tobe inserted in locations where there is limitedroom for gun operation, such as between ribs inthe wing structure.21. USE OF CHERRY RIVET GUNS 'Whenever applying a Cherry rivet gun on arivet, make certain the rivet stem fits wellinto the bottom of the drawbolt slot beforeapplying pressure. This should prevent shearingoff the head of the rivet and ensure the breakingof the stem. Care should be taken to see thatthe head of the rivet is held f i and squarelyagainst the sheet to which it is being applied.This simple precaution will do much to ensureuniform and smooth results. The more nearlythe axis of the pulling jaws is parallel tothe axis of the rivet, the better the finaljob will be. Special instructions for thecare and handling of each gun are deliveredwith the gun. Care should be used to see thatproper nosepiece is fitted to the head of theRESTRICTED[ 236 ]

iRESTRICTEDMISCELLANEOUSPAR. 21 TO 22BRAZIER-HEADADAPTER CAPUNTRIMMEDSTEMBROKEN STEMFigure 12-—Untrimmed Cherry RivetsFigureIJ-^Tr imming Cherry Rivetsgun to handle either countersunk or braziertyperivets in each diameter. It is necessaryto have a separate size of pulling head foreach diameter rivet. These heads are readilynterchangeable.22. CHERRY RIVET GUN MAINTENANCEThe maintenance and functioning of the gunare important. Care must always be taken tohave the gun in proper working condition. This^— G6H6 HEAD FOR 3/jg":>sMSs^- G 6H5 H EAD FOR y^2W G 6H4 HZ CHR FOR!/„• I/b"G-IO HAND GUNFigure14— Cherry Rivet Hand Gun[237]RESTRICTED

MISCELLANEOUSPAR . 22 TO 26RESTRICTEDcan be determined by measuring the pullinghead. In guns designed for self-plugging rivets,the movement is 3/4- inch minimum. Decreasein the length of the stroke may result in fai lureto break the stem of the rivet. Sometimes,instead of breaking the stem, the outside headwill pull off. This is not a serious fault.Drill out the remaining portion and insertanother rivet. V.lien drilling out Cherry blindrivets, use a drill a size smaller than the diameterof the Cherry rivet stem. Punch the remainingportion out with a drift punch.23. TRIMMING CHERRY BLIND RIVETS.When the self-plugging rivet is applied, thestem usually breaks from 3/8-inch to 1/2-inchoutside the preformed head. In most cases, itis desirable to cut off this stem. Such cuttingcan be done with an ordinary pair of multiplyingnippers or diagonal cutters. (See Figure 13.>!These should be ground dcwn to a flat surfaceon the outside, and the inside should be hollowground. If properly adjusted, such nipperswill do sat isfactory work.DUPONT EXPLOSIVE RIVETS21. GENERAL.Du Pont explosive rivets, manufactured byE.I. du Pont de Nemours and Co., Inc., Wilmington,Delaware, present another solution to the problemof blind riveting encountered in field repairof aircraft. These explosive rivets are similarin appearance to standard solid rivets and theriveting preparations are the sarre. After insertion,a heated rivet expanding iron is appliedto the head of the rivet and causes the rivetto explode and consequently expand on t he blindside. This expansion forms a head on the blindside which grips the metal and prevents extractionof the rivet. Du Pont explosive rivetsare suggested only as a substitution for solidrivets or Or^erry blind rivets and should notbe used indiscriminately.25. RIVETING PRECAUTIONS.grip length must be used (see Figure 25J . Therivet must be expanded in 1.5 to 6 seconds witha recommended heating iron and a proper sizetip. Parts to be riveted must be held firmlytogether at the time of rivet expansion. Donot use these rivets in steel or other metalsharder than 24ST aluminum alloy. Do not attemptto expand rivets in any other way than with thestandard du Pont riveting iron (see Figure 16).26. STORAGE AND HANDLING PRECAUTIONS.Handle du Pont explosive rivets with care.Keep the rivets in the boxes in which they areDU PONTRIVETINGIRONInasmuch as explosive rivets present a firehazard when used on fueled airplanes, all necessaryfire precautions must be taken. If theblind side of the riveting is in a closed compartmentor bay, as will general ly be the case,bloA' the compartment or bay with carbon dioxideor live steam to neutralize any vapors whichmight exist. Holes must be drilled carefullyso there is a close fit between rivet and hole.Rivet of correct head type, shank diameter, andGRIPLENGTHINSERTED —^EXPLODED-Figure 15 — Explosive Rivet Cross SectionFigure 16—Use of Expanding IronRESTRICTED[238]

woodenapp.RESTRICTEDMISCELLANEOUSPAR. 26 TO 31received and store in a dry place, preferablyin a closed cabinet. Keep open fires away andavoid temperatures in excess of 38°C (IOO°F).Do not throw unexpanded rivets into trash. Ifrivets become distorted so that they cannot beused, place them in a wire basket having holessmal ler than the rivets and destroy them byplacing the basket in a fire.27. RIVET INSTALLATION.Push rivet into drilled hole. Do not forceto the extent of scraping the protective filmon the shank. If necessary, tap lightly witha smal I or rubber mal let, taking carenot to damage the head of the rivet. Inspectto make certain the rivet head is well seatedand the rivet shank fits snugly in hole. Makecertain the rivet is of proper grip length anddiameter (see Figure i^) Apply a heated rivetingiron to the head of the rivet, holdingthe tip firmly against the rivet head. Removetip from rivet immediately after expansion.Important: Du Pont explosive rivets shouldbe expanded within I. 5 to 6 secondsafter heat is applied.28. DU PONT RIVETING IRONS.The No. 3 iron is the most common rivetingiron, for it can be used to expand all diametersand lengths of du Pont rivets (see Figure i6)Six t i ps a re ava liable fo r th i s i ron and it isimportant that the proper tip be used when expandingdu Pont explosive rivets. Read carefullythe instructions on the boxes in whichthey are shipped. The LOV setting on the wattageregulator is for short rivets. The HIGH settingon the wattage regulator is for long rivets.For intermediate length rivets, the regulatorshould operate at a point between LOW and HIGHin proportion to the length of the rivet. Trya sample in open sheet stock for proper expansion.The wattage regulator should be set so thatthe rivet will exp lode wi th in 1,5 to 6 secondsafter application of the expanding iron.RIVNUTS29. GENERAL.A Goodrich rivnut is an internally threadedand counterbored tubular rivet that can beheaded blind. Rivnuts are of a one-piece constructionand are accurately machined fromcorros ion-res istant al uminum al loy . Rivnutswere designed primarily to be used as nut platesfor the attachment of de-icers to the leadingedges of wings and stabilizers. However, sincerivnuts can be pulled up or headed while workingentirely from one side of the structure,other unique adaptations have been made insecondary and primary structures of modernairplanes. The upset portion or head formedon the far side of the metal is large enoughto resist being pulled through the metal skineven under conditions of eccentric load. Afterinstallation, rivnuts may be used as nut platesor as rivets, or both.30. TYPES OF RIVNUTS.Rivnuts are classified by the head style:namely, the countersunk and the flat head.The former is made in three different headshapes while the latter is made in only one.Each style is made in three sizes. No. 6-32,No. 8-32, and No. 10-32. Each of these sizes isavailable in six grip ranges and can be furnishedwith open or closed ends. With the exceptionof the thin-head countersunk type, all stylescan be furnished with or without keys underthe head. The closed end rivnuts should beused only in special places such as sealedcompartments for flotation bays. Keyed-typerivnuts should be used whenever screws areto be inserted to increase strength or for theattachment of fittings.31. GRIP RANGE.In order to produce an ideal bulge or headon the blind side of the metal sheet, a positiverelationship must be maintained betweenthe counterbore depth of a rivnut and the thicknessof the material in which it is beingheaded. This thickness is known as the grip.The grip range is the thickness of the materialabove and below this ideal point in whicha rivnut may be successfully headed. Sixconsecutive and nonover I ing grip rangesare available and these grip ranges are identifiedby indented markings found on the exposedsurface of the head. Absence of thesemarks classifies the rivnut in the minimumrange of its head style. One mark indicatesthe next higher range and five marks indicatethe maximum range. The grip range ofany rivnut can be determined directly fromits type number, as the figures at the rightin the rivnut type number indicate the fnaximumgrip in thousandths of an inch and theminimum grip for this rivnut is the same as[239]RESTRICTED

MISCELLANEOUSPAR. 31 TO 33RESTRICTEDmV-NUTS- KEY1^RIV-NUTFigure 17— Use of Keyseating ToolFigure18-~Inser t ing Rivnutthe maximum grip of the preceding size, exceptin the case of the first rivnut in a series,where the minimum grip is the thickness ofthe head in the countersunk style and .010inch in the flat head style. For best results,a rivnut should be used only within the griprange for which it is intended.32. RIVNUT TYPE NUMBERS.The first figure of a rivnut number indicatesthe screw size of the internal thread. Thefigure at the right indicates the maximum gripin thousandths of an inch. The minimum gripequals the maximum grip of the preceding size,except that the minimum grip of the first rivnutin a series is the head thickness for thecountersunk types and .0 10 inch for flat headtypes.A dash between figures indicates OPENEND KEYLESSA "B" between figures indicates CLOSEDEND KEYLESSA "K" between figures indicates OPENEND WITH KEYA "KB" between figures indicates CLOSEDEND WITH KEYIf the figures at the right are evenly divisibleby 5, the rivnut is the flat head type. If thenumber is not evenly divisible by 5, the rivnutis a countersunk type.Examples: An 8KB 100 rivnut is an 8-32, CLOSEDEND WITH KEY rivnut and has a maximumgrip of 100 thousandths of an inch (. 100")and a minimum grip of 75 thousandths of an inch(.075"). A 6BI06 rivnut is a ^32, CLOSED ENDKEYLESS rivnut and has a maximum grip of 106thousandths of an inch (. 106") and a minimum gripof 63 thousandths of an inch (.063"), which isthe countersunk head thickness. The right-handnumber ( 106) is not evenly divisible by 5.Therefore the rivnut is of the countersunk type.33. STRENGTH OF RIVNUTS.Before substituting rivnuts for sol id rivetsin all repair work, a comparison of the shearand the tensi le strength must be made to ensurea resultant safe repair. The shear andthe tensile strength of rivnuts may be greatlyincreased by the insertion of standard screwsinto the hoi low stem. When screws are oniitted,the shear strength of rivnuts is approximately30 to 50 percent of that of solid rivets. Theinsertion of standard SAE 2330 screws increasesRESTRICTED[240]

edI ingRESTRICTEDMISCELLANEOUSPAR. 33 TO 35Figure 19-— Squeezing Rivnutthe strength considerably. Before substitutingrivnuts for sol id rivets, consult a qua! if ledengineer. ITie ultimate shear and tensile strengthof solid rivets set forth in the ANC-5 (Strengthof Aircraft Elements) Manual should be comparedwith the shear and tensile strength setforth by the Goodrich Rivnut Company, to determinethe size and number of the rivnutsto be used.3^. PREPARATIONS FOR INSTALLATION.In general, material for the use of rivnutsshould be prepared in the same way as thoughsolid rivets were to be used. As much precisionis requ i in d ri I a hoi e for a ri v-nut as for a regular rivet, since the shankof the rivnut must fit snugly into the hole. Inorder to obtain a smooth round hole, a leadhole smaller in diameter than the body sizeof the rivet should be drilled first and followedby the correct finished hole size.SIZE

ight-duty.MISCELLANEOUSPAR. 35 TO 38RESTRICTEDtools have a threaded mandrel on which therivnut is to be threaded until the head ofthe rivnut is a9ainst the anvil of the headingtool. The rivnut is then inserted inthe hole and the key positioned with respectto the keyway. (See Figure 18.) Then, whileholding the tool so the mandrel is at 90 degreesto the surface of the metal, the mandrelis retracted by squeezing the handles together.This draws the rivnut against the anvil, thuscausing the bulge or head to form in the counterboredportion of the rivnut on the inaccessibleside of the work. (See Figure ig. ) Squeezethe handles together until solid resistanceis felt. Excessive pressure beyond this pointis unnecessary. Turn the handwheel counterclockwiseto unthread and withdraw the mandrelfrom the installed rivnut. (See Figure20.) Hand-type tools are illustrated.Pneumatic power rivnut drivers are availableand threading, upsetting, and withdrawal areaccomplished by compressed air through themanipulation of finger-tip controls.DZUSFASTENERS36. GENERAL.Dzus fasteners which are broken or damagedmust be replaced with new ones (see Figure 21)-The I type which is found on box covers,access hole coverplates, and fairing sectionsare easily removed and replaced. The heavydutytype, which is generally used on cowlingand heavy fairings, requires two additionaloperations, inasmuch as the retaining gronimetmust also be removed and replaced.37. REMOVING LIGHT-DUTY DZUS FASTENERS.The removal and replacement of light-duty dzusfasteners are accompi ished with the use of a setof TJ I033A punch and die dzus fastener toolsor facsimiles. Center the broken dzus fastenerover the hole in the anvil with the grooved endup. Place the proper-sized drift punch squarelyon the end of the fastener. Hold the punchfirmly in place and strike the punch a sharpblow with a hammer (see Figure 25? J . After thebroken or damaged dzus fastener is driven outin this manner, the hole will be too large andwill be dimpled slightly on the wrong side.Remedy this condition by flattening the holearea with the flat end of a bal l-peen hammerand an anvil (see Figure 2q)38. REPLACING LIGHT-DUTY DZUS FASTENERS.After the broken fastener has been driven outand the hole area flattened, as set forth in thepreceding paragraph, the hole must be red imp ledBENTFRACTUREDFigure 21 — Typical Damage to Dzus Fasteners Figure 22—Dzus Fastener ToolRESTRICTEDL 242 J

RESTRICTEDMISCELLANEOUSFIGURE 23DIMPLE BACK SIDE OF HOLEDRIVE IN NEW FASTENERFigure 23— Light-duty Dzus Fastener Replacement[243 JRESTRICTED

affectwax..MISCELLANEOUSPAR. 38 TO il3with a proper size and degree dimpling tool(see Figure 23 j. After the hole has been dimpled,turn the lid or coverplate over and fit it onthe anvil so the dimpled hole is directly centeredover the hole in the anvil. Insert anew fastener into the special driving tool, insertthe grooved end of the fastener into thedimpled hole, and drive into place with a hamner(see Figure 29 >

