NIAC DARE Phase 2 Final Report 510-02421-011 - NASA's Institute ...

Sailing the Planets: Science from DirectedAerial Robot Explorers (DARE)Phase II Final ReportGlobal Aerospace Corporation14 April 2006Universities Space Research Association Research Grant No.: NAS5-03110GAC Report 510-02421-011

Phase II Study ContributorsGlobal Aerospace CorporationDr. Kim M. AaronNathan C. BarnesNam NguyenKerry T. NockDr. Alexey A. Pankine, Principal Investigator and NIAC FellowR. Stephen Schlaiferii

AbstractGlobal Aerospace Corporation is developing a revolutionary architecture that opens new andexciting pathways for planetary exploration. At the core of the architecture are the DirectedAerial Robot Explorers (DAREs), which are autonomous balloons with path guidancecapabilities that can deploy swarms of miniature probes over multiple target areas. The DAREplatforms will explore the planets in concert with orbiters and surface platforms (landers, rovers,dropsondes). Flight path guidance will offer unprecedented opportunities in high-resolutiontargeted observations of both atmospheric and surface phenomena on global scales.Multifunctional microprobes will be deployed from the balloons over the targeted sites, andperform a multitude of functions, such as, atmospheric profiling or surface exploration, relayingdata back to the balloons or an orbiter. This architecture will enable low-energy, long-termglobal exploration of planetary atmospheres and surfaces. The proposed effort addressesobjectives of several NASA Enterprises and of the new vision for space exploration recentlyannounced by the President. These objectives are related to the understanding of solar systemformation and evolution, the search for life and its beginnings on other planets, the investigationof the composition, evolution, and resources of Mars, collaborative robotic missions to enablehuman exploration, and others. The technologies conceptualized in this study - such as verylong-duration-flight balloons, flight path guidance, microprobes, and planetary platformnavigation and communications - have direct relevance to future in situ exploration missions toMars, Venus and Titan. Mission cost estimates will assist in the assessment of the performanceand benefits of the concept and fitting it within future exploration mission cost categories. Theresults of this study can be applied directly to the development of future New Frontier,Discovery, or Scout-type missions to Mars, Venus and Titan. The key elements of the overallDARE architecture are listed below:• Long-duration lightweight planetary balloon platform(s)• Balloon flight path guidance• Autonomous navigation and control• Lightweight and efficient power generation and energy storage• Lightweight balloon platform science suite• Lightweight deployable science packages• Communication relay orbiter• Synergy between surface, airborne and orbital elements of the architecture.iii

8.1 APPENDIX 1. “SAILING THE PLANETS: PLANETARY SCIENCE FROM GUIDED BALLOONS”, PAPERPRESENTED AT THE 17 TH ESA SYMPOSIUM ON EUROPEAN ROCKET AND BALLOON PROGRAMMES AND RELATEDRESEARCH, MAY 30-JUNE 2, 2005, SANDEFJORD, NORWAY.................................................................................1918.2 APPENDIX 2. “SAILING THE PLANETS: EXPLORING MARS FROM GUIDED BALLOONS”, PAPER PRESENTEDAT THE AIAA 5TH AVIATION, TECHNOLOGY, INTEGRATION, AND OPERATIONS CONFERENCE (ATIO), 26-28SEPTEMBER 2005, ARLINGTON, VA......................................................................................................................1928.3 APPENDIX 3. “SAILING THE PLANETS: PLANETARY EXPLORATION FROM GUIDED BALLOONS”,PRESENTATION AT THE 7 TH ANNUAL NIAC MEETING, DENVER, CO, OCTOBER 10-11, 2005................................193vi

FIGURE 4-11. DEPENDENCE OF THE VENUS BGS V VELOCITY ON THE TETHER LENGTH AND BALLOON ALTITUDE......92FIGURE 4-12. TETHER CONFIGURATION FOR RUN 21 IN TABLE 4-7. .............................................................................93FIGURE 4-13. HISTOGRAM OF THE WIND DIRECTION DIFFERENCES BETWEEN 5 AND 10 KM ALTITUDES IN THE MARTIANATMOSPHERE ......................................................................................................................................................97FIGURE 4-14. HISTOGRAM OF WIND SPEED DIFFERENCES BETWEEN 10 AND 4 KM FOR LS=150°..................................98FIGURE 4-15. HISTOGRAM OF THE WING REYNOLDS NUMBERS....................................................................................99FIGURE 4-16. 3D VIEWS OF A PROLATE SPHEROID......................................................................................................100FIGURE 4-17. AERODYNAMIC DRAG FOR ROTATIONALLY-SYMMETRIC BODIES..........................................................101FIGURE 4-18. BGS V WINDS FOR DIFFERENT DARE PLATFORMS..............................................................................103FIGURE 4-19. VERTICAL WIND PROFILE IN THE NORTHERN HEMISPHERE ON 3/22/2014.............................................105FIGURE 4-20. VECTOR DIAGRAM OF THE DAILY RELATIVE WINDS AT THE BGS WING ALTITUDE ...............................105FIGURE 4-21. SIMULATED TRAJECTORY, LS=89º........................................................................................................109FIGURE 4-22. ANOTHER SIMULATED TRAJECTORY FOR LS=89°. ................................................................................109FIGURE 4-23. SIMULATED DARE TRAJECTORY FOR OBSERVATIONS OF CRUSTAL MAGNETIC ANOMALIES, LS=200°.110FIGURE 4-24. SIMULATED TRAJECTORY OF A FREE-FLOATING BALLOON (RED) AND A DARE PLATFORM (BLACK)PLOTTED OVER THE MAP OF CRUSTAL MAGNETIC ANOMALIES. .........................................................................111FIGURE 4-25. ALTITUDE OF THE DARE PLATFORM FOR THE SIMULATION ON FIGURE 4-23.......................................111FIGURE 4-26. SIMULATED TRAJECTORY OF A DARE PLATFORM CROSSING EQUATOR FROM SOUTHERN TO NORTHERNHEMISPHERE IN EARLY NORTHERN SPRING, LS=23º. .........................................................................................112FIGURE 5-1. DARE SYSTEM DEFINITIONS. .................................................................................................................113FIGURE 5-2. NASA TECHNOLOGY READINESS LEVELS (TRL) DEFINITIONS. ............................................................114FIGURE 5-3. FOLD THEN ROLL AERIAL DEPLOYMENT. ................................................................................................120FIGURE 5-4. INFLATING BALLOON AND INNER TETHER/REINFORCED INFLATION TUBE...............................................121FIGURE 5-5. SAMPLE OF THE COMPOSITE BALLOON ENVELOPE MATERIAL 2 ................................................................125FIGURE 5-6. QUANTUM DOT SCHEMATIC DRAWING....................................................................................................128FIGURE 5-7. SINGLE-WING BGS................................................................................................................................130FIGURE 5-8. ILC DOVER INFLATABLE WING EXPERIMENT. ........................................................................................131FIGURE 5-9. ILC DOVER GUN-LAUNCHED CONCEPT ..................................................................................................131FIGURE 5-10. INFLATABLE DARE BGS.....................................................................................................................132FIGURE 5-11. LOWER PORTION OF THE DARE BGS SHOWING THE BOOM WITH THE RUDDER, SOLAR PANELS AND THEGIMBALED PLATFORM WITH THE IMAGING CAMERAS........................................................................................133FIGURE 5-12. DUAL-WING BGS. ...............................................................................................................................134FIGURE 5-13. NORTHERN POLAR REGION MOLA TOPOGRAPHY ................................................................................140FIGURE 5-14. MAXIMUM PAYLOADS AND BALLOON RADIUS FOR DARE PLATFORM DESIGN.....................................141FIGURE 5-15. VARIATIONS OF THE HEIGHT OF A CONSTANT DENSITY LEVEL ACROSS MARS ......................................142FIGURE 5-16. ANNUAL LIMITS OF THE HEIGHT VARIABILITY FOR A CONSTANT DENSITY LEVEL AT MARS .................143FIGURE 5-17. FLOATING ALTITUDES FOR VARIOUS PAYLOADS AS A FUNCTION OF ATMOSPHERIC DUST LOADING......144FIGURE 5-18. CLOUDS OVER NORTH POLE ON MARS. ................................................................................................146FIGURE 5-19. VARIATIONS OF TA/TG WITH SEASON AND LATITUDE. .........................................................................148FIGURE 5-20. VARIATION OF MAXIMUM TA/TG WITH ALTITUDE FOR ALL LATITUDES FOR LS=90°............................149FIGURE 5-21. EXTREMES OF THE DARE BALLOON FILM TEMPERATURE FOR THE 6 DESIGNS IN TABLE 5-5. ..............150FIGURE 5-22. PEAK STRESS AND MAXIMUM PAYLOADS FOR DESIGN #6. ....................................................................151FIGURE 5-23. ALTITUDE CHANGE AND THE MASS OF THE VENTED GAS FOR DESIGN #6. .............................................152FIGURE 5-24. VARIATION IN SYSTEM MINIMUM ALTITUDE WITH PARACHUTE SIZE ....................................................154FIGURE 5-25. VARIATION IN SYSTEM MINIMUM ALTITUDE WITH INFLATION TIME......................................................155FIGURE 5-26. VARIATION IN SYSTEM MINIMUM ALTITUDE WITH INFLATION INITIATION DYNAMIC PRESSURE............156FIGURE 5-27. VARIATION IN SYSTEM MINIMUM ALTITUDE WITH SYSTEM FLOAT ALTITUDE.......................................157FIGURE 5-28. MINIMUM ALTITUDE REACHED DURING EDI FOR ATMOSPHERIC CONDITONS.......................................159FIGURE 5-29. MAXIMUM ALLOWABLE GROUND SPEED AS A FUNCTION OF THE TELESCOPE F-NUMBER AND SPATIALRESOLUTION (KLAASEN, 2000).........................................................................................................................161FIGURE 5-30. STATISTICS OF THE WIND SPEED MAGNITUDES AT ABOUT 12 KM ALTITUDE AT MARS FOR L S =270 ANDLOW DUST CONDITIONS.....................................................................................................................................161FIGURE 5-31. DEPENDENCE OF THE OPTICAL SYSTEM MASS ON THE APERTURE .........................................................163FIGURE 5-32. OPTICAL SYSTEM APERTURE IN METERS AS A FUNCTION OF DESIRED SURFACE RESOLUTION(DIFFRACTION LIMITED)....................................................................................................................................164viii

FIGURE 5-33. OPTICAL SYSTEM MASS IN KG VERSUS DESIRED SURFACE SPATIAL RESOLUTION (DIFFRACTION LIMITED)..........................................................................................................................................................................164FIGURE 5-34. DEGRADATION OF THE SOLAR PANEL POWER OUTPUT DUE TO ACCUMULATION OF DUST. ....................167FIGURE 5-35. TRANSMITTANCE OF THE SOLAR POWER THROUGH MARTIAN DUSTY ATMOSPHERE FOR VARIOUS SOLARZENITH ANGLES.................................................................................................................................................168FIGURE 5-36. ILLUMINATION OF THE GONDOLA SOLAR PANEL AS A FUNCTION OF THE BALLOON-GONDOLA TETHERLENGTH.............................................................................................................................................................169FIGURE 5-37. DARE PLATFORM SOLAR ARRAY POWER CAPABILITY ALONG THE MISSION TRAJECTORY AS A FUNCTIONOF MISSION TIME...............................................................................................................................................171FIGURE 5-38. DARE PLATFORM SOLAR ARRAY POWER CAPABILITY ALONG THE MISSION TRAJECTORY AS A FUNCTIONOF “DARE TIME”..............................................................................................................................................172FIGURE 5-39. DARE DAY DURATION AND THE LATITUDE OF THE PLATFORM. ...........................................................172FIGURE 5-40. ENERGY CONSUMPTION DUE TO THE BGS REELING..............................................................................173FIGURE 5-41. DAILY ENERGY GENERATED BY 0.41 M 2 GONDOLA SOLAR ARRAY........................................................174FIGURE 5-42. DAILY ENERGY GENERATED BY 0.79 M 2 BGS SOLAR ARRAY. ..............................................................174FIGURE 5-43. GONDOLA BATTERY CHARGE (MAXIMUM ENERGY 700 W-HR).............................................................175FIGURE 5-44. BGS BATTERY CHARGE (MAXIMUM ENERGY 200 W-HR) .....................................................................176FIGURE 5-45. ELBBO LEAK RATES SUMMARY AND DP DEPENDENCE........................................................................178FIGURE 6-1. AGU DARE CONCEPT POSTER...............................................................................................................182ix

