Wind-tunnel interference effects on a 70° delta wing - CFD4Aircraft

More documents

Recommendations

Info

NUMBER THE AERONAUTICAL JOURNAL OCTOBER 2004model, a change in dynamic pressure about the model due to solidand wake blockage, and in the buoyancy effect due to the axial pressuregradient along the <strong>tunnel</strong> test section. The magnitude of these<strong>effects</strong> increases with model size (increasing solid blockage).Weinberg (4) conducted an experimental investigation into wall<strong>effects</strong>. He tested two sets of three wings (one set with 60° sweep, andone set with 70° sweep), each wing with a different span size. Theexperiment was performed in a square water <strong>tunnel</strong> at a constant flowvelocity of 11ms –1 . The <strong>tunnel</strong> size was 45cm × 45cm. He found thatfor the three wings with 70° sweep, as the wing size was increased(kept at a constant angle-of-attack), vortex breakdown moved downstream.For the three wings with 60° sweep, he found that as the wingspan-to-<strong>tunnel</strong> width ratio increased from 0⋅175 to 0⋅35, the wall<strong>effects</strong> followed the computed trends (i.e. vortex breakdown wasshifted downstream with increasing wing size). However, when thewing span-to-<strong>tunnel</strong> width ratio was increased from 0⋅35 to 0⋅7, nosignificant change was observed. This suggested that effective camberwas not the only influence. For both the 60° and the 70° wings, thedifference in breakdown location observed from the smallest model tothe largest model, was of the order 25% c r .Thompson and Nelson (5) investigated experimentally the influenceof <strong>tunnel</strong> walls on a 70° delta wing by testing full, two thirds, andhalf scale models in a square <strong>tunnel</strong> (the largest model gave theratios S/H = S/W = 0⋅364). Due to a steady hysteresis effect the wingwas tested for a quasi-steady upward and downward stroke. It wasfound that for the smallest model tested (S/H = S/W = 0⋅124) thebreakdown location shifted downstream by as much as 15%c r onboth the quasi-steady upstroke and downstroke. For the half scalemodel and the full scale model, there appeared to be little differencein the breakdown locations. As stated by Thompson and Nelson, thisshift downstream as model size is decreased is in contrast to theresults of Weinberg (4) . It was noted that Weinberg used a Reynoldsnumber an order of magnitude lower, and a constant velocity, asopposed to keeping the Reynolds number constant (as in the experimentsof Thompson and Nelson). It was observed that the vortexsuction on the model surface increased with model size.More recently Pelletier and Nelson (6) studied the effect of <strong>tunnel</strong><strong>interference</strong> on 70° delta wings. Experiments were conducted in awater <strong>tunnel</strong> with three different sized wings. These low Reynoldsnumber tests agreed with the previous findings of Thompson andNelson (5) who tested at higher Reynolds number, in that breakdownmoved towards the apex with increasing wing size. Pelletier andNelson used the method of images to explain this effect, concludingthat the <strong>tunnel</strong> walls increased the mean incidence of the wing, thuspromoting breakdown.Verhaagen et al (7) performed Euler calculations of the flow over a76° delta wing inside wind-<strong>tunnel</strong>s of increasing size. The wingspan-to-<strong>tunnel</strong> width ratios considered were 0⋅292, 0⋅389, and 0⋅584and the test section was octagonal. To model the effect of asecondary separation, a small ‘fence’ was placed where secondaryseparation would occur. It was found that decreasing the <strong>tunnel</strong> size(increasing the wing span-to-<strong>tunnel</strong> width ratio) increased thesuction in the vortices and increased the velocities in the vortex core,due to an increase in circulation with decreasing <strong>tunnel</strong> size.Allan et al (8) performed Euler simulations of <strong>tunnel</strong> <strong>interference</strong><strong>effects</strong> on a 65° delta wing in various <strong>tunnel</strong>s for static and pitchingcases. It was observed that <strong>tunnel</strong> side walls were the most influentialfactor on breakdown location, with roof and floor having littleeffect. This is due to the fact that the roof and floor of the <strong>tunnel</strong> areunlikely to have any influence on the effective angle-of-attack of thewing, or on the development of the vortices. It was also noted that inpitching simulations, the <strong>tunnel</strong> <strong>interference</strong> <strong>effects</strong> were strongeston the downstroke, during the vortex reformation.When considering support <strong>interference</strong> <strong>effects</strong> Hummel (9)observed that the presence of an obstacle one chord length downstreamof the trailing edge of a delta wing caused vortex breakdownto shift towards the apex by up to 40%c r . The obstacle consideredwas unrealistically large (wider than