I ightlyI ightlyk's-eyes"eledIyRESTRICTEDMISCELLANEOUSPAR. H3 TO 45ALWAYS WETSANDPAPER, ANDAVOID EXCESSPRESSUREDEEP SCRATCHintended to suggest nethods which could be followedin the shortest possible time and withthe simplest possible equipment, obviously anydamaged pieces should be replaced as soon aspossible. Even when adequate repair facilitiesare available, replacement of the part may beless expensive than attempting a repair. Inqeneral, it will be found economical to buffalmost any scratched section, but patch in maynot be justified except on large and complicatedformed sections. Besides the considerationsof cost, it should be borne in mind that evena carefully patched part is not optically orstructurally equal to a new section.45. ISOLATED CRACKS.SAND WITH ROTATING MOTION OVER A WIDEAREA TO PREVENT OPTICAL DISTORTIONSNOTEDO NOT SAND UNLESS ABSOLUTELYNECESSARY TO REMOVE DEEP SCRATCHES-Figure 26—Removing Deep ScratchesTransparent Plastic Panelsby the procedure outlined for mi nor scratches,it may be necessary to sand the area around thescratches. It is to be noted, however, thatpanels should not be sanded unless absolutelynecessary. Wrap a sheet of 320A fine sandpaperaround a block, wet the sandpaper, and sandover a wide area with a free circularmotion (see Figure 26). Do not confine thesanding to too small an area, or objectionabled istort ions or "bul I will resu It in thetransparent plastic. \Afeish the surface; then sandwith a wet sheet of 400A or finer sandpaper.When the primary deep scratches are removed,remove the fine sandpaper scratches bythe procedure outlined in the paragraph on MinorScratches. If a buffing wheel is aval lable,a more satisfactory method is to apply a preliminarypolish to the deep scratches by meansof a felt disc coated with a mixture of jeweler' srouge and water. Rotate the disc at approximately250 RPM. Apply lightly and keep moist.After the depth of the scratches has been reduced,apply a final polish or turpentine andcha I with an 8-to lO-i nch d iameter silk buffingwheel rotated approximately 2000 RPM.Clean the area and apply a wax coat.inAt the first sign of crack inn, a hole 1/8- to3/16-inch in diameter should be dri I at theend of the crack. This simple operation helpsto prevent further splitting by distributingthe strain over a larger area. If the crackis small and repair facilities limited, stoppingt he c rac k w i t h t he d r i I I ed ho I us ua I w i I Isuffice unt i I a more permanent repair or a replacementcan be made. However, if more permanentrepair is to be effected, plu': the drillhole as outlined in the following paragraph,apply Lucite Cement, H-65 , (E.I. du Pont de^temours & Co., Inc.) with a brush to both surfacesof the crack, and then to the crack apply astream of intense hot air with a blower (seeFigure 27J. Play the bicwer over the crack witha circular motion for approximately 5 minutes;then let the rraterial cool. If a hot-air bloNsrELECTRICBLOWERELECTRIC .HEATING IELEMENT )m. REPAIRS - GENERAL.The repairs which could be attempted in thefield are at best makeshifts and should not beregarded in any other light. Although it isFigure 27—Applying Hot-air BlastCracked Plastic Panelto[245]RESTRICTED

I confineleI edges.MISCELLANEOUSPAR. 45 TO ^7RESTRICTEDFigure 28— Applying Torch Flameto a Cracked Transparent Plastic Panelis not available, a torch may be substituted.The torch should produce a soft bluish flameand should not be bicwy. Hold the torch severalinches from the panel (see Figure 28). Thistype of repair may be regarded as permanent.Trim the edges of the hole to a 45-degree taper;then cut a transparent plastic patch of slightlythicker material than the section to be repairedand trim its edges to a sharper angle than thehole (see Figure 29J . Heat the edges of thepatch over a hot-air bloA/er or an alcohol lampuntil the edges are \/ery soft and pliable; thenIforce the patch into the hole and allow it to cool.Remove the patch and soak the edges of the patchin Lucite Cement, H-65, (E.I. du Pont de NemoursK Co., lnc.)for 2 or 3 minutes; then insertthe patch into the hole again. Because theedges are tapered, pressure need be appliedonly on one surface and equal pressure willautomatically be applied to a I Leavethe patch under ight pressure for upwards of24 hours; then after the cement has hardened,sand or f i the edges level with the adjacentsurface. In patching a curved or a large sectionwhere it is not possible to heat the patch, itis possible to obtain a perfect fit by cuttingthe patch first and using it as a template.U6. REPAIRING HOLES.The first step in repairing holes in transparentplastic panels is to trim the hole and surroundingcracks to a circle as soon as possible. Thiswill prevent the development of radial cracksand wi I the damage to a minimum area.7 '£3ORDINARYo TRIM PATCH AND HOLETO TAPERED EDGESO THAT-PATCH HASSHARPER ANGLETHAN MATERIA^X ./=i I®. HEAT EDGES OF PATCHUNTIL SOFT, THENFORCE PATCH IN HOLEAND ALLOW TO COOLUNDER PRESSURE,REMOVE PATCH, APPLYCEMENT. THEN-® AGAIN FORCE PATCH INTOHOLE AND MAINTAINPRESSURE UNTIL CEMENTHARDENS. FILE EDGESLEVEL WITH SURFACE.Figure 29—^Patching Holes inTransparent Plastic Panels

IsightandIRESTRICTEDMISCELLANEOUSPAR. 47 TO 49Imay be prepared from flat stock. The rraterialshould consist of transparent aery late-baseplastic sheet of a thickness equal to the oricjinalpanel. In order to provide sufficient plasticityfor working, hang the trimmed flat panelin a hot-air oven at 104°+ 121 °C ( 220°+250°F).If the hot-air oven is not available, the flatpanel may be heated in hot oil, glycerin, orkerosene. Hot water is unsatisfactory and shouldnot be used. If the heated liquid method isused, immerse the panel for a few minutes, removeit, and allow it to form over a clothcoveredwood or metal form block. Since thewarm plastic readily takes impressions, it mustbe handled carefully with clean cotton gloves.After the plastic sheet has cooled on the form,remove, trim, and polish the panel. To drillthe panel for installation, ordinary drills formetal may be used. Better results are obtained,however, if the dri I are ground with \/erylittle lead (see Figure ^o) . To prevent cloggingand burning of the sheet, lift the drill fromthe hole frequently and remove shavings. Usemoderate speeds and pressure to avoid"grabbing" when the drill penetrates the transparentplastic. If high speeds are used, a coolantmay be required. If the point of the drillpenetrates thin material before the full widthhas entered, the dri I I will tear the sheet. Thefabricated panel must be installed exactly likethe or ig i na I .48. ROUTING EDGES OF PANELS.After a new panel has been cut and formed tofit an enclosure, the edges must sometimes beTHUMBSCREWFigure 32— Routing Inside Curvesrouted or skived. This may be accomplished byseveral methods, and the method used will necessarilybe dependent upon available eouipment.Two practical ways, accomplished with simpletools generally found in any shop, are describedin this and the foMoA/ing paragraph. The tools requiredare a square-ended rotary f i le of at leastI/2-inch diameter and any means of motive pcwercapable of spinning the rotary file at a speedof at least 2000 RPM. By means of metal blocksand a piece of angle iron, form a rout i nq jig(see Figure ji). Cut a slot in one flange of theangle iron and insert a thumbscrew into this slot .By means of this adjustable guide the width ofthe cut can be accurately controlled. Some meansmjst also be provided to control the depth of cutby the rotary file. This must be determined locally,as the method of applying motive paverto the rotary file will also be a local problem.An air supply nozzle at the cutting blade issuggested to keep the rotary file from clogging.Place the edge of the panel against the guideand pass the panel over the cutting blade. Ifa deep cut is to be taken, better results wi Ibe obtained if it is accomplished in two orniore operations. This jig method will satisfactorilyrout the edqes of straight panels orthe outside periphery of curved panels.49. ROUTING INSIDE CURVES.WIDTH "a" of routCONTROLLED BYADJUSTABLE GUIDEFigure 31—Panel Routing JigInasmuch as the method of routing transparentplastic panels described in the preceding paragraphcannot be appi ied to the inner surface ofa curved panel, a method such as illustrated mustbe used. (See Figure qg . ) Obtain a narrowstrip of soft rreta I fomi it to fit the contou r[247]RESTRICTED

"C"IMISCELLANEOUSPAR. i|9 TO 51RESTRICTEDis suggested that a solid rest be provided forthe motor to give added control. F^ss the rotaryfile back and forth over the exposed area, keepingthe flat end of the rotary file against themetal guiding strip (see Figure c^c^j. An alternatemethod can be used when the panels are sma I I.Clamp the motor securely in a vise and pass theexposed portion of the panel back and forth overthe rotary file, keeping the flat end of therotary file against the metal guiding strip,50. COCKPIT ROOF PANEL REPLACEMENT.Figure 33— Curve Routing Detailpanel. With srtRl I clamps or other suitablemeans, clamp this metal strip accurately intoposition along the edge of the panel, exposingonly as much of the panel edge as is desiredto rout (see Figure 99J . When this method ofrouting is applied, the motor which turns therotary f i le must be held by hand and the depthof the cut cannot be accurately controlled. ItI ingA v isua I i nspect ion of the cockp it roof wi Iread i ly determine the procedure for the removaland replacement of individual panels. To removea panel from the windshield assembly, removethe bcw and strips as one assembly by dri Iout the rivets and removing the screws securingthe assembly to the cowling. Disassemble tneframe as required and replace the defectivepanel. If the rubber strips and the spongerubber located in the grooves of the frame arein good condition, they may be used again. Applya plastic caulking compound around the outsideedges of the windshield, such as Hunt's KfesticCaulking Compound or the equivalent.51 . SIDE PANEL REPLACEMENT.To replace any of the side panels in eitherFigure 34— Cockpit Enclosure AssemblyRESTRICTED[248]

1 i ngingittleI looseRESTRICTEDMISCELLANEOUSPAR. 51 TO 5i|of the curved sliding section, pull the EMER-GENCY handle and push the defective panels outwith the hand. Replace with the new panel,utilizing the same rubber strips if in good condition.Ascertain that the EMERGENCY handlelockpins and safety wires are properly reinstalled.To replace a forward top panel of the aft si iding section, removal of the entire rear hoodassembly is necessary; this is to permit thesi iding section to be moved aft. Remove the bowand corner blocks if in good condition. To removethe aft top panel of the section, it is onlynecessary to remove the bav and corner blocks.To replace a forward side panel or top panel ofthe stationary section, remove the four blockshapedstops from the inside of both tracks.Lift the entire section clear. Remove the bowand corner blocks and replace the defective panelsas required. To remove an aft top panel, it isonly necessary to remove the bow and cornerblocks. When ordering a replacement part, besure to give the correct part number. Each panelhas a separate part number and is cal led out on itsrespective assernbly number (see Figure 94J.ALUMINUM CASTINGS ANDFORGINGS52. GENERAL.The repair of castings and forgings by weldingis not recommended. Damaged parts should bereplaced by spares. However, minor repair bywelding of parts not requiring high structuralvalues is permissible, provided that specialpractices inherent in the p/ocess be observed.It should be borne in mind that, in the nBjorityof cases, the use of castings and forgings isutilized as a means of expediting production;therefore, as a means of temporary repair, asubstitute part can be fabricated of some othermaterial and of comparable design. Where fabricationof parts is accomplished, the ultimatestrength and functional requirements of theoriginal must be maintained. See Sectionxilfor the required tensile strength of all castingsand forgings subsequent to heat treatment, andSection I for possible corresponding materialsubstitution equivalents.53. MINOR SURFACE CRACKS.Small surface cracks, scratches, and minordiscontinuities may be relieved by carefullyf i or scraping. Fi I should be such asto present a gradual, smooth filleting of thearea affected. Sharp or abrupt corners shouldbe avoided. The area should be thoroughlycleaned and a clear varnish or other transparentsolution suitable for the purpose shouldbe applied as an aid in determining any subsequentdevelopment of the discontinuity.54. MINOR WELDING OF LOWLY STRESSED MEMBERS.In general, welding of castings and forgingsI varies in practice from that of sheetmaterial, except that necessary allowances forvariations in shape, cross-sectional area, andform must be made. Prior to welding a casting,carefully investigate the nature of const met ion.Look for tapering cross sections which wouldtend to draw the welding heat away and makepreheating of the casting necessary. Observethe possibilities of thermal strains and cracks,which are most likely to occur in castings ofthin-walled or otherwise intricate design.After the proper investigation has been made,thoroughly clean the area by means of a stiffbristledwire brush and gasol ine. If the damageis a minor crack, completely cut or meltaway the crack and a portion of the surroundingstock by means of a file or puddling iron.If a smaIpiece is to be replaced inthe casting, carefully clamp the piece in position,securing loosely to prevent the thermalstrains during welding. Larger castings orcastings having intricate sections should bepreheated by means of a suitable furnace priorto welding. Small castings or castings requiringa weld near the edge of a thin-walledsection may be preheated locally by means ofa welding torch. In either case, preheatingmust be accomplished gradually and at uniformrate. Use a suitable welding flux and a weldingrod of the same material as the casting. Theadded material must be completely molten andmust be thoroughly worked with the end of thewelding rod. Remove accumulated oxides andimpurities from the surface of weld by meansof a wire brush. After welding, heat treatthe casting or forging in accordance with therequired tensile strength specified in SectionXII.L 249 JRESTRICTED

MISCELLANEOUSPAR. 55 TO 57ROPELLERRESTRICTEDFigure 35— Propeller Installed55. MINOR REPAIRS.Visual \y inspect the hub and blades for darmgessuch as nicks and sharp dents on the leadingedges, or gashes on the blade faces. Suchinjuries are particularly dangerous, as theygreatly reduce the fatigue strength at thatparticular point. A failure may result unlessthey are promptly removed. All mars on thesurfaces of the blades are "stress raisers"and cause a stress concentration which mayraise the stress beyond the endurance limit,resulting in a fatigue failure. Sharp dents andnicks or gashes may be removed locally withoutthe necessity for reworking the entire bladesurface. A curved "riffle" file is recommendedfor use in removing the sharp base of the nick.Fine emery cloth or crocus should be used forpolishing. Care should be taken in removingnicks from the blade face to ensure thatthe thickness is not reduced more than isnecessary. It is recommended, as an addedsafety precaution, that the surface be etchedafter the removal of a nick and examined witha magnifying glass to ensure that the nick isentirely removed and that a crack has not started.After inspection, polish the etched area. Propellershaving yery severe nicks or gashes shouldbe sent to an authorized repair depot or returnedto the factory for repair. (See Figure ^^.)WOODEN ACCESSORIES56. WOODEN COCKPIT SEATS.Damages to the wooden cockpit seats may berepaired by following the procedures set forthin Section II for preparing and gluing woodensplice members and patches to plywood panels.Inasmuch as the seats are a nonstructura I partof the airplane, the scope of negligible damageis much greater than that of a structuremember. However, it is essential to haveno rough edges or protruding members attendantafter repair. No repairs should be madethat will interfere with, or result in damageto, the seat harness, the safety belt, orthe seat adjusting mechanism. (See Figure 36. J57. WOODEN FLOOR BOARDS.General methods set forth in Section II forgluing, clamping and reinforcing damaged woodenRESTRICTED[250]