List of TablesTABLE 3-1. MARS POTENTIAL NEXT DECADE PATHWAYS...........................................................................................28TABLE 3-2. ENTRY VEHICLE MASS BUDGET..................................................................................................................53TABLE 3-3. CONCEPTUAL DARE ESN MISSION MASS BUDGET ...................................................................................56TABLE 3-4. GONDOLA POWER AND ENERGY BUDGET ...................................................................................................57TABLE 3-5. BGS POWER AND ENERGY BUDGET ...........................................................................................................58TABLE 3-6. CONFIGURATION OF THE DARE PLATFORM FOR ESN SCENARIO ..............................................................62TABLE 3-7. BALLOONS PARAMETERS FOR THE TRAJECTORY IN FIGURE 3-32...............................................................68TABLE 4-1. NOTATION FOR FIGURE 4-2. ......................................................................................................................74TABLE 4-2. DYNAMIC SCALING PARAMETERS .............................................................................................................77TABLE 4-3. INPUT PARAMETERS FOR THE BGS MODEL RUN AT LS=6º .........................................................................79TABLE 4-4. OUTPUT PARAMETERS OF THE BGS MODEL FOR LS=6°.............................................................................81TABLE 4-5. INPUT PARAMETERS FOR THE BGS MODEL RUN LS=89º ............................................................................84TABLE 4-6. OUTPUT PARAMETERS OF THE BGS MODEL FOR LS=89°...........................................................................86TABLE 4-7. VENUS BGS MODELING SUMMARY ...........................................................................................................91TABLE 4-8. OUTPUT OF THE BGS MODEL FOR VENUS BGS RUN 21.............................................................................93TABLE 4-9. DARE BGS PARAMETERS FOR THE TRAJECTORY ON FIGURE 4-21..........................................................108TABLE 5-1. DARE MISSIONS REQUIREMENTS ............................................................................................................115TABLE 5-2. DARE BALLOON FLIGHT SYSTEM MASSES FOR THE THREE TECHNOLOGY HORIZONS ..............................116TABLE 5-3. DARE ARCHITECTURE CONCEPTUAL TECHNOLOGY ROADMAP. ..............................................................118TABLE 5-4. INFLATION SYSTEM MASS BUDGET...........................................................................................................123TABLE 5-5. THERMO-OPTICAL PROPERTIES FOR VARIOUS DARE BALLOON DESIGNS USED IN THE ANALYSIS............147TABLE 5-6. VARIOUS ATMOSPHERES CONSIDERED FOR EDI TRADE STUDIES. ATMOSPHERES 1 AND 2 REPRESENTLIMITING CASES. ...............................................................................................................................................158TABLE 5-7. SPACECRAFT, BALLOON AND COMMERCIAL TELESCOPE CHARACTERISTICS. ...........................................162TABLE 5-8. PARAMETERS OF THE DARE OPTICAL IMAGING SYSTEMS.......................................................................165TABLE 5-9. SOLAR ARRAYS AND BATTERIES SIZING RESULTS. ...................................................................................176TABLE 5-10. GAS LOSS ESTIMATES FOR DARE MARS BALLOON FOR DIFFERENT ASSUMPTIONS................................179x

1 IntroductionThis is the final report for Phase II of NIAC Contract Number NAS5-03110 for the developmentof the concept for the new planetary exploration architecture with Directed Aerial RobotExplorers (DARE). This study includes analysis of the science objectives and of the potentialobservations enabled by the architecture, conceptual design of the architecture and developmentof conceptual mission scenarios, identification of the enabling technologies, analysis of thebenefits offered by the architecture development, and of the pathways to the concept realization.In the following sections we summarize the concept, discuss its key elements and objectives, itsrevolutionary nature and its significance to NASA.1.1 Sailing the Planets: Science from Directed Aerial Robot Explorers(DARE)Global Aerospace Corporation is developing a revolutionary architecture that opens new andexciting pathways for planetary exploration. At the core of the architecture are the DirectedAerial Robot Explorers (DAREs), which are autonomous balloons with path guidancecapabilities that can deploy swarms of miniature probes over multiple target areas. The DAREplatforms will explore the planets in concert with orbiter(s) and surface platforms (landers,rovers, subsurface penetrators), and dropsondes. Flight path guidance will offer unprecedentedopportunities in high-resolution targeted observations of both atmospheric and surfacephenomena on global scales.Figure 1-1. Key elements of the DARE architecture.11

Multifunctional microprobes will be deployed from the balloons over the target areas, andperform a multitude of functions, such as atmospheric profiling or surface exploration, relayingdata back to the balloons or an orbiter. This architecture will enable low-energy, long-termglobal exploration of planetary atmospheres and surfaces. The proposed effort addressesobjectives of several NASA Enterprises and of the new vision for space exploration recentlyannounced by the President. These objectives are related to the understanding of solar systemformation and evolution, the search for life and its beginnings on other planets, the investigationof the composition, evolution, and resources of Mars, collaborative robotic missions to enablehuman exploration, and others. The technologies conceptualized in this study - such as verylong-duration-flight balloons, flight path guidance, microprobes, and planetary platformnavigation and communications - have direct relevance to future in situ exploration missions toMars, Venus and Titan. The results of this study can be applied directly to the development offuture New Frontier, Discovery, or Scout-type missions to Mars, Venus and Titan.1.2 Keys to the ConceptThe technical and programmatic keys to this new concept are the development of the newobservational approaches, demonstration of balloon guidance capabilities, significant and costeffective scientific applications, identification of enabling technologies, and the development ofthe pathways to architecture realization. Specifically, the key elements of the overall DAREarchitecture are listed below:• Long-duration lightweight planetary balloon platform(s)• Balloon flight path guidance• Autonomous navigation and control• Lightweight and efficient power generation and energy storage• Lightweight science instruments• Lightweight deployable science packages• Communication relay orbiter• Synergy between surface, airborne and orbital elements of the architecture1.3 Concept ObjectivesThe objectives of this concept are to:1. Enable new planetary exploration architecture for low-cost, high-resolution and targeted insitu and remote sensing of planetary atmospheres and surfaces,2. Identify technologies to enable new planetary exploration applications and measurements;3. Instigate development of new observational techniques by planetary scientists, and4. Facilitate NASA meeting the objectives of the Strategic Plan and of the Vision for theNation’s Space Exploration Program by offering a new observational vantage point from whichto explore Mars and other planets.12

The primary objective of this concept is to enable low-cost, high-performance and targeted insitu and remote sensing of planetary atmospheres and surfaces through the utilization of the newplanetary exploration architecture incorporating balloon platforms with path guidancecapabilities, miniature deployable probes, surface landers/rovers and orbiters. Such architecturewould enable unique measurements and observations not feasible by any other means, wouldgreatly expand our knowledge about the Solar System and provide reconnaissance capabilitiesfor future human planetary exploration.Advances in technology in the near future will dramatically reduce the mass of balloonplatforms, extend balloon-platform life beyond several days to more than a year, provideefficient power generation and energy storage, further miniaturize sensors and electronics. Theseadvances will enable radically new measurements, such as emplacement of global seismological,geophysical, geochemical and meteorological networks on the surfaces of Mars and otherplanets; enable repeated soundings of the atmosphere with small densely packed sondesequipped with multiple sensors; monitor the state of the atmosphere continuously through severalseasons; search for life habitats on the surface of Mars, enable very high-resolution imaging ofthe Martian surface, etc.The proposed DARE architecture offers a new approach to performing in situ and remote sensingobservations in planetary environments. The unique features of the DARE architecture - longmission duration combined with low observing altitudes, targeted observations and possibilitiesfor combined surface/atmosphere measurements - were never exploited together before.Collaborative efforts of the planetary science community are needed to develop newobservational techniques and methodologies that would maximize the science return from theDARE architecture implementation.The development of the DARE architecture concept facilitates NASA meeting the objectives ofthe Strategic Plan and of the Vision for the Nation’s Space Exploration Program by providing anew observational vantage point from which to explore Mars and the other planets and a newreconnaissance robotic platform to assist human exploration of Mars and other planets.1.4 Phase II Technical and Programmatic ObjectivesThe Phase II technical and programmatic objectives are to:• Continue the development of the concept of the DARE architecture in the context of a futureNASA Mars mission and identify enabling technologies.• Develop new trajectory control and observational techniques that would increase the sciencereturn from the DARE planetary exploration architecture.• Demonstrate, by analysis, performance advantages and benefits of the proposed DAREplanetary exploration architecture compared to the traditional approaches. Estimate potentialcosts of DARE planetary exploration architecture and compare with cost of conventionalexploration approaches.13

• NASA embraces the DARE planetary exploration architecture concept and begins to use it asa framework on which to plan future planetary exploration activities.1.5 What Makes This Concept RevolutionaryA combination of several factors makes this concept revolutionary.First, the core of the new architecture is shifted from orbiters, landers or rovers to the airborneplatform. Previously, planetary explorations architectures included orbiters and landers (Vikings)or orbiters and rovers (Mars Global Surveyor (MGS), Mars Odyssey (MO) and Mars ExplorationRovers (MERs)). The proposed new architecture is built around a planetary balloon – a longdurationobservational platform that navigates and explores the planets in tandem with theobservational platforms in orbit around the planet, on its surface or in the depth of itsatmosphere. Planetary balloons offer a new vantage point for planetary exploration on a globalscale and also serve as a platform for targeted delivery of lightweight surface and atmosphericrobotic probes. They combine global coverage enabled by orbiters with the high-resolutionobservations afforded by surface vehicles. Flight path guidance revolutionizes the capabilities ofthe balloon. A balloon platform can now be used as an active observational tool rather than apassive winds tracer. Together, with low power consumption and long flight duration, the pathcontrol capability opens up possibilities for continuous long term monitoring of planetaryenvironments, targeted observations and near global coverage not possible with any otherplatform.Guided balloons can enable precision deployment of small, or micro, probes as opposed to thelow accuracy attainable for small direct-entry vehicles. In addition, the fact that massive entrysystems are unnecessary for each small probe means more can be carried to Mars. The use ofmultiple, deployable low-cost targeted micro in situ packages promises to deliver the level ofsampling of the atmospheres and surfaces that previously could not be achieved. Multiplelocations could be sampled with a single DARE platform providing a comprehensive picture ofthe composition and processes in the atmosphere and on the surface. The proposed architectureof guided DARE platforms, orbiters and surface vehicles creates the most powerful explorationsystem ever. The whole history of how planets came to be can be read.The second factor that makes the concept revolutionary is the synergy between orbiter(s),balloon(s), surface and atmospheric probes and mobile vehicles, and - possibly in the not toodistant future - human explorers. The presence of multiple observing platforms simultaneouslyexploring a planet from different vantage points presents unique opportunities for insight intodifferent temporal and spatial scales of the processes that shaped the planets, individually, andour Solar System, as a whole. The first ever simultaneous observations of the Martianatmospheric temperatures from the surface to the upper reaches of the atmosphere by the MGS,MO and MERs clearly demonstrate the advantages of the synergetic observations. The combinedobservations by the same platforms also clearly demonstrate that the difficulty in interpretingsurface geology and mineralogy from orbit and the limited range of surface vehicles necessitatean airborne platform with a global reach at intermediate altitudes. Such a platform is a planetaryballoon with flight path guidance capabilities.14