the span of the delta wing),however, this early observation highlighted the possible <strong>effects</strong> thatsupport structures can have. Recently Taylor et al (10) varied the positionof various obstacles behind the wing’s trailing edge to find theeffect support proximity had on vortex breakdown. It was concludedthat when an obstacle was placed in the vortex core, breakdown waspromoted due to the adverse pressure gradient, which formed infront of the obstacle. As the obstacle was moved away from thevortex core, breakdown was observed to move downstream. It canbe expected that the degree of support <strong>interference</strong> will be dependenton support size, vortex strength and trajectory. Although possiblynot considered as support <strong>interference</strong>, fuselages on delta wingmodels can effect the breakdown location. Some sting mounted<strong>tunnel</strong> models have upper surface fuselages to connect the model tothe sting. The effect of such fuselages has been considered in anumber of investigations, each of which have shown significant variationsin breakdown location for a given sweep angle and incidence.See for example Refs (11-13). A possible reason for the variation inthe breakdown locations was given by Ericsson (12) in which theeffect of the fuselage was described as an induced camber effect.The induced camber effect alters the location of vortex breakdown,either delaying or promoting breakdown depending on the fuselage.It is the aim of the current work to evaluate the influence of <strong>tunnel</strong>walls and downstream structures on delta wing aerodynamics. Inorder to include the influence of the secondary separation a 3DRANS model of the flow is assumed. A particular advantage of CFDfor such an investigation is that the various influences can be examinedseparately, thus the qualitative effect of each component of testfacility <strong>interference</strong> can be evaluated and conclusions drawn.2.0 COMPUTATIONAL METHODAll simulations described in this paper were performed using theUniversity of Glasgow PMB3D (Parallel Multi-Block 3D) RANS(Reynolds Averaged Navier-Stokes) solver. A full discussion of thecode and turbulence models implemented is given in Ref. 14.PMB3D uses a cell centred finite volume technique to solve theEuler and RANS equations. The diffusive terms are discretised usinga central differencing scheme and the convective terms use Roe’sscheme with MUSCL interpolation offering third order accuracy.Steady flow calculations proceed in two parts, initially running anexplicit scheme to smooth the flow solution, followed by switchingto an implicit scheme to obtain faster convergence. The pre-conditioningis based on block incomplete lower-upper (BILU) factorisationand is also decoupled between blocks to help reduce the computationaltime. The linear system arising at each implicit step is solvedusing a generalised conjugate gradient (GCG) method. For timeaccuratesimulations, Jameson’s pseudo-time (dual-time stepping)formulation is applied, with the steady state solver used to calculatethe flow steady states on each physical time step (discussed fully inRef. 14).Since the RANS equations are solved the two equation k-ω turbulencemodel is used for closure. It is well known that most linear twoequation turbulence models over-predict the eddy viscosity withinvortex cores, thus causing too much diffusion of vorticity (15) . Thisweakens the strength of the vortices and can eliminate secondaryvortices, especially at low angles of attack where the vortices arealready relatively weak. The following modification suggested byBrandsma et al (16) was therefore applied to the standard k-ω model ofWilcox (17) to reduce the eddy-viscosity in vortex cores.P k = min{P ku,(2⋅0 + 2⋅0min{0, r – 1})ρβ * kω}. . . (1)Here P ku is the unlimited production of k and r is the ratio of themagnitude of the rate-of-strain and vorticity tensors. When k is overpredicted in the vortex core, it will be limited to a value relative tothe dissipation in that region. This modification was found improvepredictions compared with the standard k-ω turbulence model (18) .
ALLAN ET AL WIND-TUNNEL INTERFERENCE EFFECTS ON A 70° DELTA WING NUMBERTable 1ONERA 70° wing test cases — fully turbulent flowTunnel S/W S/H M ∞ ReFarfield – – 0⋅2 1⋅56 × 10 6ONERA F2 0⋅49 0⋅38 0⋅2 1⋅56 × 10 6S/W = 0⋅63 0⋅63 0⋅38 0⋅2 1⋅56 × 10 63.0 TEST CASESThe subsonic case of a 70° delta wing is considered at 27° angle-ofattack.At this incidence vortex breakdown occurs over the wing.The wing has a root chord length of 950mm, a trailing edge length of691⋅5mm, flat upper and lower surfaces, and a leading edge bevel of15°. The wind-<strong>tunnel</strong> model had a blunt trailing edge 20mm thickwhich has been bevelled (15° bevel) in the computational model tosimplify the grid generation. Experimental data for this case wasobtained by Mitchell (19) . The wing was tested in the ONERA F2<strong>tunnel</strong> (situated at Le Fauga-Mauzac Centre, near