I theRESTRICTEDMISCELLANEOUSPAR. 57 TO 59members may be used to repair the wooden floorboards of the front and rear cockpits on thelater model airplanes. All plywood used forrepairs should conform to AAF Spec. AN-NN-P5IIA. Spruce used for reinforcing and splicingmembers should conform to AAF Spec. AN-S-6a. Casein glue conforming to AAF Spec.C-G-456 should be used for gluing all repairmembers to the original parts. Attach allrepair members to the underneath side of thefloor boards so as not to interfere with anymovements of the pilot's feet.58. MINOR WOODEN ACCESSORIES.Vt€PLYWOODFigure 36-^Wooden Cockpit SeatOther miscellaneous small items on the airplaneare constructed of plywood and reinforced withspruce. The majority of these items are consideredas expendables and their size and minorcost warrants replacement rather than repair.All such items may be purchased from NorthAmerican Aviation Inc., Inglewood, California.If repair is attempted, the primary considerationmust be the function of the part afterrepair. No protruding members should be attendantafter repair and the strength of the repairedpart must equal that of the original.CONTROLRODS59. GENERAL.Various units of the engine control installationand the fuel and oil system installationsare controlled by res i stance welded pushrods or, as in the case of the fuel selectorvalve, by torque rods. Repairs to these controlrods consist of smoothing out negligibledamage, splicing damaged rods, manufacturingreplacement rods, and replacing the clevis,threaded, or ball-bearing rod ends. The repairmust develop the original strength of the rodand be able to withstand a I compression,tension, or torque loads imposed upon it. Allsplices must be so made as not to interfere withthe movement of any of the controls, and specialattention must be paid to locations where controlrods pass through the firewall or throughphenolic fiber fairleads. No welds are permittedwithin a region on the control rod thatwill pass through a fairlead.60. NEGLIGIBLE DAMAGE.Smooth dents not exceeding one tenth the tubediameter or one tenth the tube perimeter, whichare without cracks, fractures, or sharp comers,may be disregarded except to satisfy appearance.Smooth the dent with a half-round finefi le and finish with steel wool. Apply theoriginal finish to the member. (See Section XL)61. SPLICING CONTROL RODS.Severely damaged or badly bent control rodsSTEEL TUBING1025. SPEC X4130. SPECAN-WW-T-846 AN-WW-T-850SEAM WELDED.SPEC AN-T-3 ^ ^ gg,QAN316-4R NUTf—i'*V1^*U2275^BI387 CLEVIS. BI388 THREADED UNITED AIRCRAFTROD END ROD END FITTING qR EQUIVAN4B6 CLEVIS, BI388 THREADED 19-48034ROD END, ADJUSTING ROD END CONNECTORFigure 37 — Control Rod Repair Components[251]RESTRICTED

I rodI alowI surfacesI aMISCELLANEOUSPAR. 61 TO 63RESTRICTEDmay be spliced by welding in a section of steeltubing to match that of the original rod. Steeltubing conforming to four d if ferent specificationsmay be used for the splice member. (SeeFigure 37. J Proceed as follows: Cut the rodon each side of the damage. Obtain a lengthof steel tubing to match the diameter and wal Ithickness of the damaged rod, and cut to sucha length that the length of the rod when splicedwill equal the length of the original rod. Cleanand burr a I ends and file to a 60-degreetaper. DrI INo. 40 hole in the tube to providefor the injection of the corrosion inhibitingfluid. Butt the rod ends together andcareful ly al ign the spl ice member with the originalrod. With suitable clamping devices, holdthe rods firmly in place. With oxyacety lenewelding equipment, apply a weld around the joint,completely filling the beveled gap. (See FiguregS. J For all information concerning oxyacetylene we Id i ng, refer to Sect ion II. Afterall welding has been completed, squirt hot linseedoil into the interior of the tube throughthe previously drilled No. 40 hole. Rotate theBURR AND BEVELEND OF RODOXYACETYLENEWELDCUT OFF DAMAGED ROD ENDFITTING AS CLOSELY ASPOSSIBLE TO ORIGINAL WELDDRILL N0.40 HOLE-BEVEL END OFNEW FITTINGAFTER WELDING, FORCE HOTLINSEED OIL INTO THIS HOLE.ALLOW TO DRAINBUTT FITTING AND ROD TOGETHER AND APPLY WELDFigure 39'-Control Rod Fitting Replacementcur DAMAGED AREA^RILL A NO. 40 (.098 DIA.) HOLEtube so that a I will be covered andthen a I the tube to drain. Plug the No. 40hole with a cadmium-plated, self-tapping screw.62. FABRICATING REPLACEMENT RODS.S»T'BURR AND BEVEL ROD ENDS— T— '^VHBHHHBHHHBHHBII)^DBURR AND BEVEL ENDS OF INSERTDuplicate rods may be fabricated from steeltubing and the proper size and type rod endfittings. Compare those of the damaged rodwith those in the illustration and determinethe type needed. (See Figure 37. J Resistancewelding should be used when this type of equipmentis available. However, oxyacety lene weldingwill suffice. When resistance welding isused, 1/8-inch of the base metal of the tubeand fitting Is used at each welded joint. Therefore,when gas welding the joints, the tubemust be cut proportionately shorter. When uti-I iz ing gas we Id ing, dri I No. 40 hole in thetube and treat the tube interior for corrosionas directed in the preceding paragraph.AFTER WELDING, FORCE HOTLINSEED OIL INTO THIS HOLEWELD INSERT TO ORIGINAL ROD STUBSFigure 38— Control Rod Splice63. REPLACING DAMAGED ROD END FITTINGS.Determine and obtain the proper sizeandtyperod end fitting to replace the damaged one. Witha hack saw, remove the damaged fitting by cuttingthe rod as close as possible to the weldRESTRICTED[252]

RESWICTEDMISCELLANEOUSPAR. 63 TO 6i|securing the fitting. (See Figure ^g.j Filethe rod end and the welding surface of the fitti ng to a 60-degree taper. Butt the fittingand rod together; and with suitable clampingdevices, clamp the new fitting and the originalrod in alignment. Apply a weld around the jointand treat the interior of the rod with hot linseedoil as set forth in a preceding paragraph.ARMAMENT INSTALLATION REPAIRS64. GENERAL.Repairs to the bomb racks and the machinegun installations are accomplished by replacementof the worn or damaged parts. If replacementparts are not available, exact duplicatesmay be fabricated, but all dimensions andmaterials must be the same as the original partto permit proper functioning of the racks andthe machine gun apparatus. (See Figure 40.)FIXED FUSELAGE GUNFLEXIBLEGUNFigure 40— ArmamentInstallations[253]RESTRICTED

BRESTRICTEDREBUSHINGPARAGRAPH ISECTION 10REBUSHING SPECIFICATIONSI. GENERAL.The fol lowing list of bushings used on the AT-6 Series Airplane is divided into four tables. Table Icontains all reducer bushings. Table II contains all spacer bushings. Table III contains all B Standardbushings except the BIOOQ bushings listed in Table II. Table IV contains all bushings which have componentpart numbers and are not standard bushings. In general, the ream tolerances are +.0005, -.001 unlessotherwise noted, and the last dash number on al I Standard bushings denotes the length of the bushing inth i rty-seconds of an inch.TABLE I —REDUCER BUSHINGSMOTE: AC895 reducer bushings are called out on all production drawings.r bush i ngsLOCATION OFBUSHINGINSTALLATION,INSTALLATION,SURFACE CONTROLSINSTRUMENTSELEVATOR ASSEMBLY, COVEREDINSTALLATION,INSTALLATION,INSTALLATION,INSTRUMENTSINSTRUMENTSINSTRUMENTSNOTE: First dash number denotes bolt size of bushing.Second dash number denotes length of bushing in thirty-seconds of an inch.R denotes reamed bushing.

REBUSHINGPARAGRAPHRESTRICTEDTABLE III — B STANDARD BUSHINGSMOTE: Ream tolerances are +.0005, -.001 except where noted.LOCATIONOF BUSHINGINSTALLATION,SURFACE CONTROLSINSTALLATION, SURFACE CONTROLSINSTALLATIONS,GENERALINSTALLATION, HYDRAULIC SYSTEMINSTALLATION, HYDRAULIC SYSTEMSTABILIZER ASSEMBLY, HORIZONTALINSTALLATION, STARTING SYSTEMINSTALLATION,TAIL WHEELINSTALLATION, LANDING GEARFRAME ASSEMBLY, FUSELAGE FRONTINSTALLATION, TAIL WHEELINSTALLATION, LANDING GEARFRAME ASSEMBLY, FUSELAGEREAR SECTION, COVEREDINSTALLATION,INSTALLATION,WINGSFURNISHINGSCOVERED ASSEMBLY, WINGCENTER SECTION

'RESTRICTEDREBUSHINGPARAGRAPH ITABLE IV — BUSHINGS, COMPONENT PARTSDl erances areLOCATION OF BUSHINGAILERON OUTBOARD HINGE BRACKETCENTER AND OUTER WING FLAPSHORIZONTAL STABILIZER, FRONTFUSELAGE ATTACHINGVERTICAL STABILIZER, FRONTFUSELAGE ATTACHINGFRAME ASSEMBLY, FRONT FUSELAGEFRAME ASSEMBLY, FRONT FUSELAGECOCKPIT ENCLOSURE PANEL ROLLERLANDING GEaRSUPPORTLANDING GEAR RETRACTING OLEOATTACHINGFITTINGLANDING GEAR LOCK PIN MECHANISMSUPPORTTAIL WHEEL ATTACHMENTCARBURETOR AIR FILTER LOCATINGSHAFTENGINE STARTER SHAFT BRACKETPITOT TUBE, OUTER WINGCONTROL SURFACE LOCK HANDLECONTROL SURFACE TRIM TAB DRUMFLAP EQUALIZER CONTROL CYLINDERHYDRAULIC LANDING GEAR AND FLAPCONTROLUNIT100 LB. BOMB RACK SWAY BRACE100 LB. BOMB SHACKLE LEVERRADIO ANTENNA THIMBLE ATTACHING

lowingRESTRICTED FINISH SPECIFICATIONPAR. I TO 2,cordance with Specification 3-IOO-H, Sec. E-l ISECTION 11FINISH SPECIFICATIONI . GENERAL REQUIREhCNTS .WHEELSWINGS,The finish for surfaces and parts is in ac-and the markinqs are in accordance with Specification98-24I05-P. Generally, the finishrequirements may be summed up as follows:9EXTERIORWINGS, INTERIOR ANDINTERIOR PARTSWINGS, INSIGNIA2. FINISH AND FINISH MATER I.SPECALUMINUMALUMINUMNO FINISH^98-2U102-LAL SPECIFICATIONS.AILERONS, EXTERIORAILERONS, INTERIOR ANDINTERIOR PARTSANTIGLARECASTINGS, FORGINGSCOCKPITSCOMPARTMENT, LANDING LIGHTCOMPARTMENT, BAGGAGECOWLING, EXTERIOR & INTERIORELEVATORS, EXTERIORELEVATORS, INTERIOR ANDINTERIOR PARTSFAIRINGFLAPS, EXTERIOR & INTERIORFUSELAGE, EXTERIORFUSELAGE, FRAMEFUSELAGE, INTERIOR PARTS,COCKPITSFUSELAGE, INTERIOR PARTS,AFT OF BAGGAGE COMPARTMENTHORNS, CONTROLINSTRUMENT PANELSLANDING GEARPROPELLER, REAR FACERUDDER, EXTERIOR BALANCEPORTIONRUDDER, INTERIOR AND IN-TERIOR PARTSRUDDER, EXTERIOR REAR OFHINGESEATSSTABILIZERS, EXTERIORSTABILIZERS, INTERIOR ANDINTERIOR PARTSSTEPS, PLATESSTRUTS, LANDING GEARTAIL WHEEL FORKTANKS, EXTERIORWALKWAYSNOALUMINUMFINISH*FLAT BRONZE-GREEN NO.TO MATCHYELLOW-GREENFLAT BLACKYELLOW-GREENALUMINUMALUMINUMSPEC.NO FINISH*ALUMINUMALUMINUMALUMINUMYELLOW-GREENYELLOW-GREENNO FINISH*TO MATCHFLAT BLACKALUMINUMFLAT BLACKNOALUMINUMFINISH*98-2U102-LALUMINUMALUMINUMNO FINISH*ALUMINUMALUMINUMALUMINUMALUMINUMBLACKThe fo I are the U.S. Army Air ForcesSpecification numbers (except as noted) for thefinish materials which are quoted in the fol-Icwinq detail finish requirements:MATERIALcadmium platecement, black "plyosyn"cleaning. prior to finishingdope, cellulose nitrate, cleardope, cellulose nitrate, clear(for aluminized dope)enamel, aircraftfinish, aircraft metal partsfinish, aircraft wood surfaceslacquer, cellulose nitratelacquer, clearlacquer, flat blacklacquer, flat bronze-green no. 9lacquer, gloss whitelacquer, yellow-greenmarkings, aircraftmg protective treatmentoil, linseedpaste, aluminumprimer, zinc chromatesealer, wood, liquidsurfacer, wood, liquidzinc plateSPECIFICATIONNO.AN-0Q-P-'i2lt2656298-20007AN-TT-D-51UAN-TT-D-5513-983-10012U115AN-TT-L-513-1583-1583-1583-1583-100-Hl98-2U10598-20010A; E-3BJJJ-0-336AN-TT-A-U61SArf-TT-P-656®1U1131U115AN-P-32tf Z i nc plating may be used in lieu of cadmiumon all parts except cable terminals and steelsheet .0625 inch or thinner which is usedstruct ural 1y.lYellow-green finish shall be formulated with asuitable tinting material. (Sherwi n-wil 1 iamsPX 3076 Tinting Paste or equivalent)*Finish for corrosion prevention shall be inaccordance with the detail sections listedbelow. Whenever t he pecul i ar i ty f a particularpart or assembly prohibits the employmentof the specified finish, the part orassembly shall be given a high degree ofprotection as is consistent with its properfunctioning for its intended use.©zinc chromate primer may be tinted with a suitable t i nt inq mater ial§The use of aluminum paste shall be restrictedto pigmentation of clear dope, lacquer, orvarnish for final coats on fabric, wood, surfacesof uncamouf 1 aged airplanes, aircraft wheels,and metal only where specifically required.®[257]RESTRICTED