These significant differences from traditional approaches promise to dramatically increase thescientific return of planetary exploration, advance our understanding of the Solar System, andprovide robotic support to human planetary exploration.1.6 Significance to NASAThe proposed concept addresses a number of strategic goals and objectives outlined in the NASAStrategic Plan 2003.The concept supports NASA’s goal of exploring the Solar System, understanding the origin andevolution of life, and searching for evidence of life elsewhere by providing a novel architecturefor unique and systematic scientific observations, through various in situ and remote sensingmeans, of the surfaces and in the atmospheres of the planetary bodies in the Solar System. Theproposed architecture will provide the means to determine if life exists or has existed on Marsand the other planets. High-resolution imaging of potential landing sites from DARE platformson Mars will help in identification of potential hazards (slopes, rocks) and resources for humanexploration (shallow underground ice or water).The proposed concept supports NASA’s goal of inspiring and motivating students to pursuecareers in science by envisioning potential future involvement of university students in thedevelopment of the elemental subsystem of the proposed architecture (microprobes).The concept also supports NASA’s goal of enabling revolutionary capabilities through newtechnology by proposing to identify and develop new technologies and capabilities, such as theBalloon Guidance System, lightweight power generation and energy storage systems, newcommunication approaches, and the miniature deployable science packages (microprobes) - thatwill enable a variety of new observational approaches. The proposed autonomous DAREplatforms implementing Balloon Guidance can provide the basis for airborne platforms on Mars,Venus, Titan and Jupiter. The same balloon path guidance technology could lead to therealization of NASA’s vision of the high-altitude, long-endurance, unmanned aerial vehicles,providing expanded capability for Earth science applications.The concept is responsive to the new Vision for the Nation’s Space Exploration Program bydeveloping new options for reconnaissance of Martian resources, landing sites and sites that mayhave harbored life in the past, for the monitoring of such environmental hazards as dust andradiation on Mars, for targeted emplacement of navigational beacons for landing crafts, forproviding beyond-line-of-sight communication capabilities.In addition to the benefit of creating new observational opportunities, the proposed architecturecan have cost and performance benefits to NASA.15

2 Concept Development Summary2.1 Summary of Phase II TasksPhase II of the Sailing the Planets: Science from Directed Aerial Robot Explorers (DARE)development includes the following tasks as originally planned and described.2.1.1 Task 1 Conceptual Architecture Design and Technology IdentificationContinue the development of the revolutionary DARE architecture with a focus on Mars.Develop conceptual architecture designs in the context of future NASA Mars exploration.Develop observational techniques created by the DARE architecture that utilize opportunities forconcerted and/or collaborative observations between the satellites, balloon platforms, rovers,landers, subsurface explorers, and microprobes. Identify technologies that are crucial for thedevelopment and successful implementation of the DARE concept. Sketch a preliminaryroadmap of necessary further development and technological milestones of the concept.2.1.2 Task 2 Balloon Guidance ResearchExpand the path guidance analysis started in Phase I to employ more complex atmosphericmodels and environments. Develop the Advanced Balloon Guidance System (BGS) performancemodel and incorporate it into the trajectory simulation code. Continue research on trajectoriesand potential mission scenarios in support of the identified science objectives and measurements.Demonstrate, by realistic simulations, the performance of guided planetary balloons.2.1.3 Task 3 Cost and Performance Benefits AnalysisAnalyze DARE cost and performance benefits by assessing the unique observations that can bemade in the context of the DARE architecture and by comparing these to observations from otherobservational platforms or with other means. Develop preliminary cost estimates for the DAREarchitectures at different planets.2.1.4 Task 4 Identify Pathways to Architecture RealizationBrief planetary scientists on the concept. Seek new science applications for the architecture.Conduct discussions on the DARE concept with NASA mission and program planners from theOffice of Exploration Systems and Office of Space Science. To facilitate discussion with NASAscientists, planners and program managers we will invite key NASA and JPL personnel tobriefings to present the results of our studies and travel to NASA facilities to engage NASApersonnel in discussions on the potential benefit of embracing a DARE transportationarchitecture.16

2.1.5 Task 5 Planning and ReportingA detailed Phase II plan will be developed at project initiation. Bimonthly status reports, a midtermreport and a final report shall be written. We shall participate in and present a status reportat the NIAC Fellows Conference. GAC will support a site visit by the NIAC Director at the GACfacilities in Altadena, CA, if deemed appropriate by the NIAC Director. Copies of all briefings,presentations or professional society technical papers pertaining to the Phase II study will beprovided to NIAC.2.2 Summary of Work AccomplishedThis section provides a concise summary of the work accomplished during the Phase II effort. Amore detailed description of the work follows in later sections. The progress made during thePhase II is illustrated by this summary of accomplishments:August-September 2004• Developed approach to continued development of the DARE architecture (Task 1)• Developed Approach to Balloon Flight Path Guidance research (Task 2)• Developed model for the Advanced BGS on Mars (Task 2)• Worked on defining parameter space for BGS model studies (Task 2)• Developed detailed schedule for the Phase II effort (Task 5)October-November 2004• Researched potential altitudes for balloon flight on Mars (Task 1)• Researched ways to increase performance of the BGS (Task 2)• Priced the development of a computer-animated movie presentation of the DARE concept• (Task 4)• Presented the concept at the NIAC meeting (Task 5)• Submitted our first bi-monthly report (Task 5)December 2004 - January 2005• Started work on detailed definition of observations and applications for DARE platforms onMars (Task 1)• Established the parameter space for the BGS analysis (Task 2)• Started work on development of the Mars atmospheric trajectory simulation software (Task2)• Started analysis of the Martian winds in support of the BGS performance analysis (Task 2)• Presented the concept at the Fall Meeting of the American Geophysical Union (AGU) (Task4)• Submitted our second bi-monthly report (Task 5)February-March 2005• Continued work on detailed definition of observations and applications for DARE platformson Mars (Task 1)• Finished developing the Single-wing Balloon Guidance System model (Task 2)• Continued work on development of the Mars atmospheric trajectory simulation software(Task 2)17

• Continued analysis of the Martian environment in support of the BGS performance analysis(Task 2)• Submitted our third bi-monthly report (Task 5)April-May 2005• Started mission scenario analysis for the surface network emplacement mission (Task 1)• Continued analysis of the Martian environment in support of the DARE conceptual designdevelopment (Task 2)• Continued work on defining enabling technologies (Task 2)• Continued work on development of the Mars atmospheric trajectory simulation software(Task 2)• Continued analysis of the Martian winds in support of the BGS performance analysis (Task2)• Presented the DARE concept to the Mars Advanced Studies Group at JPL (Task 4)• Presented the DARE concept at the 17th ESA Symposium on European Rocket and BalloonProgrammes and Related Research (Task 4)• Submitted our fourth bi-monthly report (Task 5)June - July 2005• Defined new measurements and observations for DARE (Task 1)• Researched potential science instruments (Task 1)• Calculated impact velocities for deployable probes (Task 1)• Outlined 4 mission scenarios for further analysis (Task 1)• Developed example system design for Surface Network Emplacement mission scenario(Task 1)• Researched enabling technologies (Task 1), including Entry, Descent and Inflation (EDI)approaches and Inflation Hardware options• Continued work on development of the Mars atmospheric trajectory simulation software(Task 2)• Continued work on development of the numerical model of the BGS (Task 2)• Analyzed BGS performance in representative environment (Task 2)• Studied BGS performance in low Reynolds number regime (Task 2)• Analyzed dynamic scaling of the Mars BGS (Task 2)• Contacted NASA MEP management (Task 4)• Submitted our fifth bimonthly report (Task 5)• Submitted mid-term report (Task 5)August - September 2005• Started work on detailed system and mission design (Task 1)• Continued development of the numerical model of the BGS (Task 2)• Developed approach to costing and benefits analysis (Task 3)• Issued a press-release on the development of the concept• Presented a paper on the “Sailing the planets” concept at the AIAA 5th Aviation,Technology, Integration, and Operations Conference (ATIO), 16th Lighter-Than-AirSystems Technology Conference and Balloon Systems Conference in Arlington, VA onSeptember 26-28, 2005. (Task 4)• Submitted our sixth bimonthly report (Task 5)18

• Hosted a NIAC site visit on August 8, 2005 (Task 5)October - November 2005• Completed analysis of the Entry, Descent and Inflation (EDI) sequence (Task 1)• Sized the cameras for the conceptual payload (Task 1)• Began analysis of the gondola payload power budget (Task 1)• Began defining the gondola payload mass budget (Task 1)• Modeled BGS performance with a numerical model (Task 2)• Analyzed optimal design of the DARE platform with a numerical model (Task 2)• Simulated balloon trajectories at Mars (Task 2)• Started cost analysis by defining mass budget for the conceptual mission scenario (Task 3)• Participated at Mars Exploration Program Advisory Group (MEPAG) and Venus ExplorationAdvisory Group (VEXAG) meetings (Task 4)• Established new contacts with planetary scientists and NASA managers (Task 4)• Submitted our seventh bimonthly report (Task 5)December 2005 - January 2006• Continued analysis of the mission power budget (Task 1)• Continued sizing the solar panels and the batteries (Task 1)• Studied effects of the vertical wind gusts on the DARE platform performance (Task 1)• Estimated rates of buoyant gas loss for DARE platforms (Task 1)• Studied performance of the BGS with non-spherical balloons and larger wings (Task 2)• Developed guidance techniques that optimize guidance capabilities (Task 2)• Completed simulation of balloon trajectories for conceptual Surface Network EmplacementMission at Mars (Task 2)• Studied existing and available costing models to be used in Task 3 (Task 3)• Briefed Dr. Samad Hayati of JPL (Manager, Mars Technology Program) on DARE concept(Task 4)• Search for partners for potential Mars Balloon Scout proposal (Task 4)• Submitted our eighth bimonthly report (Task 5)February - March 2006• Sized solar panel and battery for conceptual DARE platform design (Task 1)• Developed technology development roadmap (Task 1)• Researched BGS performance at Venus (Task 2)• Outlined unique DARE observations (Task 3)• Contacted ESA groups involved in Mars and Venus aerobot studies (Task 4)• Submitted a technology summary on Venus BGS to the Venus ESA Cosmic Vision Missionworking group (Task 4)• Prepared and submitted our Final report (Task 5)2.2.1 Task 1 Conceptual Architecture Design and Technology IdentificationWe have reviewed a number of recent documents on Mars exploration objectives and formulateda list of general scientific objectives for the DARE architecture. These objectives are: Marssample return assist; search for present or past life habitats (hydrothermal sources, atmospheric19

methane mapping); characterization of the water, atmosphere and climate on Mars in the pastand today; and characterization of the structure and evolution of Mars.We have also developed a list of potential observations in support of these objectives. We haveanalyzed instrumentation that would be able to support these observations and producerevolutionary science results. For example, a high-resolution camera will enable imagery with aresolution on the order of 1 centimeter per pixel – comparable to the resolution of the imagery ofthe Mars Exploration Rovers - but virtually of any place on the surface of the planet.We have sketched several conceptual mission scenarios in support of the proposed observationsand measurements. These include: Mars Sample Return Assist, Emplacement of SurfaceNetworks, Crustal Magnetic Anomalies Mapping, Search for Origin of Atmospheric Methane,Polar Night Studies. Emplacement of networks of surface seismological and meteorologicalstations was chosen for detailed analysis and for conceptual design effort. The detailed analysisof the mission scenario included simulations of the atmospheric trajectories of the DAREplatforms and analysis of the energy and power budget during a mission.We have performed a number of design trade studies, estimating potential payload masses asfunctions of altitude, potential stresses experienced by the balloon films, changes in theplatform’s altitude due to deployment of probes and others. These studies were based on theassumptions of the development of various enabling and enhancing technologies. Thesetechnologies include: advanced balloon envelopes; entry, descent and inflation techniques;inflation hardware; etc. Other enabling and enhancing technologies that were researched duringthe Phase II study included balloon guidance system, power generation and energy storage, andothers. We have developed a roadmap that outlines the necessary further development of theenabling technologies and technological milestones of the concept.We have developed a conceptual architecture design in the context of future NASA Marsexploration. We chose the Emplacement of Surface Networks mission to focus the design effort.The design effort focused on the DARE platform and assumed existence of a communicationorbiter in orbit about Mars. We have estimated masses and dimensions of the main componentsof the DARE platforms: the balloon; gondola structure; BGS wing, tether and winch; solarpanels; batteries; support instrumentation (computers, communication, thermal protection),science instrumentation, etc. Fall speeds of the deployable packages that comprise the nodes ofthe surface network were studied and design approaches to mitigate g-loads upon landing aresuggested. Simulated atmospheric trajectories were used to size the solar panels and the batteries.2.2.2 Task 2 Balloon Guidance ResearchWe have developed advanced mathematical models of the Single- and Dual-Wing BalloonGuidance Systems. These models were used to study the performance of the BGS and tosimulate atmospheric trajectories of the DARE platforms for conceptual mission scenarios.We have modified our existing balloon trajectory simulation software to use the Mars GlobalReference Atmospheric Model 2001 (Mars-GRAM 2001) and Venus Global ReferenceAtmospheric Model (Venus-GRAM 2005) as the simulation environments. This allowed us tostudy BGS performance at Mars and Venus, and to simulate atmospheric trajectories for the20