Toulouse, France)which has a working section of height 1⋅8m, width 1⋅4m, and length5m. It is a subsonic, continuous, closed-return <strong>tunnel</strong>. The wing wasalso tested in the ONERA S2Ch <strong>tunnel</strong> situated at the Chalais-Meudon Centre (ONERA), which is of quasi-circular cross-sectionwith a diameter of 3m and a test section length of 4⋅93m. Themodel/<strong>tunnel</strong> ratios are S/W = 0⋅49 and S/H = 0⋅38 for the ONERAF2 <strong>tunnel</strong>, and S/W = 0⋅23 and S/H = 0⋅55 (taking into account theoff <strong>tunnel</strong> centreline location of wing) for the ONERA S2Ch <strong>tunnel</strong>.Despite the flow conditions being similar in both <strong>tunnel</strong>s, the meanbreakdown locations were closer to the apex in the ONERA F2<strong>tunnel</strong> (the smaller of the two) when compared with those measuredin the ONERA S2Ch <strong>tunnel</strong>. This suggests some test facility <strong>interference</strong>which could be due to either the support structure or <strong>tunnel</strong>wall <strong>interference</strong>.The flow conditions for which extensive experimental data fromthe ONERA F2 <strong>tunnel</strong> are available, are an incidence of 27°,Reynolds number of 1⋅56×10 6 , and a freestream velocity of 24ms –1(M ∞ = 0⋅069). For these flow conditions, flow visualisation ofvelocity components, vorticity components, turbulent kinetic energycomponents and local static pressure are available in various planes.Vortex breakdown locations were also obtained based on where theaxial component of velocity becomes negative. In the current worksteady state simulations are performed in all cases, with theReynolds number being matched, however the Mach number is 0⋅2,compared with the lower Mach number of 0⋅069 in the experiment.Since the current work is based on a high speed flow solver, afreestream Mach number of 0⋅2 was used to avoid any possibleconvergence issues. Transition was observed to occur at around40%c r in experiment, however the CFD simulations assume a fullyturbulent flow. As will be described the fully turbulent flow assumptionwill influence the validation of the results. However, the <strong>tunnel</strong><strong>interference</strong> will mainly be dependent on vortex strength, thus thefully turbulent assumption is unlikely to influence the trendsobserved (the leading edge separation location is fixed due to thesharp leading edge and breakdown is known to be independent ofReynolds number (20) ).Three boundaries have been chosen. The first has the wing in freeair (farfield conditions), the second represents the ONERA F2<strong>tunnel</strong>, and the third the ONERA F2 <strong>tunnel</strong> side walls brought closerto the wing (increasing the S/W ratio). Full details of these configurationsas well as flow parameters can be found in Table 1.To assess the <strong>effects</strong> of downstream structures in the <strong>tunnel</strong> avertical (generic) structure was placed in the centre of the ONERAF2 <strong>tunnel</strong>, downstream of the wing. Since supports in dynamictesting tend to be fairly large (driving mechanisms must also behoused) a thick cross-sectional area for the support was chosen. Thegeneric structure considered is a cylinder with a straight taper in theFigure 1. ONERA F2 <strong>tunnel</strong> setup with downstream structure.Figure 2. Geometry of downstream support structure (plan view).downstream direction. Figure 1 shows the support structure and itsplacement within the ONERA F2 <strong>tunnel</strong>. To avoid grid generationcomplications the mounting attaching the wing to the vertical structureis omitted (it may be expected that there is further <strong>interference</strong>from this mounting). The downstream structure (shown in Fig. 2with dimensions) gives a frontal area blockage of around 12%.The support structure was placed at two downstream locations inthe ONERA F2 <strong>tunnel</strong>, at 0⋅5c r and 1c r from the trailing edge of thewing. The mesh over the wing is identical to that for the ONERA F2<strong>tunnel</strong> without downstream structures. Therefore there is no changein grid resolution in the vortical region above the wing. The ONERAF2 <strong>tunnel</strong> mesh was altered downstream of the wing to allow thepresence of the structure.In all cases considered the <strong>tunnel</strong> walls and downstream supportstructures have been modelled with inviscid wall boundary conditions.Eliminating the <strong>tunnel</strong> wall boundary layers reduces grid sizes,however, a consequence of this is that the favourable pressuregradient within the test section (due to <strong>tunnel</strong> wall boundary-layer
Page 1: THE AERONAUTICAL JOURNAL OCTOBER 20
Page 5 and 6: ALLAN ET AL WIND-TUNNEL INTERFERENC
Page 7 and 8: ALLAN ET AL WIND-TUNNEL INTERFERENC
Page 9: ALLAN ET AL WIND-TUNNEL INTERFERENC

Wind-tunnel interference effects on a 70° delta wing - CFD4Aircraft

Create successful ePaper yourself

Delete template?

Save as template?