I other.Icw-qreen*receiveotherFINISH SPECIFICATIONPAR. 3 TO 10RESTRICTED3. FINISH CODE.The fol Icwing code of finishes, applicable tothis airplane and conforming in detail to theabove finish specification, facilitates themethod of specifying the complete finish andcolor on all North American Aviation, Inc.,d rawings.CODEF - FINISH, PREFIX DESIGNATIONA - ALUMINUM AND ALUMINUM ALLOYSC - COPPER AND COPPER ALLOYSG - GENERALM - MAGNESIUM AND MAGNESIUM ALLOYSS - STEELT - TEXTILEEXAMPLE: FS^20 - Finish No. 20 for steelALUMINUM AND ALUMINUM ALLOYS4. GENERAL FA-0Before assembly, do not anodize any structure,but prime all structures with one coat of zincchromate primer. However, it is to be notedthat alclad, 2S, 3S, and 52S shall receive noprimer nor other finish except between the adjoiningsurfaces of riveted skin butt or skinlap joints or as required to match adjoiningsurfaces. After ccmplete assembly, finish structuresto match adjoining surfaces.5. INTERIOR CLOSED MEMBERS FA-20.Before assembly, do not anodize any structure.Do not apply primer or subsequent ly finish alclad,2S, 3S, or 52S, except between riveted skin buttor skin lap joints or where it forms part ofa finish surface. Apply zinc chromate primerto a I members. After ccmplete assembly ,apply two coats of ye I lacquer to thecockpits and baggage compartment.6. INTERIOR SURFACES, PARTS, AND OPEN ^E^€ERSFA-21.Before assembly, do not anodize any structure.Do not apply primsr or subsequently finish alclad,2S, 3S, and 52S, except between riveted skinbutt or skin lap joints or where it forms partof a finished surface. Apply zinc chromateprimer to all other members. After completeassembly, apply one coat of zinc chromate primerto the surfaces of the ailerons, elevators,rudder, and attaching parts contacting dopedfabric, with the exception of alclad, 2S, 5S,and 52S. Apply two coats of ye I I OA/-green* lacquerto the cockpits, baggage compartment, attachingparts, and two coats of flat black lacquerto the landing light compartment. There areno further requirements for the bolting anglecover strip, fuselage rear of the baggage compartment,wings, and tail group. The carburetorair scoop shall not be finished on the interiorand no finish or buffinq is required for theseats. Buff the entire frame of 53S cockpite nc I OS u reSee footnotes, Paragraph 2.7. ENGINE COMPARTMENT SURFACES FA-23.Before assembly, do not anodize any surface.Do not apply primer to alclad, 2S, 35, and 52S,except between the surfaces of riveted skin buttor skin lap joints. Prime al I structures.After complete assembly, apply two coats of clearlacquer* to all surfaces with the exception ofa Ic lad, 2S, 35, and 525.8. EXTERIOR SURFACES,FA-25.PARTS, AND MEMBERSBefore assembly, do not anodize any part. Alclad,25, 35, and 525 wi II receive no primer nor furtherfinish, except where it forms a part of afinished member or surface or between the adjoiningsurfaces of riveted skin butt or skinlap joints. Also, the cockpit enclosure (535and corner cast i ngs ) shall receive no primer;however, all other surfaces should receive onecoat of zinc chromate primer. After completeassembly, apply two coats of flat bronze-greenNo. 9 antiglare. Finish entire frame of cockpitenclosure (535 and corner castings) by buffing.With the exception of alclad, 25, 35, and 525,apply two coats of clear lacquer* to cowling,fairing, fuselage, landing gear, tail group,wings, and attaching parts (including flaps, flapbays, and wheel wells). Only the rear surfacesof the propeller shal I two coats of flatblack lacquer.9. INSTRUMENT PANELS FA-28.Before assembly, do not anodize any structure.Apply one coat of zinc chronate primer to panels.After complete assembly, apply two coats of flatblack lacquer to the d i rect lighting portionand two coats of gloss white lacquer to the indirectlighting portion.10. ELECTRICAL AND RADIO JUNCTION BOXES ANDCONDUIT FA-29.Before assembly, do not anodize any structure.No finish is required, except for the boxes andconduit in cockpits and baggage compartmentwhich shall receive one coat of zinc chromateRESTRICTED[258]

RESTRICTED FINISH SPECIFICATIONFIGURE IITEMNO.

Before assembly, cadmium plate* all partsexcept brass and bronze parts in the enginecompartment, brass screws and nuts used in proximit.yto the compass, brass safety lockwire,and internal brass and bronze parts contact-h.ydrau lie fluid.i ngsurfaces.I ow-g.coilnFINISHSPECIFICATIONPAR. 10 TO 18RESTRICTEDprimer. After complete assembl.y, boxes andconduit in cockpits and baggage compartmentshall receive two coats of yellow-green* lacquer.No other members require finish.II.GENERAL FC-G.12. COPPER LINES FC-20.No finish requ i red13. DISSIMILAR METALS FC-23.Before assembly, cadmium plate* all partsand appl.y one coat of zinc chromate primer toall surfaces coming in contact with dissimilarmetals. After complete assembly, apply twofinish coats to match adjacent surfaces. Allowsurfaces to dry before joining.m. CHROMIUM PLATE FG-O.GENERALHard chromium plate all internal steel partscontacting hydraulic fluid and communicatingto exterior with the exception of corrosionresistant stee I15. DISSIMILAR METALS FG-5.Before assembly, apply one coat of zinc chromateprimsr to al I coming in contact with dissimilarmetals. After complete assembly, applytwo finish coats to match adjacent surfaces.16. STEPS FG-7.Before assembly, apply one coat of zinc chromateprimer to steps. After complete assembly,finish with two coats of clear lacquer*.*See footnotes, Paragraph 2.FINISHES17. WALKWAYS FG-8.After assembly, apply one coat of black Plyosynfcement and cover with 30-mesh carborundumgrits. Finish with one coat of blackPlyosyn cement as a sealer.18. COIL SPRINGS FG-IO.Before assembly, do not cadmium plate thecoil springs. With the exception of springssubmerged in hydraulic fluids which shall notbe primed nor otherwise finished, apply onecoat of zinc chrormte pririBr to al I springs.After complete assembly, apply two coats ofye I reen* lacquer to a II spr i ngs i thecockpits and baggage compartments, and applyone coat of zinc chromate primer to all coil*See footnotes, Paragraph 2."for equivalent.MANIFOLDPRESSURESTEAMHOT AIR DUCTSEXHAUST ANALYZERCOLD AIR DUCTSANTI-ICINGPURGINGVENT[ 1VACUUMSMOKE SCREEN EQUIPMENTHYDRAULIC OIL PRESSUREFUEL —VAPOR TOHEATERSI IAIR PRESSURE 20 PS I MAX.HEATEREXHAUST1^ ^ AIR PRESSURE 25 PSI MIN.FLOTATIONEQUIPMENTFigure 2~'Line Color Coding ChartRESTRICTED[ 260]

low-greenbett.flatflatJRESTRICTED FINISH SPECIFICATIONPAR. 18 TO 24springs in the fuselage aft of the baggage compartment,wing, and tail group.19. FLAT SPRINGS FG-IIBefore assembly, do not cadmium plate anyflat springs. With the exception of springssubmerged in hydraulic fluid, which shall notbe primed nor otherwise finished, apply onecoat zinc chromate primer to al I springs.After complete assembly, apply two coats ofye I lacquer to al Isprings inthe baggage compartment and cockpits.20. ARMAMENT PROTECTION FG-13.Ejection chutes, gun barrels, and blast tubesshal I weather-sealed by covering with a suitableeel lulose acetate tape.21. MARKINGS: LINES FG-21.The width of each color in the marking bandshall be approximately 1/2-inch. Cellulosetapes in appropriate colors coated with clearlacquer shall be used for the rrarking band. Thefollowing colors shall be used as the means ofidentification for lines. (See Figure 2.LINEAIR SPEEDPI TOT®STATIC®COMPRESSED AIR20 P.S. I. MAX.25 P.S. I. MIN.ANTI-ICER®EXHAUST ANALYZER®FIRE EXTINGUISHER^FLOTATION AND BILGE tFUEL IHYDRAULIC tMANIFOLD PRESSURE tOIL§OXYGEN tPURGING tSTEAMVACUUMLT. BLUE -BANDBLACKiLACK - LT. GREENLT. BLUE -

iFINISHSPECIFICATIONFIGURE 3RESTRICTEDLOWER SIDE OF RIGHT WINGRIGHT SIDE OF REAR FUSELAGE SECTIONFigure 3—Airplane Insignia LocatonsRESTRICTED[ 262]

Is).I partsRESTRICTED FINISH SPECIFICATIONPAR. 25 TO 3425. GENERAL FM-0.MAGNESIUM ANDBefore assembly, subject magnesium parts toprotective treatment as set forth in U.S. ArmyAir Forces Spec. 98-200! OA; E-3B. Appl.y onecoat of zinc chromate primer to a I andappl.y finish to match adjacent surfaces.26. EXTERIOR SURFACES, PARTS, AND hCMBERSFM-20.Before assembly, appl.y protective treatmentto all exterior surfaces, parts, and members,as indicated in U.S. Army Air Forces Spec. 98-200 lOA; E-3B; and appl.y one coat zinc chromateprimer. After complete assembly, appl.y an antiglaretreatment of two coats of flat bronzegreenlacquer No. 9- Appl.y two coats of clearlacquer* to the cowling, fairing, fuselage,landing gear, tail group, wings, and attachingparts (including flaps, flap bays, and wheelwe I On the SNJ-4, apply two coats of orange-yellowlacquer to the wing leading edgeand upper surfaces.MAGNESIUM ALLOYS27. ENGINE COMPARTMENT SURFACES, PARTS, ANDOPEN MEMBERS FM-23.Before assembly, apply protective treatmentto all surfaces, parts, and open members, asindicated in U.S. Army Air Forces Spec. 98-200I0A; E-3B; and apply one coat zinc chromateprimer. After complete assembly, apply twocoats of clear lacquer* to all surfaces.28. INTERIOR SURFACES, PARTS, AND MEMBERSFM-25.Before assembly, apply protective treatmentto all surfaces, parts, and members, as indicatedin U.S. Army Air Forces Spec. 98-2001 OA;E-3B; and apply one coat of zinc chromate prinier.After complete assembly, finish cockpits,landing light compartment, baggage compartment,and attaching parts with two coats of flatblack lacquer. To parts in view over cockpitenclosure trim, apply two coats of clear lacquer*.PLASTICS29. INTERIOR SURFACES, PARTS, AND MEMBERSFP-21.With the exception of parts that are notnormally visible, apply one finish coat tomatch adjacent surfaces.30. EXTERIOR SURFACES, PARTS, AND MEMBERSFP-23.Apply one coat of clear lacquer*.31. INTERIOR OF CLOSED MEMBERS FP-24.No finish required.STEEL32. GENERAL FS-0.Before assembly, cadmium plate* all steelparts except corrosion-resistant steel. Fillclosed members with hot (I09°C, I65°F) rawlinseed oil, drain thoroughly, and close withcadmiuoh-plated self-tapping screws. Apply onecoat of zinc chrormte primer to all parts.33. EXTERIOR SURFACES, PARTS, AND OPEN ^CM-BERS THAT CAN BE PLATED FS-20.Before assembly, cadmium plate* all exteriorsurfaces, parts, and open members, except thoseof corros ion- resistant steel. Apply one coatof zinc chromate primer. After complete assembly,apply an antiglare treatment of two'see footnotes, paragraph 2.coats of flat bronze-green lacquer No. 9. Withthe exception of corrosion-resistant steel,apply two coats of clear lacquer* to the cowling,fuselage, landing gear, tail group, wings,and attaching parts (including flaps, flapbays, and wheel wells). On t he SNJ-4 only,apply two coats of orange-yel low lacquer tothe wing leading edge and upper surfaces.3i|. EXTERIOR CLOSED MEMBERS THAT CAN BEPLATED FS-21.Before assembly, cadmium plate* all exteriorparts except corrosion-resistant steel. Applyone coat of zinc chromate primer. Withthe exception of corrosion-resistant steel,after complete assembly, apply two coats of'see footnotes, paragraph 2[ 263 ]RESTRICTED

low.,.I.beightRESTRICTEDFINISH SPECIFICATIONPAR. 34 TO i|3clear lacquer* to the cowling, fuselage, landinggear, tail group, wings, and attaching parts(including flaps, flap bays, and wheel wells).On the SNJ-4 only, apply two coats of orangeyeI lacquer to the wing leading edge andupper surfaces35. EXTERIOR CLOSED MEMBERS THAT CANNOT BEPLATED FS-22.Before assembly, fill with hot ( [09°C, I65°F)raw linseed oil, drain thoroughly, and closewith cadmium-plated self-tapping screws, t Ap -ply one coat of zinc chromate primer to a Iparts except corros ion-res istant steel. Withthe exception of corrosion-resistant steel,after complete assembly, apply two coats ofclear lacquer* to the cowling, fuselage, landinggear, tail group, wings, and attachingparts. On the SNJ-4 only, apply two coats oforange-yel low lacquer to the wing leading edgeand upper surfaces36. ENGINE COMPARTMENT SURFACES FS-23.Before assembly, cadmium plate* all steelparts except corros ion- res istant steel and themotor mounts. Closed members shall be filledwith hot ( f09°C, I65°F) raw linseed oil, drainedthoroughly, and closed with cadmium-platedself-tapping screws. t With the exception ofcadmium-plated parts, apply one coat of zincchromate primer to a I I steel parts. After completeassembly, apply two coats of clear lacquer*to a II parts that have not been cadmiumplatedwith the exception of corrosion-resistantsteel37. INTERIOR SURFACES, PARTS, AND OPEN MEM-BERS THAT CAN BE PLATED FS-26.Before assembly, cadmium plate* all steelparts except corrosion-resistant steel. Applyone coat of zinc chromate primer. After completeassembly, apply two coats of yellow-green*lacquer to the cockpits, baggage compartment,and attaching parts. The parts in view overthe cockpit enclosure trim shall receive twocoats of clear lacquer*. Apply two coats offlat black lacquer to the landing I cornpartment38. INTERIOR CLOSED MEMBERS THAT CANNOT BEPLATED FS-27.Before assembly, the interior closed membersshall be filled with hot (109°C, I65°F) rawlinseed oil, drained thoroughly, and closedwith cadmium-plated self-tapping screws. t Applyone coat zinc chromate primer to al I steelparts except corros ion- res istant steel. Aftercomplete assembly, apply two coats of ye I lewgreen*lacquer to the cockpits, baggage compartment,and attaching parts. The parts inview over the cockpit enclosure trim shall receivetwo coats of clear lacquer*. Apply twocoats of flat black lacquer to the landinglight compartment.39. INTERIOR CLOSED MEMBERS THAT CAN BEPLATED FS-28.Before assembly, cadmium plate* all interiorclosed members except corros ion- res istant steel.The i nte rior c losed members shall be filledwith hot (I09°C, I65°F) raw linseed oil, drainedthoroughly, and closed with cadmium-platedself-tapping screws. Apply one coat of zincchrorrate primer to a I I steel parts. After corrvpleteassembly, apply two coats of yel lowgreen*lacquer to the cockpits, baggage compartment,and attaching parts. The parts inview over the cockpit enclosure trim shal I receivetwo coats of clear lacquer*. Apply twocoats of flat black lacquer to the landinglight compartment.TEXTILES»I0.TANK PADS AND STRAPS FT-20.Before assembly, cover with Tolex 205 DC-FPcloth (Texti le Leather Corp. ) or equivalent.m. ELEVATORS AND AILERONS FT-21RUDDER FT-23.After complete assembly, apply two brushcoats and one spray coat of clear nitrate dopefol lowed by two spray coats of clear nitratedope containing a minimum of six ounces of aluminumpaste per gallon of spraying material.The weight of the coats shal Iand 5.0 oz./sq.yd-between 4.01*2. ANTENNA MAST FT-2^.After complete assembly, apply three coatsof clear dope and then apply two coats of aluminizeddope.i|3. COTTON WEBBING AND FABRIC FT-25.No treatment required for cotton webbing andfabric.*See footnotes. Paragraph 2.flash welded parts require no oiling.See footnotes, Paragraph 2.tplash welded parts require no oiling,RESTRICTED[264]