DARE platforms in a variety of realistic Martian environmental conditions. We have studiedseasonal, diurnal, altitudinal and geographic variations in the Martian atmosphere and analyzedBGS performance in these varying atmospheric conditions.We have developed techniques that maximize the control capabilities of the BGS. Various modesof BGS operation can be employed to achieve optimal performance: reeling the BGS to thealtitude of the maximum relative wind; forcing the BGS wing into a stall to employ it as a dragsurface; employing the lift of the BGS wing.Numerical analysis indicates that the BGS wing on Mars will operate in an aerodynamic regimecharacterized by a very low Reynolds numbers (of the order of thousands). This does not presenta problem for the BGS performance. We have studied special low-Reynolds-number aerofoils tooptimize BGS performance. This analysis was used in the BGS conceptual design effort.We have analyzed BGS performance with non-spherical (elliptical) balloons. While nonsphericalballoons have higher weight than spherical balloons of equivalent volume, thereduction in balloon drag improves the BGS performance. An additional benefit in employing anon-spherical balloon is the possibility to use the lift generated by non-spherical balloons toimprove the BGS performance, but this possibility was not explored in this study.2.2.3 Task 3 Cost and Performance Benefits AnalysisWe have outlined unique observations enabled by the DARE architecture and describedperformance benefits of the architecture over conventional mission scenarios. We have studiedavailable costing models to use in estimating the cost of the conceptual DARE mission.However, due to resource limitations, we were unable to complete this Task and estimate thecost of the conceptual DARE missions for comparison with conventional missions.2.2.4 Task 4 Identify Pathways to Architecture RealizationIn Phase II we identified pathways for realization of DARE architecture and for continueddevelopment of the enabling technologies. We approached managers at NASA Mars Pre-Projects, Advanced Studies Office Advanced Planning, and at Mars Technology Program at JPL,submitted a white paper to the NASA Robotic & Human Exploration of Mars StrategicRoadmapping Committee, briefed NASA JPL Mars Advanced Studies Group, and participated inthe meeting of the Mars Exploration Program Analysis Group (MEPAG) and Venus ExplorationAnalysis Group (VEXAG).Due to the ongoing budget cuts in the Mars Exploration Program we believe that DARE concept,being a low-cost concept, has a good chance of being noticed by the NASA planners andmanagers. We continue to seek opportunities to get involved in the process through which newtechnology can be inserted into the Mars Program.Interest in Venus exploration is picking up with ESA’s Venus Express due to arrive at Venus inApril of 2006, NASA’s Messenger flyby of Venus in June of 2007, and JAXA’s Venus ClimateOrbiter (Planet-C) planned to be launched in 2010. NASA is at the early stages of developing aroadmap for Venus exploration. There is a lot of interest in balloon platforms for Venus21

exploration within the Venus exploration community and within NASA. We continue to seekopportunities to promote new technologies identified in this study in the context of Venusexploration.2.2.5 Task 5 Planning and ReportingGAC participated in conferences, prepared technical papers, prepared and participated in abriefing to NASA JPL. These are listed below:Technical Papers“Sailing the Planets: guided balloons for suborbital operations at Mars, Venus and Titan”, awhite paper submitted to Robotic Access to Planetary Surfaces subcommittee of the NASARobotic & Human Exploration of Mars Strategic Roadmapping Committee.“Sailing the Planets: Planetary Science from Guided Balloons”, paper submitted to 17 th ESASymposium on European Rocket and Balloon Programmes and Related Research, Sandefjord,Norway, May 30-June 2, 2005.“Sailing the Planets: Exploring Mars from Guided Balloons”, AIAA 2005-7320, AIAA 5thAviation, Technology, Integration, and Operations Conference (ATIO), 26-28 September 2005,Arlington, VA.Conferences6 th Annual NIAC meeting in Seattle, WA, October 19-20, 2004.Fall Meeting of the American Geophysical Union (AGU) in San Francisco, CA, December 13-17, 2004.17 th ESA Symposium on European Rocket and Balloon Programmes and Related Research,Sandefjord, Norway, May 30-June 2 2005.AIAA 5th Aviation, Technology, Integration, and Operations Conference (ATIO), 26-28September 2005, Arlington, VA7 th Annual NIAC meeting in Denver, CO, October 10-11, 2005.Mars Exploration Program Analysis Group (MEPAG) meeting #14, Monrovia, CA, November2-3, 2005.Venus Exploration Analysis Group (VEXAG) meeting #1, Pasadena, CA, November 4, 2005.BriefingsPresentation to the Mars Advanced Studies Group, NASA JPL, May 12, 2005.Presentation to Dr. Samad Hayati, Mars Technology Program, NASA JPL, December 21, 2005.22

GAC hosted a NIAC site visit on August 8, 2005 at its world headquarters in Altadena, CA. Inaddition, we have submitted 8 bi-monthly reports, a midterm report and this final reportsummarizing the technical progress of the work to NIAC.Media Interviews, Articles and Press ReleasesGAC issued a press release on the development of the DARE concept on September 27, 2005.The February 14, 2006 issue of New Scientist.com published an article on planetary ballooningbased in part on the interview with Dr. Pankine (“See Mars and Venus by balloon”, 14 February2006, Kurt Kleiner, Magazine issue 2538).Dr. Pankine was interviewed by J. Papalardo, Associate Editor of the Smithsonian Air and SpaceMagazine. The article is scheduled to appear in the June/July 2006 edition of the magazine.Dr. Pankine was interviewed by Leonard David, writer for the popular science web siteSpace.com.23

3 DARE Architecture Exploration Capabilities3.1 IntroductionThis section describes our work in Phase II on defining the new science and explorationcapabilities enabled by the DARE architecture. In this section we provide the DARE architecturedesign summary, review the exploration capabilities enabled by DARE architecture, review theMars Exploration Program objectives and how DARE fits within NASA’s vision of Marsexploration. We discuss potential DARE-enabled observations and measurements, and theinstruments that could be used in these observations. We conclude by discussing potentialmission scenarios.3.2 DARE Architecture Design SummaryThis section summarizes our vision of the DARE architecture. This section is included to providecontext for the discussions in the following chapters of the report.DARE architecture is comprised of multiple Mars exploration assets: landers, rovers, satellitesand guided long-duration flight balloons. The guided balloons or Directed Aerial RobotExplorers (DARE) platforms, as we shall call them, occupy a central role in the architecture.DARE platforms enable new scientific observations (both remote and in situ) and tie togetherobservations from different vantage points of satellites and in situ platforms (landers and rovers).The architecture creates a “virtual planetary scientist” – it is capable of providing a wealth ofinformation on the state of a planetary surface, atmosphere and interior at a wide range of spatialand temporal scales.The unique capabilities of the “virtual planetary scientist”, such as the ability to take a very close“look” – on a 1-10 cm scale – at geological formations at virtually any place on the surface of theplanet, and augment this first “look” with in situ analysis – is made possible by the DAREplatforms. The balloon guidance system enables observations of specific targets across theplanet. The proximity to the surface of the planet enables observations with very-high resolution.Small droppable science sondes enable in situ observations through the atmosphere and on thesurface.The DARE platforms will stay aloft for one to several Martian years, at altitudes as high as 20km above the surface. The superpressure balloons can be fairly large – from 30 to 70 m indiameter – to float relatively heavy scientific payloads weighing between 50 to 150 kg. Ourconceptual mission scenario employs a 34-meter diameter balloon with 90 kg of payload. Theballoon envelopes will be made of a very lightweight composite material that would providesufficient strength for the envelope and provide sufficient gas barrier to reduce diffusion of thebuoyant gas. The strength of the envelopes will eventually be provided by a material woven outof carbon-nano-tubes (CNT), making the envelope extremely light and very strong. Advances inmanufacturing balloon films will lead to a significant reduction in buoyant gas leakage throughthe seams of the balloon envelope. The balloon’s envelope will be aluminized at the top andpainted white at the bottom to reduce the possibility of atmospheric CO 2 condensations in the24

polar atmosphere. New materials with variable optical properties may become available in thefuture. A balloon made out of such materials would be able to change the reflectivity of itsenvelope depending on the environment and, in this way, to allow a reduction in stress due toreduced diurnal variations of the film and buoyant gas temperatures. This would enable areduction in the mass of the envelope due to less stringent envelope strength requirements.Advanced lightweight power generation and energy storage systems onboard DARE platformswill enable operations in the very cold atmosphere of the Polar Regions during polar nightconditions. Excess heat can be used to warm up the buoyant gas and the envelope film to preventatmospheric CO 2 condensation that is undesirable. Power-demanding science instruments, suchas powerful lasers and radars, could be carried on board.The DARE platforms will communicate with Earth and with the other elements within thearchitectures using a dedicated communications orbiter. The existence of the telecommunicationsorbiter in orbit around Mars will make possible the use of small and lightweight antennas(similar to those of the Mars Exploration Rovers) on DARE platforms. The orbiter could alsoprovide the means for communications with microprobes and deployed surface stations.The Balloon Guidance System (BGS) will provide capabilities to steer the DARE platformvirtually to any location on Mars and to perform targeted imaging or precise delivery ofdroppable surface probes. The wing of the BGS will be suspended on a 5 to 15-km-long tethermade out of Zylon treated with a UV-resistant agent. The BGS will employ an inflatablelightweight wing. Eventually, lightweight and very strong materials incorporating CNT might beused in the fabric of the inflatable BGS wing. Due to the high strength of the tether material, thelow weight of the BGS wing and the low gravity of Mars, the tether will have a cross-section ofjust a fraction of a millimeter and weigh just a few kilograms. Special airfoils will be used in thedesign of the BGS to optimize performance in the low-Reynolds number environment.The science payload of the DARE platform could consist of a multitude of instruments. It couldinclude high-resolution imaging cameras and spectrometers, laser altimeters for observations ofthe surface, radar, magnetometers, gravimeters for studies of the Mars interior, in situatmospheric chemistry instruments and wind Lidars for in situ atmospheric studies. To reducethe masses of the optical instruments they will be housed in the supporting structure of the BGSwing, just a kilometer or two above the surface. Attitude sensors and actuators will point thecameras to the surface targets. The cameras will be also used for platform autonomousnavigation and guidance.In addition to these instruments, the DARE platforms will carry dropsondes and deployablesurface probes that would be used for atmospheric profiling and surface in situ studies. Multiplelightweight dropsondes (10-100 g) will be used to get atmospheric profiles of pressure andtemperature, winds, and atmospheric constituents (dust, aerosols, water, etc). The deployablesurface probes would include surface stations for networks of seismological and meteorologicalnetworks; miniature geo-chemistry, mineralogy, age dating, and bioactivity detectinglaboratories; miniature rovers with limited roving range and a suite of simple instruments (e.g. amicroscope, a drill); etc. Seismological and meteorological networks would be emplaced on thesurface of Mars on a global scale. The mini-labs will be dropped with pinpoint accuracy to studyhydrothermal vents and to search for life, to determine the age and mineralogical composition of25

interesting geological deposits. Small rovers could be dropped in the vicinity of an interestingregion and perform a limited site survey (possibly focusing on just a single science task). DAREplatforms would have the ability “to look” at the layering on the sides of the cliffs at Mars – sitesnot visible from space due to observational geometry.New scientific observations and approaches will become possible in the context of the DAREarchitecture. An orbiter may detect a surface target with intriguing properties with its lowresolutioncamera. A DARE platform can be dispatched to the target to perform high-resolutionimaging and/or deploy a droppable probe or a small rover for a closer look at the target.In another scenario, an orbiter may detect elevated concentrations of water and methane in theatmosphere. A DARE platform can be dispatched to the location to provide high-spatialresolution detection and verification of the orbiter signal. The DARE platform may be able tofollow the trace amounts of gases in the atmosphere to their source on the ground. Thisimmediate response to discovery avoids waiting for several years to prepare and launch adedicated mission.DARE platforms can calibrate global orbital datasets by providing high-resolution data in thecontext of the low-resolution orbital data. Combined with the in situ data from rovers anddeployable science probes, the datasets can be calibrated at different levels of resolution andenable a deeper understanding of the existing and future global datasets created by orbiters. Inthis way DARE architecture represents a practical implementation of the recently proposedmulti-tiered planetary exploration concept 1 .The DARE architecture also enables new approaches to sample return. For example, the DAREplatform could visit sites where large rovers work for months collecting samples that would bethe most interesting to the scientists on Earth. These sample collection sites could be scatteredacross the surface of Mars and include sites in the vicinity of the polar terrains, volcanic terrains,outflow channels, old terrain, young terrain, etc. The DARE platform would fly over these sitesand collect the samples by deploying a collection vehicle or snatching the samples in specialcontainers from the ground, in the manner that mail was collected from trains in the past. TheDARE platform would then deliver and drop these samples to a single site where a rover wouldcollect them from the ground and place them into the return vehicle for the flight to Earth. In thisapproach samples from many sites on Mars could be returned with a single return vehicle savingan enormous amount of resources and time.3.3 DARE Science Objectives and Potential ObservationsThe DARE architecture is built around a DARE platform – a very versatile platform at a uniquevantage point below the orbit altitude and above the surface of a planet. Science objectives for1 W. Fink et al., “Next-generation robotic planetary reconnaissance missions: A paradigm shift”, Planetary and Space Science, inpress26

Mars balloons were formulated in the recent past 2,3 . A wide range of unique observationscovering a range of NASA objectives can be enabled by the DARE architecture:- in situ analysis of the atmosphere on a global scale (altitudinal profiles, isotopic ratios);- high spatial resolution imaging (1-10 cm) of the surface and subsurface on a global scale(visible and IR cameras, spectrometer, radar, radiometer, etc);- high-spatial resolution (1-10 km) magnetic and gravity surveys on a global scale;- emplacement of networks of surface stations on a global scale - seismological, geophysical(surface heat flux, soil properties, subsurface structure - magnetometers), meteorological (P,T, near surface altitudinal profiles, surface winds)- targeted delivery of surface probes on a global scale – penetrators, biological, mineralogical,chemical in-situ analyzers.DARE architecture enables these observations at Mars, Venus, Titan, Jupiter, Saturn, Neptuneand Uranus. The basic structure of the architecture and its capabilities remains the same for eachplanet, while the details of implementation vary. For example, while a superpressure balloon willbe used on Mars, Venus and Titan, a “hot air” Montgolfiere balloon would be more appropriatefor Jupiter, Saturn, Neptune and Uranus.In this Phase II study we further developed the DARE concept in the context of a Mars mission.As a starting point for the development of the conceptual mission scenarios, we assume that theDARE architecture at Mars can be enabled in the period 10-20 years from now. The scientificobjectives, the observations and measurements for the missions enabled by the DAREarchitecture at Mars in the second and third decade of the 21 st century will depend on theaccomplishments of Mars Exploration during the previous years.3.3.1 NASA’s Mars Exploration VisionBy the time the DARE architecture will be implemented, Mars Global Surveyor, Mars Express,Mars Odyssey and Mars Reconnaissance Orbiter will complete global reconnaissance of Marsand a more detailed analysis of several target sites. Mars Exploration Rovers, Mars ScientificLaboratory and Phoenix will complete detailed in situ geological, geophysical, mineralogical andother studies of four regions of the planet (including one polar site).NASA envisioned several pathways of exploration for the next decade depending on theoutcome of the missions in the current decade. These pathways are listed in Table 3-1.2 Mars 2001 Aerobot/Balloon System Overview, AIAA 97-1473 R. Greeley et al., The Mars Aerial Platform (MAP) Concept, AIAA 96-0335, 1996.27

Table 3-1. Mars Potential Next Decade Pathways 4PathwayLine of Scientific InquirySearch forEvidence ofPast LifeExploreHydrothermalHabitatsSearch forPresent LifeExploreEvolution ofMars• Science from First Decade missions plus Ground Breaking SampleReturn confirms ancient Mars was wet and warm.– Locating and analyzing water-laden sedimentary rock is primarygoal.– Pathway includes search for evidence of past life.• Exploration in First Decade discovers hydrothermal deposits (active orfossil).– Probability of hydrothermal regions being discovered is potentiallyhigh.– Hydrothermal habitats are focus of second decade of Marsexploration.– Potential for discovery of evidence of past and present life isgreatly improved.• Commits to search for present life at sites determined to be modernhabitats by First Decade missions.– Search for life at active hydrothermal deposits or polar caps.– Path would be taken only following a revolution in policy and/orprogrammatic interest in Mars.– MSR with mobility is included as only reliable, validatible meansof detecting of life.• Science of First Decade determines that Mars was never globally wet.– Determine the loss mechanisms and sinks for water and CO2 overtime.– The terrestrial planets evolved very differently, much more so thanwe had thought. Why?– Were the initial conditions on Venus, Earth, and Mars similar orvery different?Each of these pathways could be followed by a sequence of potential missions in the nextdecade.The bulk of observations from the current Mars missions (MERs, Mars Express, Mars Odyssey)suggest the existence of the surficial water and persistent ground water in the Martian past. Thisdiscovery in turn suggests Pathway #1 (Search for Evidence of Past Life) and possibly Pathway#3 (Search for Present Life) as the preferred pathways of Mars exploration. The goals ofNASA’s exploration of Mars will remain Life, Climate, Geology, and Preparation for HumanExplorers – with Life being “first among equals”.4 Mars Exploration Strategy 2009-2020, Mars Science Program Synthesis Group (MSPSG), Ed. D. J. McCleese, NASA JPL,2004.28

Figure 3-1 illustrates the timeline for the Mars Exploration Program (MEP) for the next decades.Note, at the time of this writing the Mars Exploration Program is experiencing budget cuts,which could stretch out or change the plan.Figure 3-1. Mars Exploration Program timeline for the next decade and beyond 5 .3.3.2 DARE and the NASA’s VisionHow does the development of the revolutionary DARE architecture fit into this vision? Thefollowing answers are possible:- Technology development for the DARE architecture leads to a Scout mission;- Elements of the DARE architecture replace or augment some of the major potential missionsin Figure 3-1 – saving time and resources, enabling new science;- Missions up to 2020 emplace the elements of the DARE architecture at Mars. DAREarchitecture is operational at Mars after 2020;- DARE architecture enables new pathways of exploration.3.3.3 DARE Science Objectives and Possible ObservationsAfter analyzing Mars exploration objectives 6,7 , reviewing exploration plans and the past Marsballoon concepts 2,3 , and discussions with planetary scientists we have formulated the followinggeneral science objectives that DARE can directly address or support:- Mars sample return assist5 MEPAG.6 Space Studies Board, National Research Council, New Frontiers in the Solar System, National Academies Press, 20037 Space Studies Board, National Research Council, Assessment of Mars Science and Mission Priorities, National AcademiesPress, 200329

- Search for present or past life habitats- Characterization of the water, atmosphere and climate on Mars in the past and today- Structure and evolution of MarsIn the context of these scientific objectives and in the context of the observations accomplishedby the current and upcoming missions (where applicable) the DARE architecture will enable thefollowing observations and applications:- Collection of samples from different remote sites on a global scale for delivery to the Earthreturningspacecraft (Section 3.5.3);- Mapping of the atmospheric methane and search for its origin (measurements of fraction ofmethane isotopes in the Martian atmosphere) (Sections 3.5.2);- Pinpoint delivery of micro-labs and mobile micro-explorers to the sites that may harbor lifetoday or may have have harbored life in the past;- Emplacement of global networks of surface stations (see Section 3.6);- Measurements of the crustal magnetic field from sub-orbital altitudes (see Section 3.5.1);- Gravity field measurements from sub-orbital altitudes – crustal structure on the scale of 10km;- High-resolution visible (1-10 cm) and infrared (10-100 cm) imaging: polar layered terrain,layering in crater/canyon walls origins of outflow channels, dichotomy boundary, waterrelatedsurface morphology; aqueous minerals; sources of water on the surface (hydrothermalvents, ice); small craters (10 cm – 10 m); boulders and large rocks;- Atmospheric water and CO 2 cycle (Section 3.5.4);- In situ rock and soil age dating;- In situ chemical analysis of atmospheric constituents;- In situ measurmenets of the atmospheric winds and vertical temperature-pressure profiles;- In situ and remote (Lidar) wind measurements on a global scale;- Radar soundings of the subsurface to provide high-vertical resolution and high signal-tonoisestructure of the underground;- High-resolution laser altimetry to quantify the surface topography and morphology on asmall scale;- Cooperative observations.This is obviously just a small subset of the total spectrum of possible observations. The planetaryscience community can suggest other observations in the future. Some of the suggestedobservations are described in more detail below.3.3.3.1 Fractionation of Methane Isotopes in the Atmosphere of MarsA DARE platform at Mars can measure fractionation of methane isotopes in the atmosphere ofMars using a Tunable Laser Spectrometer (Figure 3-2).30

Figure 3-2. Tunable Laser Spectrometer (TLS) for Atmospheric and Sub-surface gasmeasurements on Mars (NASA JPL).The measurement would look at the C 12 /C 13 ratio in the methane. This measurement can provideclues to the presence of life at Mars. Methane-making organisms discriminate between isotopesas they feed on a global reservoir of CO 2 . The lighter isotope of CO 2 – C 12 O 2 is being consumedby these organisms, while the heavier isotope C 13 O 2 is avoided. If the ratio of C 12 /C 13 in methanewere different from that in CO 2 , it would offer strong evidence for a biological source.By enabling this measurement on a planet-wide scale and at very high accuracy, DARE enablesthe search for biological sources. By mapping the distribution isotopic ratios on a global scale,gradients of concentrations and the locations of their surface or subsurface sources can beestimated. A single surface lander would produce a single measurement, which will not enablelocating the source.TLS is a very small and light instrument, as can be seen from the Figure 3-2. Several of theseinstruments can be positioned along the tether and on the gondola, providing measurements ofvertical profiles of the isotope fractionations. Alternatively, the TLS can be a part of adeployable probe and make measurements of the isotope fractions while falling through theatmosphere after being deployed from the gondola of the DARE platform.3.3.3.2 Cooperative ObservationsWe have developed a list of cooperative measurements – measurements done in cooperation withthe other platforms comprising the architecture. These are:• Validation of orbital observations:o High resolution characterization of subsets of orbital surface imagery datao In situ validation of atmospheric measurements• High-resolution traverse maps in support of a rover operations.31

• Multi-scale simultaneous atmospheric profiling from surface, atmosphere and orbit.• In situ characterization of the local atmospheric conditions in support of aerobraking andEntry, Descent and Landing activities.• Deployment of a surface probe over a target detected from the orbit.• Sample return assist.As an example of a cooperative observation, consider the recent discovery by the ESA MarsExpress orbiter of the “ice lake” near the North Pole of Mars (Figure 3-3). The bright whitematerial in the center of crater is residual water ice. Traces of the water ice can also be seen onthe slopes of the crater walls. The surface of the “ice lake” is about 200 meters above the floor ofthe crater. This difference cannot be due solely to water ice. Presumably, the water ice lake “sits”on top of the large dune field at the floor of the crater. Indeed, some darker material can be seenexposed at the edge of the ice cap.Figure 3-3. Water ice lake inside a crater on Mars (ESA).Were the DARE platform present at Mars at the time of this discovery, it could have beendirected to fly over this feature and to deploy a surface probe to study water ice and observeatmospheric conditions at the crater floor. In this way scientists would be able to compareimagery from orbit, from the DARE platform, and from the surface probe to make conclusionsabout the origin of this patch of water ice and regarding the possibilities that life may beharbored underneath it.Another example of a cooperative mission is the Sample Return Assist, which is described inmore detail in Section 3.5.3.32