ller.iarityI I I surfacesTypeesTypeI andI I surfaces,I owIRESTRICTED FINISH SPECIFICATIONPAR. ^^ TO b\WOOD44. GENERAL REQUIREMENTS.Finish all wood surfaces in accordance withNA Spec. SP-I707-A. Prior to gluing or finishingany wood surface, thoroughly clean the surfaceto remove sawdust, grease, and any otherforeign matter from the wood. Wherever possible,accomplish all gluing prior to finishing, andremove all excess glue from v is ib le interioror exterior parts before applying any finish.Whenever the pecu I of the part is suchthat it is more conveniently glued subsequentto finishing, mask all areas to be glued withScotch Tape (or suitable equivalent). No adhesivemust be left on the wood when the tapeis removed. Apply doped sealing tape to allareas previously taped. Before attaching metalfittings, supports, etc., protect them fromdirect contact with the wood by applying twocoats of sealer, or one coat of sealer and onecoat of filler to the wood under the fitting.45. TYPES OF WOOD SURFACES.Type I- Interior of Closed Members, - Surfacesnormally remaining unseen, such as winginteriors, hoi low spars, the underside of floorboards,-etc.Type II - Interior Open Surfaces, Parts, andMembers. - Surfaces visible to occupants ofthe airplane, such as fuselage compartmentsand parts attaching.Type III - Exterior Surfaces, Parts, and Members.- Surfaces exposed to the weather and toview from the outside.46. DETAIL REQUIREMENTS.Type I- Interior of Closed Members FW-2t. -Apply two coats of sealer by dipping, the memberinto the sealer or by applying the sealerwith a bnjsh or spray gun. Al low each coat toair-dry one hour. The second coat of sealermay be aluminized with 8 ounces of aluminumpaste per gallon of unthinned sealer.Type li - Interior Open Surfaces, Parts, andMembers FW-21. - Apply one coat of sealer andone coat of f iThe coat of filler maybe replaced by a second coat of sealer on closegrainedwoods such as birch. Apply one finishcoat to match adjacent surfaces.Type III - Exterior Surfaces, Parts, and MembersFW-23. - Apply two coats of sealer, onecoat of filler, and one coat of surface r. Applyone finish coat to match adjacent surfaces.47. APPLICATION OF SEALER.After all gluing has been accomplished, applysealer, conforming to AAF Spec. 14113, to allType I, II, and III surfaces. Dip the surfacesinto sealer when possible or spray or brush onas practical. Do not thin sealer to less than25% nonvolatile content. Allow each coat ofsealer to air-dry at least one hour. To assurecomplete coverage of all areas and as a guidein sanding, it is permissible to add a suitabledye to the sealer.48. APPLICATION OF FILLER.I Iwith a brush, apply one coat of filler (WipeonFormula No. 1398 or equivalent) to al I Typeand wherever required on Typesurfaces. Allow to remain 2 to 5 minutes andwipe off across grain with excelsior or burlaprag. All Type II surfaces of open-grain woodsuch as mahogany and all exposed end-grain surfacesthat are not to be taped require one coatof filler over the regular sealer coat. Onal I I If i Mall naiholes, staple holes, and countersunk screwho I wi th filler. A I the filler coat toair-dry 2 to 3 hours as required.49. APPLICATION OF SURFACER.with a brush or with a spray gun, apply onecoat of surface r (conforming to AAF Spec. 141 15)to al I III surfaces and allow to air-dryfor 6 hours. With dry sandpaper, sand the surfacerafter it has thoroughly dried. Sand soas to produce the minimum possible thicknessof surface r coat but not deep enough to touchthe filler coat beneath.50. APPLICATION OF FINISH COATS TO TYPE IISURFACES.with a soft-bristled brush or with a spraygun, apply one finish coat to match adjacentsurfaces. Use Cellulose Nitrate Lacquer (Spec.AN-TT-L-51 ) or Aircraft Enamel (Spec. 3-98)as required. Apply a coat of sufficient thicknessto entirely cover the finishing materials.51. APPLICATION OF FINISH COATS TO TYPE IIISURFACES.with a soft-bristled brush or with a spray[265]RESTRICTED

I lerFINISH SPECIFICATIONPAR. 51 TO 54RESTRICTEDgun, apply one finish coat of clear AircraftEnamel (Spec. 3-98) pigmented with 16 ouncesof aluminum paste per gallon of unthinned enamel.52. FINISHING EXPOSED END-GRAIN SURFACES NOTSUITABLE FOR TAPING.Apply an extra coat of sealer in additionto the regular finish to all drain holes andend-grain surfaces where taping is not specified.When finishing at drain holes and other inaccessibleplaces, use a stiff wire coveredwith cloth or use an ordinary pipe cleaner.53. FINISHING EXPOSED END-GRAIN SURFACES SUIT-ABLE FOR TAPING.With a hand pinking machine or with pinkingshears, cut all sealing tapes required fromregulation glider fabric or equivalent thathas been predoped with one heavy brush coatof clear dope (Spec. AN-TT-D-51 4) . Cut thetape to such a size as to give a minimum lapat all points of 1/2-inch onto the straightgrain surface. Apply a second brush coat ofclear dope to the tape and immediately applythe tape to the exposed surface. Press thetape down firmly and eliminate all air pocketsbeneath the tape. Allow the tape to dry forone hour and apply an additional coat of cleardope if the nap of the material is not satisfactorilyfilled and sealed down.Note: Doped tape will not adhere satisfactorilyto finished areas. The finishshould be sanded off, sealer coated, and thetape doped on over the sealer.54. SANDING SEALER, FILLER, AND SURFACER COATS.with dry fine sandpaper, lightly sand thesealer and f i coats on Type II and Type IIIsurfaces. After the surfacer coat has completelydried on all Type III surfaces, sandthe coat so as to leave the minimum possiblethickness of surfacer coat, but do not sand deepenough to touch or bare the filler coat beneaththe surfacer.RESTRICTED[ 266 J

RESTRICTED HEAT-TREATED PARTSPAR. TO I2SECTION 12HEAT-TREATEDPARTSI. GENERAL.The following list of heat-treated parts includesonly heat-treated castings, forgings,and steel members used in the construction ofthe airplane. Generally, the repair of these partsis not recommended and any extent of damagenecessitates complete replacement of the damagedmember. Hcwever, it is to be noted that in thePRINELLDIAMETER OFIM-

crankHEAT-TREATEDPARAGRAPH 3PARTSRESTRICTED3. HEAT-TREATED PARTS LISTPart No. Title MaterialCodeHeat Treat toTensi le Strengthin P.S. I . or asNotedIBI226 Eye-bolt - Ball Bearing OJ\-FBII90 Ring - Ai rp lane Mooring CM-F66-13013 Angle - \Mng Centersect ion Upper Bolting AA-E66-13104 Angle - Wing Centersect ion Front Bolting AA-E77-13303 Angle - Wing Centersect ion Fuel Tank Cover Bolting AA-E55-14027 Angle - Outer Wing Upper Bolting AA-E54^14028 Angle - Outer Wing Nose Bolting AA-E88-14029 Angle - Outer Wing Lower Front Bolting AA-E54-14030 Angle - Outer Wing Lower Rear Bolting AA-E55-14065 Bracket - Aileron Outboard Hinge AA-F ( I7ST)55-14066 Bracket - Aileron Inboard Hinge AA-F ( I4ST)55-14069 Bracket - Aileron Center Hinge AA-F ( I4ST)55-14244 Support - Aileron Bel I Bearing AA-F ( I7ST)55-16080 Bracket - Aileron Inboard Hinge AA-F ( I4ST)55-16081 Fitting - Aileron Center Hinge AA-F ( I7ST)55-18022 Hinge Fitting - Flap A/WF (A5IST)55-18024 Eye-bolt - Wing Flap CM-B52-21018 Fitting - Elevator Center Hinge Support AA-F52-21019 Fitting - Elevator Center Hinge Support AA-F55-21020 Fitting - Stabilizer Attaching AA-F (A5IST)52-21026 Bracket - Elevator Inboard Hinge Support AA^F55-22037 Hinge Fitting - Trim Tab AA-F ( I7ST)52-23012 Fitting - Rudder Upper Hinge Support AA-F52-23013 Fitting - Rudder Center Hinge Support AA-F52-23014 Fitting - Rudder Lower Hinge Support AA-F28-31053 Fastener - Engine Ring Cowling Rear AA-S66-31058-2 Catch - Fire Extinguisher Door Release Alv^S66-310102 Hook - Engine Ring Cowling AAr-E57-31103-2 Tube - Fuselage Station No. Lower Cross CM-T19-31105 Fitting - Fuselage Connection CM-F55-3)174 Fitting - Fuselage Connection CM-B55-31220 Step - Cockpit MA-C19-31231 Bracket - Fuselage Pilot's Seat Support CM-F5^31383 Guide - Fuselage Front Step Tube CM-B55-31385 Spring - Fuselage Front Step Tube Keeper MAf49-31894 Finger - Enclosure Release Handle CM-S25-31902 Fitting - Engine Mount Attaching QvkF66-33109 Pin - L.G. Retract Arm Attach. CM-B36-33112 Pin - L.G. Pivot Qv^T36-33117 Bushing - L.G. Support CM-B36-33119 Bushing - L.G. Attach. Fitting CM-B3^33120 Nut - L.G. Pivot Pin CM-B19-33447 Wrench - Special CM-S55-33467 Spring - Brake Cylinder Ext. MW36-33515 Latch - L.G. Lock Cam CM-S55-33525 Pin - Landing Gear Lock CM-B44-33534^2 Fitting - L.G. Lock and Hoist CIv^F125-145,000125-140,00058,00058,00058,00058,00058,00058,00058,00055,00065, 00065,00055,00065, 00055,00043,000125,00065,00065,00043,00065,00055, 00065,00065,00065,00062,000125,00062,000Normal izeNormal Ize125-140,00032,000Normal ize125,000( 10 Minutes)160-180,000Norma I i ze145-160,000160-180,000160-180,000125-140,000125,000180-200,000( 20 Minutes)125,000160-180,000160-180,000* - Heat treat spring at 500-550° F for period noted.RESTRICTED[268]

RESTRICTEDHEAT-TREATED PARTSPARAGRAPH 33. HEAT-TREATED PARTS LIST (CONTD.)Part No. Title44^33534^3 Fitting - L.G. Lock and Hoist66-33536 Bolt - L.G. Lock and Hoist Fitting Attach.36-33541 Rod - L.G. Lockpin Push36-33542 Bolt - L.G. Mechanism55-33544 Be II crank Assemoly - L.G. Locking Mechanism66-42012 Shaft - Carburetor Mixing Chamber51-47032 Stud - Oil Tank Sump49-51047 Support - Oil Separator Outer55-521 19 Spring - Surface Lock55-52302 Bellcrank - Ai leron55-52316 Clevis - Aileron Control Rod55-52428-2 Rod - Rudder Brake Pedal19-52440 Spring - Tension Coil49-52441 Spring - Rear Cockpit Rudder Pedal Adjust.55-52610 Arm - Flap Indicator55-52649-2 Rod - Flap Connecting55-52653 Plunger - Flap Control Handle Lock Spring58-53004^3 Tube - Pilot's Cockpit Seat Support49-53060 Clamp - Alemite Grease Gun55-53061 Pin - Observer's Seat Latch49-53073 Plunger - Seat Adjusting55-53075 Tube - Observer's Seat Adjusting55-53076 Brace Assembly - Observer's Seat55-53096 Spring - Observer's Seat Latch55-53104 Axle - Observer's Seat36-54029 Bearing - L.G. Position Indicator Switch Box36-54033 Guide - L.G. Position Indicator Swltcn Box55-54083 Bracket - Ignition Switch Hand Support55-54125 Spring - L.G, Position Indicator Switch36-55009 Fitting Assembly - ^ing Center Section Hoist36-58008-2 Fitting - L.G. Hy d rau I i c St rut88-58101 Body - Hydraulic L.G. & Flap Control Unit55-58137 Spring - Hydraulic One-Way Valve55-58143 Spring - Hydraulic L.G. & Flap Relief Outer55-58144 Spring - Hydraulic L.G. & Flap Rej ief Inner55-58152 Spring - Control Unit One-Way Valve55-58196 Spring - Pressure Control Return55-61040 Screw- Ammunition Box & Ejection Chute Adjust.59-73006 Bow- Instrument Flying Hood59-73007 Bow- Instrument Flying Hood