3.3.3.3 High-Resolution Imaging of the Martian Surface FeaturesThere are many targets for high-resolution imaging on Mars – here we describe just a smallsubset of possible targets for observation.Figure 3-4. Layers exposed by erosion on a slope in the Martian South Polar Region.Extensive layering in the polar terrains at Mars (Figure 3-4 and Figure 3-5) is attributed tomultiple episodes of ice and dust deposition. The duration of each deposition episode, the totaltime period over which the layered terrain formed or the mechanism of formation of the layersare not known. Thickness of the layers varies from several to 100 meters (Figure 3-5). Highresolution(at cm/pixel range) observations of the layers’ morphology and thermal properties willhelp to answer many question pertaining to Mars climate history, the atmospheric and surfacewater cycle and the role of dust in regulating Martian climate. In addition to high-resolutionimaging afforded by the DARE platforms, the DARE platforms will also enable better viewinggeometry by looking at the sides of the layered deposits, while satellite observations observethese formations from above.33

Figure 3-5. Analysis of the layers thickness in the polar layered terrain. 8The geological history of Mars written in the Martian rocks is hidden from view by the layer ofsurface dust. The Martian surface looks almost the same everywhere across the planet to orbitalinstruments due to the cover of dust. Only in a few places where the dust is blown off the surfaceby the winds orbital instruments can peer at the exposed rock.Walls of craters, channels, cliff edges or depressions are the places where surface and subsurfacerocks are exposed. In addition, depressions expose the subsurface of Mars, allowing theobserving of the geological sequences with changing depth. Walls of craters and depressionsshow layering that indicate complex geologic history of Mars (Figure 3-6).Observing the layers and exposed rock in canyon walls at very high resolution (1 cm) will enableresolving the morphology of these features. High-resolution spectral observation will distinguishminerals in each layer or formation. Geological maps of Mars will be refined and redrawn on afiner scale enabling deeper understanding of the geologic history of Mars.8 Lori K. Fenton, Ken E. Herkenhoff, “Topography and Stratigraphy of the Northern Martian Polar Layered Deposits UsingPhotoclinometry, Stereogrammetry, and MOLA Altimetry”, Icarus 147, 433–443 (2000), doi:10.1006/icar.2000.6459.34

Figure 3-6. Gullies and layers in the wall of a depression located on the floor of the Rabe Crater(NASA/JPL/Malin Space Science Systems)35

Figure 3-7. Outflow channels on MarsFigure 3-8. Details of the Ares Vallis outflow channel (ESA Mars Express).High-resolution observations of the surface morphology at the sites of the origin of the outflowchannels at Mars (Figure 3-7 and Figure 3-8), the details of the channel walls, the existence orabsence of aqueous minerals within the channels will shed light on the origin of the channels, themechanism of their formation and on the possibility of the existence of habitable environmentson early Mars.36

Figure 3-9. Martian topography.One of the striking features of Mars is the difference between the surface topography in theNorthern and Southern hemispheres. The Northern hemisphere of Mars is several kilometerslower than the Southern hemisphere (Figure 3-9). The southern hemisphere appears to be mucholder and contains many more craters. The reasons for such a dichotomy are not clear at present.Observing the dichotomy boundary (Figure 3-10) at very high resolution may provideexplanations about the mechanisms that shaped the Martian hemispheres.Figure 3-10. Mars hemispheric dihotomy boundary as viewed by Mars Express (ESA)3.3.3.4 Pinpoint Delivery of Surface ProbesDARE platforms will be able to deliver surface probes with a pinpoint accuracy at Mars, whilepinpoint delivery of small probes from space presents significant difficulties. Figure 3-11demonstrates the scales of Martian geological features compared to the landing accuracy of theexisting surface probes.37

Figure 3-11. Olivine outcrop and DS 2 landing ellipse (NASA/JPL/ASU).The figure shows a map of Martian surface mineralogy developed with the THEMIS instrumentonboard Mars Odyssey orbiter plotted over an image of Martian topography. Purple streaks onthe bottom of the canyon near the walls that run along the length of the canyon are the olivineoutcrops. The existence of the olivine at the surface of Mars is significant, because olivine isquickly destroyed in the presence of water. Here it demonstrates the scales of importantgeological features seen on Mars. The yellow ellipse is the landing error ellipse for the DeepSpace 2 (DS 2) surface probes. It is 120 km long and 30 km wide. It is clear from Figure 3-11that DS 2 probes would be unable to land on the geological feature of the spatial dimensioncomparable to that of the olivine outcrop. The DARE platform can fly over the region of thetargeted geological formation and deliver the surface probe with pinpoint accuracy.3.3.3.5 Radar Sounding of the SubsurfaceThe ESA Mars Express orbiter carries a Martian radar (MARSIS) that recently began returningstunning images of the Martian subsurface (Figure 3-12). The subsurface structure seen in theMARSIS data on the above figure are interpreted as a buried impact basin that may be filled withwater ice. DARE platforms carrying radars much closer to surface will be able to distinguishvery small subsurface features and aid in the search for subsurface water and life habitats.38

Figure 3-12. MARSIS “radargram” showing echoes obtained from an approximately 250 kmdiameter circular structure in the subsurface of Mars (ASI/NASA/ESA/Univ. of Rome/JPL)3.3.3.6 Human Mission Precursor ObservationsIn the context of the precursor to a human mission to Mars the DARE architecture will enablethe following observations:- High-resolution characterization of hazards (large rocks, terrain slopes, wind gusts) at thepotential manned landing sites;- Precise emplacements of surface radio beacons for targeted landing of manned spacecraft;- In-situ, high-resolution search for usable resources: delivery of miniature geochemicallaboratories; sounding radars (subsurface water ice); atmospheric abundances (H 2 O);- Monitoring atmospheric conditions to improve accuracy of aerobraking and landing.DARE platforms would float over selected landing sites and image the sites at very highresolution. These high-resolution observations would provide information on the presence ofhazards at the landing sites – large rocks, terrain slopes, - that may interfere with landing. DAREwould also deploy atmospheric profiling probes while over the landing sites to characterizeatmospheric winds. DARE would return to the site selected for manned landing and deploy radiobeacons that would be activated right before the arrival of the manned spacecraft and guide thecraft to the select site.39

DARE platforms would assist in selection of landing sites by identifying surface sites that haveresources that can be used by the crew of the arriving manned spacecraft. Using a sounding radarDARE may identify sites that have shallow water ice deposits that would be accessible toastronauts, or, using deployable in situ probes, it may determine that the soil has highconcentrations of specific chemical elements that can be extracted and utilized (as fuel or asource of breathable oxygen) by the human crew.DARE platforms can be relocated to the vicinity of the landing site at the day of the arrival of themanned spaceship and monitor atmospheric conditions in the region. This information can beused to correct the landing sequences or abort landing due to an unexpected atmosphericanomaly (very strong wind gusts or wind shear layer at some altitude above the surface). Theobservations would improve the safety of the manned mission.3.4 Potential InstrumentsOne of the goals of Phase II of this study was to investigate the potential of various instrumentsto enable revolutionary science from DARE platforms. In view of the required measurementsand objectives described in Section 3.3 we have considered the following instruments:• High-resolution and medium-resolution (context imaging) cameras• High-resolution thermal emission spectrometer (TES)• Radar• Magnetometer• In situ atmospheric chemistry instruments• Surface geo-chemistry, mineralogy, age dating, and bioactivity detection deployablepackages• Surface seismological/meteorological stations• Laser altimeter• Laser-Induced Breakdown Spectroscope (LIBS)While some of these instruments have a reach heritage (e.g. cameras), others are at the earlyconceptual stages of the development process (e.g. age dating deployable packages). Wherepossible, we made projections of the instrument mass, power requirements and other limitingfactors.The next sections discuss some of the instruments in more detail.3.4.1 High-Resolution ImagingHigh-resolution imaging with resolutions of 1 to 10 cm in the visible part of the spectrum andwith resolutions of 10 cm to 10 m in the infrared are needed to advance our understanding ofMars and aid in search for past and present life in the epoch after Mars Reconnaissance Orbiter(beyond ~2020). As it turns out (see Section 5.4.4) to achieve these high-resolutions from 10 kmaltitude (nominal floating altitude of the DARE platform) requires very massive and largecameras. By positioning the cameras on the BGS wing – about 1 km above the surface – itreduces the size and mass of the cameras, while achieving the required spatial resolution.40

The super-high resolution imaging that would be enabled with high-resolution instruments wouldrevolutionize our understanding of Mars. Recall, that MER imaging of the Martian geology onthe scales of centimeters and even millimeters (“blueberries”) allowed the MER science team tomake conclusions about the existence of a shallow lake at Opportunity’s landing site. If imageryof this resolution becomes available for virtually any site on Mars, the scientific outcome wouldbe truly revolutionary.Figure 3-13. High-resolution imagery from DARE platforms3.4.2 Laser-Induced Breakdown Spectroscopy (LIBS)LIBS technique is based on vaporizing a sample with a laser beam and analyzing the vapors witha spectrometer. A ChemCam instrument proposed for the Mars Science Lander (MSL, 2009)would ablate surface coatings from materials at up to 10 meters and measure elementalcomposition of underlying rocks and soils (Figure 3-14).Recent reports 9 indicate that the range of LIBS operations can be extended to 1 km. The requiredlaser is a Terawatt laser using 250 mJ of energy. The impulse duration is in 80 fs.If the range of the LIBS instrument can be extended to 10 km, then LIBS can be flown as part ofthe DARE gondola payload. If the LIBS instrument can be made very small and lightweight itcan be placed on the BGS wing (closer to surface).A LIBS instrument on a DARE platform9 Applied Physics Letters, 85, 3977, 200441

would enable planet-wide remote identification of rock and soil elemental composition from aballoon.Figure 3-14. ChemCam on MSL. Can LIBS be flown on a Mars balloon?3.4.3 Optical Stimulated Luminescence (OSL)The greatest obstacle to unlocking and interpreting the geologic and climatic record on Mars isthe need for absolute age dating.OSL is an in situ technique for age-dating recent sediments – up to 1 Myr old. Luminescencedating is based on solid-state properties of mineral grains that allow them to record theirexposure to radiation. The recorded radiation exposure can be measured by stimulating thesample with light of one wavelength and monitoring the emitted luminescence in anotherwavelength (optically stimulated luminescence, OSL). The intensity of luminescence is afunction of the absorbed natural radiation dose.The technique is well established for Earth. Low cost, simplicity, and potential forminiaturization make OSL more feasible than isotopic methods for age dating on Mars.An OSL instrument can potentially be incorporated into DARE deployable probes.3.4.4 Deployable ProbesDeployable probes are small vehicles that enable bringing the science payload in contact with thetarget when remote observations are inadequate. Small Deep Space 2 dropsondes (DS 2, Figure3-15) were deployed at Mars (unsuccessfully). Larger probes were proposed for deployment of42

surface networks of seismological and meteorological stations.(ESA’s Netlander 10 ). Smalldropsondes measuring atmospheric profiles or analyzing chemical compositions of the surface,can be carried by the dozens by DARE platforms and deployed over selected targets to augmentthe remote sensing capabilities of the platforms. In the future, probes dropped on the surfacewould be able to determine the age of the surface, look at the soil with high magnification,determine mineralogy of the soil and dig into the subsurface (see Section 5.3.11).Figure 3-15. Mars Microprobe 11 (NASA)Landing probes would impact the surface of Mars with relatively high velocities, even thoughthey would be deployed from a 10 km altitude and not from space. Figure 3-16 shows the fallvelocity of a probe deployed from a 10 km altitude on Mars as a function of altitude. Two casesare shown, one for a “thin” and the other for a “thick” atmosphere. Thin atmosphere has surfacedensity of 0.015 kg/m 3 and the pressure scale height of 11 km, while the thick atmosphere hasthe surface density of 0.03 kg/m 3 and the atmospheric scale height of 8 km. These twoatmospheres represent two extreme cases: a hot atmosphere during a dust storm and a coldatmosphere at the peak of the CO 2 sublimation from the polar caps.A small parachute (1-m in diameter) is proposed to decelerate the probes to impact velocitiesbetween 40 to 60 m/s. The time to impact is 150 and 200 seconds for the thin and thickatmosphere cases, respectively. Without the parachute the impact velocities would be similar tothose of the Deep Space-2 (DS 2) probes – 160-180 m/s. Deploying from a lower altitude (5 km)10 http://smsc.cnes.fr/NETLANDER/11 http://nmp.jpl.nasa.gov/ds2/43

educes the time to impact but does not reduce or increase the impact velocity, because theprobes reach the terminal velocity very quickly after deployment (within 2-2.5 km).Figure 3-16. Impact velocities of the ESN probes.3.5 DARE Conceptual Mission ScenariosWe have sketched several conceptual mission scenarios. These include:• Crustal Magnetic Anomalies• Search for Origin of Atmospheric Methane• Sample Return Assist• Polar Night Studies• Emplacement of Surface NetworksThe details of these conceptual mission scenarios are provided below. The Emplacement ofSurface Networks (ESN) mission was chosen for further analysis and for conceptual designwork.44