REMENTS,RESTRICTEDNDEXALPHABETICALINDEXACCESSFUEL TANK, 198FUEL TANK VALVE, 195OUTER WING LEADING EDGE, 1U7ACCESSORIESWOODEN, 250AGE HARDENINGALUMINUM, 19AILERONSAREA, 2, 164CABLES, C RAN KS, AND HORNMARKING S, 261164CONSTRUCT ION ,COVERING MAT ERIALS, 193FABRIC CO VER ING REQUIRE-MENTS,FABRIC,182F INISHES, Nl 264FINISH, 2 57GENERAL R EPA IR, 169LEADING E DGE SKIN REPAIR, 175MOVEMENT, 16 4NOSE RIB REP LACEMENT, 173ORIGINAL FAB RIC COVERING. 164SPAR OUTE R W ING. 119SPAR SPLI CE, 171TAB AREA, 16 4TRAILING EDG E, 172TRAILING EOG E SKIN REPLACE-MENT, 1 71TRAVEL, 2TRIM TAB CABLES, CRANKS, ANDHORN MA RKINGS, 261AIRFOIL, WING, 2AIR FURNACE, ALUMINUM ALLOY,H.T., 19AIRPLANECONSTRUCTION, 1DATA, 2OmENSIONS. 3INSIGNIA, 262MARKINGS, 257PART NUMBERS, 4SPECIFICATIONS, 2AIR PRESSURE20 P.S. I. , MAXIMUM LI NES,COLOR CODING, 26025 P.S. I. , MINIMUM LINES,COLOR CODING, 260AIR PRESSURE TESTENGINE MOUNT, 51FUEL TANK, 199FUSELAGE TRUSS, 51OIL COOLER, 202OIL TANK, 199AIRSPEEDPITOT LINE, COLOR CODE, 260STATIC LINE, COLOR CODE, 260ALCLAD24S0, 1824ST, 19ANNEALING, 20BENDING, 20REPAIR MATERIAL FOR FIXEDSURFACES, 158ALCOA, EXTRUSION EQUIVALENTS,21, 70, 158ALIGNMENTENGINE MOUNT, 41, 51FUSELAGE TRUSS, 41, 51ALLOY2S, 173S, 1714ST, 1717ST, 1724ST, 1752S, 1753ST, 1753SW, 17ALUMINUM, SPEC. EOUIV.. 32BRITISH REPLACEMENT, 32BRITISH SPEC. , 32COLD WORKING, 17COPPER, SPEC. EOUIV., 33HEAT-TREATABLE, 17NONHEAT-TREATABLE, 17STEEL SPEC. EOUIV., 32TENSILE STRENGTH, 32U.S.A. SPEC, 32ALUMINUMALLOY SPEC. EOUIV. , 32ANNEALING, IBANODIZING, 258CASTINGS, 12, 249EXTRUDED REPAIR MATERIAL, 158EXTRUSIONS, 18FORGINGS, 249HEAT-TREATING, 19WROUGHT STOCK, 17ALUMINUM FINISHCONDUIT, 258ELECTRICAL JUNCTION BOXES,258ENGI NE COMPART. SURF. , 258EXTERIOR SURFACES, fARTS,AND MEMBERS, 258GENERAL, 258INSTRUMENT PANELS, 258INTERIOR CLOSED MEMBERS, 258INTERIOR SURFACES, PARTS,AND OPEN MEMBERS, 258RADIO JUNCTION BOXES, 258ALUMINUM SHEET2S-i/2H, 173S-1/2H, 1752S-1/2H, 182S0, 173S0, 1724S0 ALCLAD, 1852S0, 1724ST ALCLAD, 19MARKINGS, 12PRIMING, 22REPAIR MATERIAL FOR FIXEDSURFACES, 158REPAIR MATERIALS FOR FUSE-LAGE, 70-71REPAIR MATERIALS FOR MOVABLESURFACES, 191ALUMINUM TUBE52S0, 1752S0 SPLICING, 229ANNEALINGALCLAD, 20STEEL TUBES, 36TEMPERATURES, 20TUBE 52SO, 18ANODIZING ALUMINUM, 258[i]ANTENNA MASTFINISH, 264HEIGHT, 2ANTI-CORROSION FINISH, 22, 257ANTI-GLARE FINISH, 257ANTI-ICING LINE COLOR CODE, 260ARC WELDINGAMPERAGES, 36ELECTRODE REOU I 36EOUIPMENT, 36PRECAUTIONS, 36AREAAILERONS, 2, 164BALANCE, 2BOOSTER TAB, 2ELEVATOR, 2. 164P IA p C 9HORIZONTAL STABILIZER, 2RUDDER, 2. 161TRIM TAB. 2VERTICAL STABILIZER, 2WING, 2ARMAMENT, INSTALLATION REPAIRS,253ASSEMBLY PART NUMBERS, 4AXLES, LANDING GEAR, 206BBALANCEAILERON, 186. 188AREA, 2CONTROL SURFACES, 186CORRECTION OF CONTROL SURFACESTATIC, 188ELEVATOR, 186, 188INNER TUBE, 209RUDDER. 186. 188TIRE, 209BANDINGCONTROL CABLE, 222LINE, 260BARS, BUCKING, 7BEADSFIXED SURFACE Rl B, 152TUBING, 229BENDINGALCLAD, 20Ml NIMUM RADIUS, 20PLYWOOD SKI N, 85BLAST TUBES, PROTECTION, 261BLIND RIVET HOLESCHERRY, 235LOCATING, 140LOCATING BY IMPROVISEDMETHOD, 143BLOCKS, FORMING, 25BOLTSAN4 SPECIFICATIONS, 13EDGE DISTANCE, 9FITTING OF, 11RESTRICTED

INDEXRESTRICTEDSTANDARD HOLES FOR, 12THREADS IN BEARING, 11TORQUE MEASUREMENT, 11TYPES, 12BOLTING ANGLESCOVERS, 157WING JOINT, 155BOMB RACKS, REPAIR OF, 253BONDINGBRAIDS, 227FLEXIBLE CONDUIT, 225METAL PARTS, 225REQUIREMENTS, 225RIGID CONDUIT, 225BOOSTER TABCONSTRUCTION, 166TRAVEL, 2BOXESJUNCTION, 223JUNCTION FINISH, 258BRAIDS, BONDING, 227BRAKE, HAND, 23BRINELL HARDNESS, 267BRITISH, SPEC. EQUIVALENTS,32BUCKING BARS, 7BULKHEADPART NUMBERS, ALUMINUM REARFUSELAGE, 59PLYWOOD WEBS, WOODEN REARFUSELAGE, 86WEBS, ALUMINUM, 68BURRS, HOLE, 5BUSHINGS"B" STANDARD, 255"B" STANDARD CODING, 255CABLE, 216COMPONENT PART, 256LOCATIONS OF, 25UPART NUMBERS, 254REAM NOTES, 254REDUCER, 254SPACER, 254CABLESBUSHINGS FOR, 216COLOR BANDING, 222CONTROL, 216CUTTING OF, 216DIAMETERS OF, 216FABRICATION OF, 216MARKINGS, 261NEGLIGIBLE DAMAGE TO, 216REPAIR OF, 216REPLACEMENT OF, 216RUST PREVENTION, 216SPARE, 216SPECIFICATIONS, 216TERMINALS FOR, 216TESTING OF, 222THIMBLES for; 216TYPES OF, 216WRAPPING OF, 221CADMIUM PLATINGCOIL SPRINGS, 260COPPER, 260FLAT SPRINGS, 261STEEL, 263CAPCENTERSECTION FRONT SPARLOWER. 126CENTERSECTION FRONT SPARUPPER, 124CASEIN GLUE, 75APPLICATION, 75MIXING, 75CASTINGSALUMINUM. 12, 249FINISH, 257REPAIR OF, 249WELDING OF, 249CEMENTAIRSHIP RUBBER, 215CAVALPRENE, 196DOLPHIN NO. 1625, 198LUCITE. 246TIRE BALANCING, 210CENTER PUNCH MARKS, 5CENTERSECTIONBOLTING ANGLE PART NUM-BERS, 156FLAP SPAR, 119FRONT SPAR LOWER CAP, 126FRONT SPAR UPPER CAP, 124FRONT SPAR WEB, 128-130INTERMEDIATE RIBS, 150LEADING EDGE RJBS, 151REAR SPAR, 122CHERRY RIVETSDIMPLING PREPARATIONS, 235DRILLING PREPARATIONS, 235EXPANDING, 236GRIP LENGTH OF, 235GUNS FOR, 236INSERTION OF, 235TRIMMING, 238TYPES OF, 235USE OF GUNS FOR, 236CHORDFLAPS, 2. 161MEAN AERODYNAMIC, 2STABILIZER, HORIZONTAL, 2WING, 2CHROMIUM PLATING, 260CLAMPS"C", 6PRERIVETING, 6SKI N, 6WOOD, 77CLEANINGOIL COOLER, 202TANKS, 195TIRE INTERIOR, 210TRANSPARENT PLASTIC PANEL,224CLEARANCE, PROPELLER, 2CLECOSCOLOR CODING, 7USE OF, 7CLOSING STRIP, OUTER WING, 147COCKPITFINISH, 257WOODEN SEATS, 250COIL SPRINGSCADMIUM PLATING, 260FIN^SH, 260COLD AIR DUCT LINES, COLORCODING, 260COLD WORKING ALLOYS, 17COLOR CODINGCABLES, 222CLECOS, 7LINES, 260COMPARTMENT, LANDING LIGHTFINISH, 257CONDUITBONDING, 225FINISH REQUIREMENTS, 258INSTALLATION DRAWING NOS. , 225REPAIR OF, 232SUPPORT OF, 225CONSTRUCTIONAIRPLANE, 1FUSELAGE, 34, 60, 71LANDING GEAR, 206MOVABLE SURFACES, 161OIL COOLER, 202TANKS, 194WING, 91CONTOURS, WING LEAD. EDGE, 146CONTROL CABLESSPLICING, 220TERMINAL SOLDERING, 220TERMINAL SWAGING, 218CONTROL RODSEND FITTING REPLACEMENT, 252FABRICATION OF, 252NEGLIGIBLE DAMAGE TO, 251SPLICING OF, 251CONTROL SURFACESFABRIC PATCHING, 176PART NUMBERS, 163, 165, 167RECOVERING, 183STATIC BALANCE, 186STATIC UNBALANCE CORRECTION,188UNBALANCED DETERMINATION, 187COOLERS, OILCLEANING, 202INSTALLATION, 196REPAIR OF, 202REPLACEMENT TUBES, 203SUPPORTS, 196TESTING, 202COPPERALLOY, SPEC. EQUIV., 33FINISH, GENERAL, 260FINISH, IN CONTACT WITHDISSIMILAR METALS, 260RESTRICTED[n]

RESTRICTEDINDEXLINES, FINISH, 260COREOIL COOLER, LEAKS, 203REMOVING OIL COOLER, 203CORROSION PROTECTIONALUMINUM, 22, 258CABLES, 216DISSIMILAR METALS, 260STEEL TUBES, 51COTTON WEBBING, FINISH, 264COUNTERSINKINGCUT, 6DIAMETERS, 6DIMENSIONS, 6DIMPLE, 6COVER FORMERS, FUEL TANKCOMPARTMENT, 149COVERINGMATERIALS, AILERON, 193MATERIALS, ELEVATOR, 192MATERIALS, RUDDER, 191REMOVAL OF FABRIC, 179REQUIREMENTS, AILERON FAB-RIC, 182REQUIREMENTS, RUD. & ELEV.FABRIC, 180, 181COVERS, WING JOINT, 157COWLINGCONSTRUCTION, 55PART NUMBERS, 55REPAIR, 55CRACKSALCLAD SKIN, 138CASTING SURFACE, 249LOCATING FUEL TANK, 195LOCATING OIL TANK, 195METAL SURFACE, 22REPAIR OF LARGE TANK, 199SEALING SMALL FUEL TANK, 197SEALING SMALL OIL TANK, 197STEEL TUBES, 43TRANSPARENT PLASTIC PANEL, 245CRANK MARKINGS, 261CRATER ELIMINATOR, WELDING, 37CUTOUTS, FIXED SURFACE RIBS, 152DAMAGEEXTENT OF, 22PRIMARY, 22SECONDARY, 22DECIMAL EQUIV. , DRILL SIZES, 5DEGREES MOVEMENTAILERON, 2. 164ELEVATOR, 2. 164LANDING FLAP, 2. 161RUDDER, 2. 161TRIM AND BOOSTER TABS, 2DENTS, SKIN SURFACE, GENERAL, 22DIAMETERSCABLE, 216COUNTERSINK, 6PROPELLER, 2RIVET CODING, 1STEEL TUBES, 35DIEALCOA EQUIVALENT NUMBERS,21, 70, 158SWAGED CABLE FITTINGS, 216DIHEDRAL, WING, 2DIMENSIONSAIRPLANE, 3BUSHING REAM, 255RIVET, 1RIVET COUNTERSINKI NG, 6DIMPLE COUNTERSINKING, 6DISCONNECT POINT MARKINGS,ENGINE, 261DISSIMILAR METALS,PROTECTION, 260DISTANCESBOLT EDGE, 9RIVET EDGE, 9CORROSION-DOPING FABRICMOVABLE SURF., 185, 186, 264WOOD END GRAIN FINISH, 89DRAWING NUMBERS,BLIES, 4DRILLCHART, 5RIGHT-ANGLED, 5SIZES, 5DRILLING, 1DZUS FASTENERS, 242COWLING, 55HEAVY-DUTY, 242INSERTING NEW, 244LIGHT-DUTY, 242REMOVAL- OF, 39REPLACEMENT OF, 242MAJOR ASSEM-EDGE DISTANCE, RIVET AND BOLT, 9EJECTION CHUTES, PROTECTION, 261ELECTRIC ARC WELDING, 36ELECTRIC SYSTEMGENERAL, 223NA DWG. NOS., 223REPAIR TOOLS AND MATERIALS, 227ELEVATORAREA, 2. 164CABLES, CRANKS, AND HORNMARKINGS, 261CONSTRUCTION, 163COVERING MATERIALS, 192COVERING REQUIREMENTS, 181FINISHES, GENERAL, 257, 264GENERAL REPAIR, 169LEADING EDGE SKIN, 170MOVEMENT, 2. 164ORIGINAL FABRIC COVERING, 164RIB REMOVAL, 169RIB REPLACEMENT, 170SPAN. 2TAB AREA, 164TRAILING EDGE STRIP REPLACE-MENT, 171.TRAVEL, 2, 164EMPENNAGE(see elevator, rudder, andstabilizers)ENGINEcompartment SURFACES, FINISHREQUIREMENTS, 258COMPARTMENT SURFACES, STEELFINISHES, 264COWLING construction, 55COWLI NG PART NUMBERS, 55COWLING REPAIR, 57DISCONNECT POINT MARKINGS, 261FIREWALL CONSTRUCTION, 52STEEL TRUSS, 3'^EQUIVALENTSALUMINUM ALLOY, 32BUSHINGS, 254COPPER ALLOY, 33DECIMAL, 5EXTRUSION ALCOA, 21. 70. 158STEEL ALLOY, 32WIRE AND CABLE, 33EXHAUSTANALYZER LINES COLORCODING, 260MANIFOLD COLLAR REPLACE-MENT, 52MANIFOLD CONSTRUCTION, 51MANIFOLD REASSEMBLY, 52MANIFOLD REPAIR, 52EXPLOSIVE RIVETEXPANDING IRONS, 239HANDLING PRECAUTIONS, 238INSTALLATION, 239RIVETING PRECAUTIONS, 238STORAGE PRECAUTIONS, 238EXTRUDED REPAIR MATERIALFIXED SURi^ACES, 158FUSELAGE, 68EXTRUSIONSALUMINUM, 18EQUIVALENTS, 21. 70. 158FABRICFLEXIBILITY, 175LARGE TEARS AND HOLES, 177PARTIAL RE-COVERING. 178PATCHING, GENERAL, 175REPAIR TOOLS, 189SECTION REPLACEMENT. 177TEARS, SMALL, 176TEARS, V-SHAPED, 176WEBBING FINISHES, 264FABRIC COVERING MATERIALSAILERON, 193ELEVATOR, 192RUDDER, 191[iii ]RESTRICTED