3.5.1 Crustal Magnetic AnomaliesFigure 3-17 shows the map of the crustal magnetic anomalies detected by the Mars GlobalSurveyor during the aerobraking phase. The enigmatic east-west linear structure of the crustalanomalies is apparent.Figure 3-17. Map of the Martian crustal magnetic field strength in the Southern HemisphereThe objective of this mission is to study the history of the Mars internal magnetic field and thecrustal accretion process, and the detection of subsurface water reservoirs.The mission would carry 1 or 2 vector magnetometers and attitude sensors positioned at thegondola and/or along the tether. Simultaneous high-resolution visible and infrared imaging of thesurface would be performed to link the measurements of the magnetic field and subsurface waterto topographic features.DARE platform would complete several “orbits” over southern highlands, producing multipletraverses of the anomalies. As a result, the measurements would enable characterization of thespatial frequency of the magnetic anomaly pattern, the depth, thickness and lateral extent of thesources. The relations of the magnetic features to topography would be established and thesubsurface water reservoirs will be mapped.3.5.2 Search for Origin of Atmospheric MethaneThe objective of this mission would be to determine the nature of the source of the methane inthe atmosphere of Mars – is the source biologic or not? – and to locate the source of methaneemission. To distinguish between biologically and abiologically produced methane, instrumentsonboard of the DARE platform would measure isotopic fractionation of the methane isotopes in45

the atmosphere of Mars using a Tunable Laser Spectrometer (TLS). More details on theproposed measurements can be found in Section 3.3.3.1. This measurement can provide clues onthe presence of life at Mars.The TLS instruments will be housed in the gondola of the DARE platform. Several instrumentsmay be needed as isotopes of the other atmospheric constituents may need to be measured.DARE platforms would also carry a magnetometer for simultaneous measurements of themagnetic field anomalies. Crustal magnetic anomalies at Mars may be indicative of preservationof subsurface structures that harbor methane-producing biota or outgas methane into theatmosphere 12 .The DARE platform would float over the southern highlands performing isotopic and magneticfield measurements. The preferred seasons for the observations is late southern spring (L s =200º -250º) or late southern summer (L s =330º-360º), to avoid the dust storm season. The platformwould circumnavigate Mars several times at different latitudes, mapping isotopic ratios andmethane concentrations. This would enable establishing gradients of methane concentrations inthe atmosphere and to locate the source of methane on the surface.3.5.3 Sample Return AssistThe objective of the mission would be to enable sample return from multiple sites on a singlereturn vehicle. This mission illustrates cooperative capabilities offered by the DAREarchitecture. Mission concept is illustrated on Figure 3-18. Mission profile is as follows:• Multiple rovers collect samples at different sites. The rovers are able to search vast spans ofterrain and to select samples for return to Earth.• The samples are packaged into a container and transferred to the DARE platform flying overthe sample collection sites. The actual transfer mechanism may involve a long line that isbeing snatched from the air by the DARE platform BGS system (see Figure 3-18) or otherapproaches.• DARE platforms with the sample return canister fly over the site of the Sample ReturnVehicle.• Sample canisters are dropped from the DARE platform, collected by a special rover andtransferred to the Sample Return Vehicle.• Sample Return Vehicle takes the samples to Earth.In this way several samples from distinct sites on the planet can be delivered to Earth on a singlereturn vehicle.12 M. Mumma, NASA GSFC Center for Astrobiology, private communication46

3.5.4 Polar Night StudiesFigure 3-18. DARE sample return mission concept.The objective of the Polar Night Studies mission would be to study polar winter processes – suchas CO 2 accumulation, formation of clouds, etc. The DARE platform will carry a thermalspectrometer to measure atmospheric and surface temperatures, deployable probes anddropsondes to sample the underlying terrain in situ.The DARE platforms will need to rely on non-photonic power sources, which may includeRTG’s or wind power approaches.The produced observations will validate CO 2 cycle models and provide insights into formation ofthe polar terrains.3.5.5 Emplacement of Surface NetworksThe objective of the mission is to establish a network of 4 meteorological/seismological stationson the surface of Mars. DARE platform deploys 4 surface stations of 5 kg each. It is assumedthat the stations remain operational on the surface of Mars for 1 Martian year.ESA considered delivering 4 surface stations equipped with seismometers, a meteorologicalpackage, and other instruments from orbit in the Netlander mission. The Netlander mission wascancelled. The surface package had a mass of about 20 kg, which included 6 kg of scienceinstruments. Figure 3-19 shows in green color the possible landing sites for Netlander 1313 Global Seismological Performance of Networks Proposed for NetLander, M. Knapmeyer, T. Spohn, J. Oberst, 200347

Figure 3-19. Green: Netlander possible landing sites determined by climate and altitudeconstraints (after Ferri, 2002).Figure 3-20. Several sets of landing sites proposed by NetLander Landing Sites Work Group.Figure 3-20 shows several sets of proposed landing sites for the Netlander mission. Thesenetwork configurations focus on studies of the Martian interior underneath the Tharsis uplift.DARE platforms will be able to deliver the surface stations anywhere (except over the very hightopography of Tharsis volcanoes), not just in the "green" areas limited by altitude constraints.This will improve the performance of the network and the interpretation of the data registered bythe network, as an almost optimal configuration of the network can be achieved.The mission profile is as follows:• DARE platform flies to pre-selected sites and deploys stations. Figure 3-21 shows thetetrahedron configuration for a surface network on Mars. In the tetrahedron network thesurface stations are separated by the same distance from each other (the colors on the Figure3-21 indicate the distance from a station with the dark blue being the farthest away from astation). The tetrahedron configuration optimizes interpretation of the seismologicalinformation.48

Figure 3-21. Tetrahedron network of stations on Mars• Probes fall to the surface; the aft body stays on the surface, while the probe separates fromthe aft body and penetrates in to subsurface.• Surface station remains operational for 1 Martian year, receiving power from RHU or solarpanels• DARE platform continues with imaging mission after delivering probes.The network will provide soundings of the interior of Mars and measurements of the atmospherictemperature, pressure, humidity, winds and maybe other parameters.Emplacement of the Surface Networks mission scenario is described in more detail in thefollowing section.3.6 Emplacement of Surface Networks Mission Scenario AnalysisThis section describes in detail one of the envisioned mission scenarios outlined in the previousSection. This conceptual mission is Emplacement of Surface Networks (ESN). This mission wasalso used to focus the design of the conceptual DARE platform.Emplacement of Surface Networks mission scenario was chosen for detailed analysis because itoffers an alternative to conventional implementation of one of the core missions in the NASAMars Exploration Roadmap – namely, Planetary Evolution and Meteorology Network (seeFigure 3-1). Thus, the results of the analysis of the mission scenario can be useful in inserting the49

enabling technologies outlined in the effort into the Mars exploration technology developmentprogram. In addition this mission scenario demonstrates the exploration capabilities of theDARE architecture – the ability to target locations on the planet to deploy surface stations ordropsondes and the global planetary reach of the DARE platforms. In the course of the missionthe DARE platform visits different geographic regions of the planet and operates under differentsolar illumination conditions, which enables studying solar power generation and energy storagecapabilities in different conditions.In the following sections we describe the mission scenario, starting from the atmospheric entryof the entry vehicle and aerial deployment of the DARE platform, and concluding with simulatedatmospheric trajectories. The technologies envisioned in this mission scenario are described inmore details in Chapter 5.3.6.1 Mars Arrival Date and SiteWe consider the 2013 Mars arrival opportunity for the analysis of the ESN mission scenario.This date corresponds to a possible Scout mission in the NASA Mars Exploration Roadmap (seeFigure 3-1). The arrival date of the lowest energy trajectory for this opportunity is around July13, 2013. This date corresponds to equinox at Mars – Ls=351º. The sun is directly above theequator of Mars and is moving into the northern hemisphere. The arrival opportunities in later (orearlier) years will not change our main conclusions significantly: they would correspond todifferent seasons at Mars and may require more (or less) energy to reach Mars.Without the detailed analysis of the arrival trajectories and geometries we assume that the sitesspanning latitudes from –50º to 50º can be targeted for entry in each opportunity (polar regionscan not be targeted at every opportunity). We have chosen to have the atmospheric entry in theNorthern hemisphere because of the favorable solar power generation conditions there at thisseason (Northern spring). We have also chosen to have the entry site in the high latitudes in theNorthern hemisphere – at 50º latitude. This latitude can be targeted at any arrival opportunity andis close to the latitude of one of the surface sites for deployment of the surface network station(60º latitude, Figure 3-21). In addition, this latitude is “safe” for operation of even an unguidedballoon– if the BGS of the DARE platform fails to deploy or function, the DARE platform canstill operate for some time as an unguided platform, without the risk of crashing into hightopography (see Figure 3-9).3.6.2 Atmospheric Entry, Deployment and Inflation (EDI)Entry, Descent and Inflation (EDI) are the sequence of events that occur over a short time-spanfollowing the arrival of the spacecraft carrying the DARE platform to Mars and before theplatform achieves its floating altitude in the atmosphere of Mars.In this section we describe the EDI sequence. This is intended to be somewhat analogous to theEntry, Descent and Landing (EDL) systems used for the recent Mars Exploration Rovers, butreplacing the landing phase by the inflation phase appropriate to a balloon system. The series offigures below illustrate the sequence of events that take place during EDI.50

Figure 3-22 shows the entry vehicle as packaged prior to entry. The entry vehicle is the same asin the Mars Pathfinder mission. Figure 3-22 shows containers for the drogue chute, main chute,and balloon envelope, the gondola, stowed BGS and the inflation hardware.Figure 3-22 Cutaway Entry VehicleFigure 3-23 is an artistic representation of the proposed EDI sequence. The entry and descentphases would be very similar to previous missions. A blunt conical heat shield reduces theenormous interplanetary velocity, using shock waves to convert most of the kinetic energy intoheating of the atmospheric gases. The heat shield itself still reaches a very high temperature, butthe entry phase lasts a relatively short time, and the heat pulse does not penetrate completelythrough the heat shield to the encapsulated vehicle inside before the heat shield is ejected. Priorto dropping the heat shield, the aft cover is released and supersonic parachute is deployed at aMach number of ~2 at altitudes between 13 and 28 km (depending on atmospheric conditions).This slows the descent rate significantly into the low subsonic range to the point the internalcomponents no longer need protection from the ram air and the heat shield can be dropped.As the system approaches terminal velocity on the parachute, the inflation phase commences.The gondola and attached equipment are lowered, using a stitched tear-away bridle to control therate of extension, and the weight is used to stretch out the balloon envelope vertically. Theballoon envelope is packaged in a protective sleeve that helps to control the envelope duringinflation. If the envelope was not constrained, it could form a “sail” in the descent winds thatcould create forces that could damage the envelope51