INDEXRESTRICTEDFABRIC COVERING REMOVAL, 179FABRIC COVERING REQUIREMENTSAILERON, 182ELEVATOR, 181RUDDER, 180FABRIC DOPING, MOVABLE SUR-FACES, 185, 186, 264FABRIC ENVELOPEATTACHMENT, 183HAND SEWI NG, 18UMACHINE SEWING, 184PREPARATION, 183FAIRINGSFINISH, 257FUSELAGE, 68FASTENERSDZUS, 242PRERIVETING, 6SKIN, 6FILLETS, FUSELAGE, 68FINISHESALUMINUM, 258COPPER, 260MATERIAL SPECIFICATION, 257SPECIFICATION, 257STEEL, 263STEEL TUBE, 51TEXTILES, 264WOODEN STRUCTURES, 88, 265FIRE EXTINGUISHER LINES, COLORCODING, 260FIREWALLCONSTRUCTION, 52COWLING ATTACHING ANGLEREPAIR, 54STIFFENER REPAIR, 54WEB REPAIR, 53FIVE- TUCK CABLE SPLICE, 220FIXED SURFACESGENERAL SKIN REPAIR, 136GENERAL SPARLARGE SKI N HOREPLACEMENTSKI N CRACKS,SKIN PANEL RESMALL SKIN HOSPLICING SKI NSTRINGER COMBC366T, 107STRINGER REPASTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGER TYPESTRINGERS, GESTRINGERS, TYC204T, 205STRINGERS, TYC250T, 106STRINGERS, DOUBLED C366T, 111STRINGERS, DOUBLED K7 7 A, 112SUMMATION OF STRINGER TYPESUSED, 103FLANGES, FIXED SURFACE RIB, 153FLAPAREA, 2CHANNEL SPAR, 167CHORD, 2CONSTRUCTION, 161FINISH, 257LEADING EDGE, 168SKIN REPAIR, 166TRAILING EDGE, 168FLAP REPAIR, GENERAL, 166FLAP SPARCENTERSECTION, 119OUTER WING, 119FLAPPER VALVES, FUEL TANK, 195FLAT SPRINGS, CADMIUM PLATING,261FLOOR BOARDS, WOODEN, 250FLOTATION LINES, COLOR CODING,260FORGINGSALUMINUM, 249FINISH, 257WELDING OF, 249FORMERSALUMINUM REAR FUSELAGE, 61FRONT FUS. SIDE PANEL, 57, 60FUEL TANK COMPART. COVER, 149PART NUMBERS, ALUMINUM REARFUSELAGE, 59PART NUMBERS, FRONT FUSELAGESIDE PANELS, 57WOODEN REAR FUSELAGE, 86FORMINGFRAMES, 62

RESTRICTEDINDEXGUN BARRELS, PROTECTION, 261HAMMERS, 8HHEATER LINES, COLOR CODING, 260HEAT-TREATABLE ALLOYS, 17HEAT- TREATED PARTSHARDNESS TABLE, 267MATERIAL CODE, 267TENSILE STRENGTH, 267HEAT-TREATINGAGE HARDENING, 20AIR FURNACE METHOD, 19ALUMINUM, 19QUENCHING, 19SALT BATH METHOD, 19TEMPERATURES, 19TIMING, 19HOLESLOCATING BLIND RIVET, 140OIL COOLER SHELL, 2O3REPAIR OF LARGE TANK, 199REPAIR OF TRANSPARENT PLASTICPANEL, 246RIVET, 5HORIZONTAL STABILIZERALUMINUM FRONT SPAR, I32ALUMINUM REAR SPAR, 135FINISH, 257GENERAL SKIN REPAIR, I36LARGE SKIN HOLES, I38-I39PART NUMBERS, 91SKIN ARRANGEMENT, I38SKIN CRACKS, I38SKIN PANEL REPLACEMENT, 148SMALL SKIN HOLE REPAIR, 137SPLICING SKIN PANELS, 141TIP SKIN DENTS IN, 149WOODEN CONSTRUCTION, 91WOODEN REPAIR, I32HORNSCONTROL FINISH, 257MARKINGS, 261HOT-AIR DUCTS, COLOR CODING, 260HYDRAULICLINE REPAIR, 23OOIL LINES COLOR CODING, 260OIL TANK, 199HYDROGEN WELDING, 197INCH-POUNDS TWIST, 11INCIDENCESTABILIZER, 2WING, 2INSIGNIA LOCATIONS, 262INSTRUMENT PANELSFINISH, 257FINISH REQUIREMENTS, 258INSULATION, ELECTRIC WIRE, 224INTERCOSTAL. WOODEN REAR FUSELAGEBAGGAGE COMPARTMENT, 79INTERCOSTAL ANGLE, REAR ALUMINUMFUSELAGE, 64INTERMEDIATE, RIBS C SECTION, 150JACK, HYDRAULIC, 207JIG, FUEL TANK TESTING, 197JOINT BOLTING ANGLES, WING, 155JOINT COVERS, WING, 157JOINTSFORMING BEADED TYPE TUBE, 229FORMING FLARED TYPE TUBE, 229TESTING STEEL, 51TYPES OF TUBE, 229JUNCTION BOXES, FINISH REQUIRE-MENTS, 2 58LACINGMARLIN HITCH FOR WIRE, 223WIRE, 223LANDING FLAPSCHANNEL SPAR, 167CHORD, 2, 161CONSTRUCTION, 161LEADING EDGE, 168MOVEMENT, 2, 161REPAIR, GENERAL, 166SKIN REPAIR, 166TRAILING EDGE, 168LANDING GEARCONSTRUCTION, 206FINISH, 257MAIN, 206REMOVAL OF WHEELS, 207SPECIFICATIONS, 2TAIL WHEEL, 206TIRES, 207TIRE DEFECTS, 207TREAD, 2TUBES, 207LANDING LIGHT COMP. FINISH, 257LEADING EDGEACCESS TO OUTER WING, 147AILERON SKIN REPAIR, 175CENTERSECTION RIBS, 151CONTOURS, OUTER WING, 146ELEVATOR AND RUDDER SKIN, 170LANDING FLAP, 168PART NUMBERS, FLAP, 162WING SKIN, 143LEAKSFUEL AND OIL TANKS, 195LOCATI NG OIL COOLER, 202OIL COOLER CORE, 203OIL COOLER SEAM, 203REPAI R OF OIL COOLER, 203LINESCOLOR CODING, 260COPPER, FINISH, 260MISCELLANEOUS, 232LINSEED OIL, APPLICATION TO STEELTUBES, 51SPECIFICATION, 263LOCATING BLIND RIVET HOLESIMPROVISED METHOD, 143WITH TOOL, 140LONGERONSALUMINUM REAR FUS., LOWER, 66ALUMINUM REAR FUS.^ PART NUMBERS,59ALUMINUM REAR FUS., UPPER, 64WOODEN REAR FUS., UPPER, 80WOODEN REAR FUS., LOWER, 82LOWER CAP, CENTERSECTION LOWERSPAR, 126LOWER LONGERONSALUMINUM REAR FUSELAGE, 66WOODEN REAR FUSELAGE, 82MMAGNESIUM FINISHENGINE COMPARTMENT SURFACES,PARTS, AND OPEN MEMBERS, 263EXTERIOR SURFACES, PARTS, ANDMEMBERS, 263GENERAL, 263INTERIOR SURFACES, PARTS, ANDMEMBERS, 263MAIN OUTER WING SPAR, 114-116MANIFOLDCOLLAR REPLACEMENT, 52CONSTRUCTION, 51LINES, COLOR CODING, 260REASSEMBLY, 52REPAIR, 52MARKINGSAIRPLANE, 257ALUMINUM SHEET, 12ENGINE DISCONNECT POINT, 261INNER TUBE, 209MISCELLANEOUS AIRPLANE, 261TIRE, 209MARKS, CENTER PUNCH, 5MAST, ANTENNA FINISH, 264MATERIALSAILERON FABRIC COVERING, 193ALCLAD SHEET FUS. REPAIR, 70ALCLAD SHEET METAL REPAIR,FIXED SURFACES, 158ALUMINUM EXTRUDED REPAIR, 158ELECTRICAL REPAIR, 227ELEVATOR FABRIC COVERING, 142EXTRUDED ALUMINUM FUSELAGE RE-PAIR, 68OIL COOLER REPAI R, 205RUDDER FABRIC COVERING, 191STEEL FUS. REPAIR, 54LORESTRICTED

INDEXRESTHICTEDTANK REPAIR, 199TIRE REPAIR, 215TUBING REPAIR, 233WOOD FINISH, 88WOOD FUSELAGE REPAIR, 88MOISTURE CONTROL,FUSELAGE, 73REAR WOODENMOVABLE SURFACES, 1 61Al LERON FABRIC CO VERING RE-OUIREMENTS, 182ATTACHING FABRIC ENVELOPECOVER, 183DOPING FABRIC, 18 5-186ELEVATOR FABRIC C OVERING RE-OUIREMENTS, 181FABRIC FLEXIBILIT Y, 175FABRIC REPAIR TOO LS, 189FABRIC SECT. REPL ACEMENT, 177GENERAL FABRIC PA TCHING, 175HAND SEWING ENVEL OPE, 184LARGE FABRIC TEAR S, 177MACH. SEWING FAB. ENVELOPE, 18uMETAL STRUCTURE R EPAIR MATERIAL191PARTIAL FAB. RE-C OVERING, 178REPAIRING FABRIC ENVELOPE, 183REMOVAL OF FABRIC COVERING, 179RUDDER FABRIC COV ERING REQUIRE-MENTS, 180SMALL FABRIC TEAR S, 176V-SHAPED FABRIC T EARS, 176MOVEMENTAILERON, 2, 164ELEVATOR, 2, 164LANDING FLAP, 2, 161RUDDER, 2, 161TRIM AND BOOSTER TABS, 2NNOSE RIB REPLACEMENTAILERON, 173WING, 155NUMBERING, ELECTRICAL WIRE, 223NUMBERSMAJOR ASSEMBLY, 4STATION, 34, 100NUTSBASKET, ANCHOR, 12CORNER, ANCHOR, 12DIMENSIONS OF, 16GANG CHANNEL, 12PART NUMBERS, 16SELF-LOCKING, 12SIZE, 16THREADS, 16TYPES OF, 12, 16OILCOOLERS, 202LINE, COLOR CODING, 260LINSEED, 51, 263SYSTEM, 194OUTER WINGAILERON SPAR, 119BOLTING ANGLE PARTCLOSING STRIP, 1U7NOS., 156FLAP SPAR, 119LEADING EDGE, ACCESS, 147LEADING EDGE CONTOURS, 146MAIN SPAR, 114-116OXYACETYLENE WELDING, STEEL, 37OXYGENDISTRIBUTION LINE, COLOR COD-ING, 260FILLER LINE COLOR CODING, 260PANELSCLEANING PLASTIC, 244COCKPIT ROOF REPLACEMENT, 248DEEP SCRATCHES ON TRANSPARENTPLASTIC, 244FRONT FUSELAGE SIDE PARTNUMBERS, 57S, GENERAL REPAIR OF TRANSPARENTPLASTIC, 245MINOR SCRATCHES ON TRANSPARENTPLASTIC, 244PREPARATION OF NEW TRANSPARENTPLASTIC, 246ROUTING EDGES OF TRANSPARENTPLASTIC, 247ROUTING INSIDE CURVES OF TRANS-PARENT PLASTIC, 247SIDE, REPLACEMENT, 248TRANSPARENT PLASTIC, 244PART NUMBERSAILERONS, 164AIRPLANE, 4ALUMINUM REAR FUSELAGE, 60BUSHINGS, 254C'SECTION BOLTING ANGLE, 156ELEVATOR, 163ENGINE COWLI NG, 55FIXED SURFACE Rl BS, 91FIXED SURFACE SPARS, 91FRONT FUSELAGE SIDE PANELS,HORIZONTAL STABILIZER, 9157LANDING FLAP, 161NUTS, 16OUTER WING BOLTING ANGLE, 156RUDDER, 161VERTICAL STABILIZER, 91WING JOINT COVERS,WING STRUCTURE, 91157PARTSHEAT-TREATED, 267TAIL WHEEL COMPONENT, 207ORDERING SPARE (INTRODUCTION)PATCHINGALCLAD SKIN, 137-141FUEL AND OIL TANKS, 195FABRIC, MOVABLE SURFACES,GENERAL, 175INNER TUBES, 209PLASTIC PANELS, 245PLYWOOD SKIN, 84PLASTICSFINISH, 263FUSELAGE FAIRINGS, 70TRANSPARENT, 244PLYWOOD SKINFORMING, 85PATCHING, 84SECTION REPLACEMENT, FUSE. ,85PLYWOOD WEBS, FUSE. BULKHEAD, 86PNEUMATICCHERRY RIVETRIVET GUN, 8GUN, 235RIVET SQUEEZERS, 9PRESS, ARBOR, 25PRESSURE FOR GLUINGNAILING STRIPS, 76WOOD CLAMPS, 76PRESTONE LINES, COLOR CODE, 260PRIMER, ZINC CHROMATE, 22PRIMING, ALUMINUM SHEET, 22PROPELLERCLEARANCE, 2DIAMETER, 2FINISH, 257INSPECTION, 250MINOR REPAIR, 250TYPE, 1PROTECTION, CORROSION, 22PURGING LINES, COLOR CODING, 260QUENCHING, ALCUD SHEET, 19REAMER, SPIRAL, 26REAR FUSELAGEALUM. FORMERS, 61ALUM.ALUM.62ALUM.ALUM.ALUM.ALUM.PART NUMBERS, 60STRINGER TYPE C107LT-40,STRINGER TYPE C180T, 63STRINGER TYPE C234LT, 63STRINGER TYPE C364LT, 63STRINGER TYPE C366T, 64ALUM. STRUCTURE CONSTRUCTION, 60BAGGAGE COMPARTMENT ALUMINUMINTERCOSTAL ANGLE, 64BAGGAGE COMPARTMENT WOODENINTERCOSTAL, 79LOWER ALUMINUM LONGERONS, 66PLYWOOD SKIN REPLACEMENT, 85SUBSTITUTES FOR REPAIR WOOD, 74UPPER ALUMINUM LONGERONS, 64WOODEN CONSTRUCTION, 71WOOD FINISH REQUIREMENTS, 88WOODEN FORMERS, 86WOODEN LOWER LONGERON, 82WOODEN SKIN ARRANGEMENT, 81WOODEN STRINGERS, 77WOODEN UPPER LONGERONS, 80REAR SPARALUM. HORIZONTAL STABILIZER,CENTERSECTION, 122135RESTRICTED[vi]

RESTRICTEDINDEXVERTICAL STARILIZER,I3O-I32REBUSHING SPECIFICATIONS, 254RE-COVERING, PARTIAL FABRIC, 178REPAIR MATERIALFIXED SURF. ALCLAD SHEET, 158FIXED SURFACES EXTRUDED ALUMI-NUM, 158FUSELAGE ALCLAD SHEET, 70FUSELAGE EXTRUDED ALUMINUM, 68STEEL FUSELAGE STRUCTURE, 54WOODEN REAR FUSELAGE, 88RIBSAILERON TRAILING EDGE, 172AILERON NOSE, 173AILERON PART NUMBERS, 164BEADS, 152C'SECTION INTERMEDIATE, 150C'SECTION LEADING EDGE, 151CUTOUTS, 152ELEVATOR AND RUDDER, 169ELEVATOR PART NUMBERS, 163FIXED SURF. PART NUMBERS, 91FLANGES, 152LANDING FLAP PART NUMBERS, 161RUDDER PART NUMBERS, 161WEBS, 153WING TRAILING EDGE, 151RIVETINGCORRECT METHOD, 9EQUIPMENT, 1INCORRECT METHOD, 9IRONS, DU PONT, 239RIVETSAN426-AD, 1AN442-AD, 1B1219-AD, 1B1227-AD, 1CHERRY BLIND, 233DIMENSIONS, 1DRIV ING PRACTICES, 9DU PONT EXPLOSIVE, 238EDGE DISTANCE, 9HOLES, 5IDENTIFICATION, 1MATERIAL, 1REMOVAL OF, 8SELF-PLUGGING CHERRY, 234SETS, 7SQUEEZER, 8TYPES, 1TYPES OF CHERRY, 234RIVNUTSGRIP LENGTH OF, 239HEAD STYLES OF, 239IDENTIFICATION OF, 239INSTALLATION OF, 241PREPARATION FOR INSTALLATIONOF, 241STRENGTH OF, 240TYPE NUMBERS OF, 240TYPES OF, 239USES FOR, 239ROCKWELL HARDNESS, 267RODS, CONTROL, 251ROLLER, HAND, 23RUDDERAREA, 2, 161CABLES, CRANKS, AND HORN MARK-INGS, 261CONSTRUCTION, 161COVERING MATERIALS, 191FABRIC COVERING REQ. , 180FABRIC FINISHES, GENERAL, 264FINISH, 257GENERAL REPAIR, 169LEADING EDGE SKIN, 170MOVEMENT, 2, 161ORIGINAL FABRIC COVERING, 164RIB REMOVAL, 169RIB REPLACEMENT, 170TAB AREA, 2, 161TRAILING EDGE SKIN, 171TRAVEL, 2SALT BATH FURNACE, 19SCLEROSCOPE, SHORE HARDNESS, 267SCRATCHESCOCKPIT PLASTIC PANELS, 245SKIN SURFACE, 22SCREWB1248 SPECIFICATIONS, 14B1251 SPECIFICATIONS, 15FITTING OF, 11STANDARD HOLES FOR, 12TYPES, 12SEALANTFUEL TANK SMALL CRACK, 197SYNTHETIC RUBBER, 197SEAMS, REPAIR OF SPLIT TANK, 198SEATSFINISH, 257WOODEN COCKPIT, 250SETS, RIVET, 7SETTING, VERTICAL STABILIZER, 2SHEET, ALUMINUM, 17SHORE HARDNESS, 267SHRINKER, METAL, 23SIDE PANELSCONSTRUCTION OF FRONT FUS., 57FRONT FUSELAGE FORWARD FORMERREPAIR, 57FRONT FUSELAGE REAR FOUR FOR-MERS REPAIR, 60PART NUMBERS, FRONT FUS., 57SKIN ARRANGEMENTFUSELAGE, 69STABILIZERS, 138WING, 137WOODEN REAR FUSELAGE, 81SKI NALUMINUM FUSELAGE, 68CRACKS, 138CLOSING STRIP, OUTER WING, 147ELEVATOR AND RUDDER LEADINGEDGE, 170LARGE HOLE REPAIR, 138-139LEADING WING EDGE, 143PANEL REPLACEMENT, 148PANEL SPLICING, 141PLYWOOD PATCHING, 84REMOVING DENTS IN HORIZONTALSTABILIZER TIP, 149SKIN REPAIRAILERON LEADING EDGE, 175FIXED SURFACES, GENERAL, I36LANDING FLAP, 166SMALL HOLE, 137WOODEN REAR FUSELAGE, 85SMOKE SCREEN EQUIPMENT LINES,COLOR CODING, 260SPANELEVATOR, 2HORJZONTAL STABILIZER, 2WING, 2SPARALUMINUM HORIZONTAL STABILIZER,FRONT, 132ALUMINUM HORIZONTAL STABILIZER,REAR, 135CENTERSECTION FLAP, 119C'SECTION FRONT LOWER CAP, 126C'SECTION FRONT UPPER CAP, 124C'SECTION FRONT WEB, 128-130CENTERSECTION REAR, 122LANDING FLAP CHANNEL, 167OUTER WING AILERON, 119OUTER W ING FLAP, 119OUTER WING MAIN, 114-116VERTICAL STABILIZER FRONT, I3OVERT. STAB. REAR, I3O-I32SPAR PART NUMBERSAILERONS, 164FIXED SURFACES, 91LANDING FLAP, 161SPAR REPAIRS, GENERAL, FIXEDSURFACES, 113SPAR SPLICE, AILERON, 171SPECIFICATIONSAIRPLANE, 2ALUMINUM ALLOY, 32AN4 BOLT, 13B1248 SCREW, 14B1251 SCREW, 15CABLE MATERIAL, 216COPPER ALLOY, 33EMPENNAGE, 2FINISH MATERIAL, 257LANDING GEAR, 2REBUSHING, 254STEEL ALLOY, 32WING, 2WIRE AND CABLE, 33SPLICINGCONTROL CABLE, 220CONTROL ROD, 251SKIN PANEL, 141WING LEADING EDGE, 143[vii]BESmiCTU)

NDEXRESTRICTEDSPLICING STEEL TUBESGENERAL, 4UINNER SLEEVE METHOD, 46LARGER DIAMETER REPLACEMENTTUBE, 47OUTER SLEEVE METHOD, 47SQUEEZER, PNEUMATIC RIVET, 8STABILIZER

RESTRICTEDINDEXTIRESBALANCING, 210CLEAN ING INTERIOR OF, 210INSPECTION OF, 207LANDING GEAR, 207MOUNTING OF, 211REPAIR OF, 208REPAIR MATERIALS, 215TALC FOR, 212TOOLS FOR MOUNTING, 215TREADING, 209TOOLSARBOR PRESS, 25BLIND RIV. HOLE LOCATING, 140"C" SHEARS, 26CABLE SPLICING, 222ELECTRICAL REPAIR, 227FORMING BLOCKS, 25GENERAL, 22HAND BRAKE, 23HAND ROLLER, 23LANDING GEAR, 215METAL SHRINKER, 23METAL STRETCHER, 23OIL COOLER REPAIR, 205RIVNUT, 241SPIRAL REAMER, 26STANDARD, 23STANDARD METAL WORKING, 30STANDARD POWER AND HAND, 26STEEL FUS. STRUCTURE REPAIR, 55TANK REPAIR, 199TIRE, 215TRIMMING MACHINE, 26TUBE BEADING, 230TUBE FLARING, 229TUBING REPAIR, 232USE OF, 26TORCH TIPS, OXYACET. WELDING, 37TORQUE, MEASUREMENT OF BOLT, 11TORQUE TUBE PART NUMBERS, ELE-VATOR, AND RUDDER, 162-165TRAILING EDGEAILERON PART NUMBER, 164AILERON RIB REPLACEMENT, 172FLAP, 168FLAP PART NUMBERS, 162RUDDER, ELEVATOR, AND AILERONSTRIP REPLACEMENT, 171WING, 113WING PART NUMBERS, 93WING RIBS, 151TRAVELAILERON, 2BOOSTER TAB, 2ELEVATOR, 2RUDDER. 2TRIM TAB, 2TREAD, LANDING GEAR, 2TRIM TABSAREA, 2CONSTRUCTION, 166ELEVATOR AND RUDDER PART NUM-BERS, 161-163TRAVEL, 2TUBE JOINTSREPAIRING SHARP DENTS, 43REPAIRING SMALL CRACKS, 43TUBE REPLACEMENT. 49TUBE SIZESENGINE MOUNT, 34FRONT FUSELAGE, 34TUBESALIGNMENT, 51ALUMINUM, 17BALANCING OF INNER, 210CHECKING ALIGNMENT, 41CORRECTION OF BOWING, 41CRACKS IN LENGTH, 44EXTENT OF DAMAGE, 40GENERAL SPLICING, 44INSPECTION OF INNER, 207LANDING GEAR, 207MAXIMUM ALLOWABLE BOW, 41MOUNTING OF, 211NEGLIGIBLE DAMAGE, 40REMOVING DENTS, 40REMOVING OIL COOLER, 203REMOV ING OVAL SHAPE, 41REPAIR OF INNER, 209REPLACING 01 L COOLER, 203SPLICE, INNER SLEEVE METHOD, 46SPLICE, LARGER DIAMETER REPLACE-MENT TUBE, 47SPLICE, OUTER SLEEVE METH0D,47SPLICE RESTRICTIONS, 46TYPES OF JOINTS, 229TUBINGIRVOLITE INSULATION, 227MATERIAL USED FOR, 229REPAIR OF HIGH PRESSURE, 230REPAIR OF LOW PRESSURE, 23IREPAIR MATERIALS, 233REPAIR TOOLS, 232SIRCO INSULATION, 227TWIST, INCH-POUNDS, 11TYPESBOLTS, 12NUTS, 12SCREWS, 12UNBALANCECONTROL SURF. CORRECTION, 18(CONTROL SURFACE, 187DETERMINATION OF CONTROLSURFACE STATIC, 187UPPER CAP, CENTERSECTION FRONTSPAR, 124UPPER LONGERONSREAR ALUMINUM FUSELAGE, 64WOODEN REAR FUSELAGE, 80VACUUM LINES, COLOR CODING, 260uVAPOR LEAKSFUEL TANK, 19501 L TANK, 195REPAI R OF, 197SEALANT FOR, 197VENT LINES, COLOR CODING, 260VERTICAL STABILIZERFIN ISH, 257FRONT SPAR, I3OGENERAL SKIN REPAIR, I36LARGE SKIN HOLES, I38-I39PART NUMBERS, 91REAR SPAR, 130-132SETTING, 2SKIN ARRANGEMENT, I38SKIN CRACKS, 138SKIN PANEL REPLACEMENT, 148SMALL SKIN HOLE REPAIR, 137SPLIC ING SKIN PANELS, 141wWALKWAYS, FINISH, 257. 260WASHOUT, WING, 2WEBSALUMINUM BULKHEAD, 68C'SECTION FRONT SPAR, 128-130FIXED SURFACE RIB, 153WOODEN REAR FUSE. BULKHEAD, 86-WELDINGCONDITION OF COMPLETED, 38DEPTH, 37ELECTRIC ARC, 36HYDROGEN, 197OVER WELD BEAD FAILURE, 36OXYACETYLENE, 37PREPARATION, 36PROCEDURE, 36TANK INSERT PATCHES, 199TANKS, 198TEMPERATURE, 36WIRE, 37WHEELSFINISH, 257REMOVAL OF LANDING GEAR, 207REPAIR OF, 208TAfL, 207TIRE BALANCE, 210WINGAIRFOAREA,BOLTlCHORDCLOSICONSTIL, 22NG ANGLE PART NOS., 156NG STRIRUCTIONDIHED RAL, 2FINIS H, 257INCID ENCE, 2INSIG N lA FIN ISH, 257JOINTNG ANGLES, 155, BOLTIJOINT COVERS , 157JOINT COVER PART NUMBERS, 157LARGE SKIN H OLES, 138-139LEADI NG EDGE ACCESS, 147LEADI NG EDGE CONTOURS, 146LEADI NG EDGE SKIN, 143OUTER AI LERO N SPAR, 119OUTER FLAP S PAR, 119OUTER MAIN S PAR, 114-116PART NUMBERSSKIN, 91ARRANGE MENT, 137SKIN CRACKS, 138SKIN PANEL R EPLACEMENT, 148SKIN REPAIR, GENERAL, I36SMALLSPAN,SKIN2H OLE REPAIR, 137SPECI FICATIO NS, 2P, OUTER, 147. 91[Ix]RESTRICTED

INDEXRESTRICTEDSPLICING SKIN PANELS, 141STATION, 91SWEEPBACK, 2TRAILING EDGE, 113TRAILING EDGE RIBS, 151WASHOUT, 2WING STRINGERTYPE C123LT, 100TYPE C148T, 101TYPE C180T, 104TYPE C204T, 104TYPE C250T, 105TYPE C265T, 108TYPE C266T,. 109TYPE C274T, 109TYPE C366T, 110TYPE COMBIN., C250T - C366T, 107TYPE DOUBLED C148T, 102TYPE DOUBLED C204T» 105TYPE DOUBLED C250T, 106TYPE DOUBLED C366T, 111TYPE DOUBLED K77A, 112TYPE K77A, 112WIREINSULATION, ELECTRICAL, 224LACING, 223NUMBERING, 223REPLACEMENT, 223ROUTING OF, 223SPECIFICATIONS, 33SPOT TYING, 223TERMINALS, 223WELDING, 2S0, 198WELDING, Ml LD STEEL. 37WIRINGELECTRICAL, 223TYPICAL SWITCH BOX, 226WOODCOCKPIT SEATS, 250FIN ISH MATERIALS, 88FLOOR BOARDS, 250FORMING BLOCKS, EXTRUSIONS,FORMING BLOCKS, RIB, 15625GENERAL FIN ISHES, 265GLUE JOINTS, GENERAL, 75GLUING PRESSURE, CLAMPS, 76GLUING BY, NAILI NG STRI PS, 76MINOR ACCESSORIES, 251MOISTURE, REAR FUS. , 73REPAIR MATERIALS, FUS,, 88WOODEN HORIZONTAL STABILIZERCONSTRUCTION, 91REPAIR, 132WOODEN REAR FUSELAGEBAG. COMPART. INTERCOSTAL, 79BULKHEAD PLYWOOD WEBS, 86FINISH REQUIREMENTS, 88FORMERS, 86LOWER LONGERONS, 82PLYWOOD SKIN PATCHING, 84PLYWOOD SKIN SECTION REPLACE-MENT, 85REPAIR SUBSTITUTES, 74SKIN ARRANGEMENT, 81STRINGERS, 77UPPER LONGERONS, 80ZINC CHROMATE PRIMERDISSIMI LAR METALS, 260SPECIFICATIONS, 258USE OF, 22, 51, 258ZINC PLATING, 257iRESTRICTED[x]

NOTES

NOTES