Figure 3-23. DARE EDI sequence.Once the balloon envelope is extended to its full vertical length, inflation from the top occurs.The weight below the balloon assists in the inflation process by keeping the balloon envelopeunder control and reducing the possibility of “sailing” in the descent winds. Therefore, theinflation equipment is mounted below the balloon, and will be jettisoned when the inflation ofthe balloon is complete. An inflation tube is used to carry the gas up to the initial inflation bubbletowards the upper end of the balloon.Once the balloon has reached a size comparable to the parachute, the drag on the balloon willapproach that of the parachute, making the parachute ineffective. At this point, the parachute willbe cut away. As inflation continues, and buoyancy increases, the rate of descent will slowsignificantly.Upon completion of inflation the now extraneous inflation hardware can be released, and thesystem will ascend towards an equilibrium float altitude, at which point the trajectory controlsystem can be deployed.Possible altitudinal profiles of the DARE platforms during deployment and inflation in differentatmospheric conditions are studied in Section 5.4.3 and are shown in Figure 5-28. The seasonand location of the atmospheric entry chosen in this study are close to the atmospheric conditions52

of those in case 9 in Table 5-6. The descent profile of the DARE platform for this season andlocation is close to that marked as COSPAR-short. The analysis in Section 5.4.3 indicates thatthe DARE platform will continue descending for the total duration of the EDI sequence. It willreach floating altitude of 8 kilometers in about 20 minutes after atmospheric entry.An alternative approach that was considered and rejected was to land prior to inflating theballoon. This would allow more time for inflation of the balloon. However, there are severaldisadvantages. Naturally, some kind of landing system would be needed, which would addsignificant unnecessary mass to a system that is intended to operate exclusively within theatmosphere. Furthermore, significant horizontal winds would likely interfere with the inflationprocess, and there is a significant danger of snagging the balloon on surface terrain. For earthbasedballoon systems, crews can often wait days or weeks for low enough winds beforeinflating.3.6.3 Entry Vehicle Mass BudgetWe assume a Delta 7325 launch vehicle that is capable of delivering 614 kg payload to the orbitof Mars (C 3 = 10 km 2 /s 2 ) 14 . The mass of the entry vehicle containing the DARE platform isassumed to be 340 kg, similar to the MABS analysis 2 .Table 3-2. Entry vehicle mass budgetComponentMain parachute 25Heat shield 62Back shell 67Inflation hardware 16Balloon flight system 140Contingency (9%) 30Total 340Mass, kgWe are not assuming any advances in spacecraft or entry technologies. We focus our study onenabling technologies for balloon platform systems, and descent and inflation systems. The entryvehicle mass budget is summarized in Table 3-2.When compared to the previously developed Mars balloon mission concepts, the increase in theballoon flight system mass (140 kg vs. 80 kg in MABS) seen in our design is achieved throughthe use of the advanced lightweight high strength material for the balloon envelope and mainchute; and by the use of a cryogenic system for storage of the buoyant gas (hydrogen).14 http://www1.jsc.nasa.gov/bu2/ELV_US.html53

3.6.4 DARE PlatformFigure 3-24. Cartoon (left) and correct (right) scale DARE Flight SystemFigure 3-24 shows the deployed DARE balloon flight system. The correct scale rendering (on theright) shows the gondola and BGS dimensions appropriately scaled to the diameter of theballoon. The tether length is not to scale.54

Figure 3-25 Gondola of the DARE platform.Figure 3-26. Bottom-View of the DARE gondolaFigure 3-25 shows the conceptual design of the gondola of the DARE platform. The gondolalooks like a thick disk with its top surface covered by the solar panel. Science and engineeringpayloads are housed completely inside the gondola. Having a closed gondola will help tomaintain thermal control of the payload. Imaging payload is positioned on the BGS boom, so55

there is no need for opening in the gondola housing for imaging windows. This drawing does notshow antennas or other payload elements that would be protruding through the walls of thegondola. This view of the gondola shows three deployable probes stowed in their containers atthe bottom and edge of the gondolaFigure 3-26 shows the bottom view of the gondola. The bottom of the gondola has an openingfor the BGS tether. This drawing shows all 5 deployable surface network stations carried on theconceptual DARE platform3.6.5 DARE Platform Mass and Power BudgetsTable 3-3 summarizes the power budget of the DARE platform:Table 3-3. Conceptual DARE ESN mission mass budgetBGS & Envelope -------------------------------> 69BGS 10Winch 2Tether 1Wing 5Battery 0.5Solar Array 1.5Buoyant Gas 9Envelope 50Gondola -------------------------------> 30Structure 15Comms. 5Power 5Controller 2Battery 2Solar Array 1Thermal 5Insulation 3Heater 2Science Payload -------------------------------> 35Dropsondes (5) 25Cameras 9Optics 6Attitude 2Actuators 1Magnetometer 1TOTAL -------------------------------> 134TOTAL AllowableMass -------------------------------> 14056

The power budget for the ESN mission is the sum of the power budgets of the subsystemshoused at the gondola of the DARE platform and of the subsystems of the BGS wing (cameras,attitude sensors, actuators, batteries, etc.). The power budgets for the gondola and BGSsubsystems are shown in Table 3-4 and Table 3-5. It is inefficient to transmit power over thelength of the tether connecting the BGS to the gondola, therefore, both the BGS and the gondolamust have power generating capabilities.Table 3-4 shows the estimated power and energy budget for the gondola. Not included in thisbudget is the power required to reel-up the BGS, as may be required during operation. The BGSreel-up power requirements are estimated in Section 5.4.5.4 (see Figure 5-40).Table 3-4. Gondola power and energy budgetGONDOLA POWER AND ENERGY BUDGETEnergy Power Energy Power Energy(mWhr) (mW) (mWhr) (mW) (mWhr)Day 8 Night 16 TotalCOMM 2200 275 4400 275 6600SCIENCE 8000 1000 16000 1000 24000Magnetometer 8000 1000 16000 1000G & N 3000 375 1120 70 4120Inertial 2700 338 720 45Sun 100 13 0 0Star 100 13 200 13Alt 100 13 200 13COMP 8625 1078 7865 492 16490CPU 4630 579 1810 113SRAM 570 71 160 10A/D 525 66 825 52SSR 2100 263 3470 217Buffer 800 100 1600 100Power 800 100 1600 100 2400Cond 800 100 1600 100Total 14625 1828 14985 937 2961057

Table 3-5 shows the power and energy budget for the BGS. Actuation of the BGS is not shown,but it is thought that any actuators will require minimal power (of the order of mW).Table 3-5. BGS power and energy budgetBGS POWER AND ENERGY BUDGETEnergy Power Energy Power Energy(mWhr) (mW) (mWhr) (mW) (mWhr)Day 8 Night 16 TotalCOMM 800 100 1600 100 2400SCIENCE 152000 19000 115200 7200 267200NAC Camera 48000 6000 1600 100WAC Camera 48000 6000 1600 100IR TES 56000 7000 112000 7000COMP 4320 540 4800 300 9120CPU 2400 300 1600 100SRAM 400 50 160 10A/D 320 40 640 40SSR 800 100 1600 100Buffer 400 50 800 50Power 800 100 1600 100 2400Cond 800 100 1600 100Total 157920 19740 123200 7700 281120Reel-up energy requirements vary from 0 to 350 Whr for the conceptual mission. Averaged overthe whole mission the BGS reel-up energy requirement is about 80 Whr.3.6.6 Balloon Guidance SystemThe Balloon Guidance System or BGS is stowed below the gondola of the DARE platform. Forthe Near term Mars mission we envision an inflatable wing that enables deploying a relativelylarge and light aerodynamic surface. The canister that houses the deflated BGS wing is attachedto the boom. The 10 km long tether that attaches the BGS wing to the balloon’s gondola duringoperation is rolled up on the winch. Upon reaching the floating altitude of the platform, the boomof the BGS is released and starts to descend on the tether attached to the winch. The rate ofdescent is controlled to reduce deployment shock on the boom, the tether and the winch (seeSection 5.3.9.4). At an altitude of about 100 m below the balloon the BGS power generation andstorage subsystems, communication and science suites are checked. The BGS wing is inflatedand lowered to the operational altitude. At that point control and observational sequences cancommence. The onboard computer at the gondola issues a control command to the BGScomputer via a radio link. The BGS computer commands the rudder motor to change the rudderangle, which changes the angle of attack of the BGS wing. The BGS wing starts to generatelifting forces. The BGS computer gathers telemetry from attitude sensors and wind sensors, andestimates the forces being generated by the wing. The BGS computer adjusts the BGS wingangle of attack via rudder so that the generated forces are of the magnitude and in the direction58

prescribed by the main computer at the gondola. As the winds and other environmentalparameters vary, the BGS computer autonomously controls the BGS wing.The BGS computer receives data from attitude sensors and from the gondola computer to assessthe attitude of the BGS boom that houses the cameras. The main computer at the gondolaestimates the location of the DARE platform via star and sun sensors and receives surfaceimagery from the cameras at the BGS. The imagery is used to identify features on the Martiansurface and to pinpoint the platform location by matching these features on the digital terrainmap stored in computer memory.In the vicinity of a surface target, the main computer activates the cameras in the BGS boom,which point at the surface target via actuators using the attitude knowledge estimated fromattitude sensors at the gondola and BGS. The cameras perform context and high-resolutionvisible imagery, and surface thermal spectrum observations. To improve the probability tocapture the target, observations may be performed by scanning a path on the ground. Thecameras are continuously pointed at the target by the actuators as the platform moves by andabove the target.3.6.7 Surface StationsDARE platforms would carry 5 deployable surface stations that would form the nodes of theSurface Network upon deployment at the pre-selected sites.The payload of the deployable surface stations would accomplish the same scientific objectivesas were proposed for the Netlander mission, namely provide first systematic in situ observationsof the Martian lower atmosphere with a network of stations and to study the internal structure ofMars via detection of the seismological signals. The surface stations that would be carried byDARE would be lighter than those envisioned for Netlander due to using new power generatingand storing technologies, and new instrument technologies (such as penetrating seismometers). Itwould be possible to deliver stations at more polarward locations that were originally consideredin Netlander (there, the highest latitudes were limited by power input from the sun).The penetrating seismometer technology looks very promising for reducing the overall mass ofthe surface package and avoiding using parachute or other decelerating devices. Instead the soilof the surface would act to decelerate the surface package. The probe dropped from the DAREaltitude of 10 km would impact the surface with terminal velocity of about 200 m/s, roughly theimpact velocity of the Deep Space 2 (DS2) probes designed by NASA JPL. Penetratingseismometers were considered before, for the Mars 96 ESA balloon mission. Hence, there isconsiderable technological heritage for this technology.Conceptual view of the deployable probe is shown on Figure 3-27. The design is based on theDS 2 Mars entry probes. The probes designed for DARE have similar geometric relationships tothe DS2 probes. The mass of the stations is higher than the mass of the DS 2 probes (2.3 kg), butlower than the mass of the NetLander surface probe (25 kg). The added mass (compared to DS2)accounts for the larger science payload and improved power generation and storage capabilities.The reduction of mass compared to Netlander probes will come from advances in power59

generation and storage capabilities, and removal of the context camera and ground penetratingradar from the payload.Figure 3-27. Surface station/penetrator for ESN mission scenarioThe station consists of the aft body and the penetrator (the protruding lower part on Figure 3-27).The probe would carry a seismometer, a meteorological package (P, T, winds, water vapor), asoil analysis experiment; communication package and power generation and storage package.After dropping a surface station the DARE platform would become lighter and rise in theMartian atmosphere. The balloon’s envelope would become more stressed, as the balloon wouldnow be at an altitude with a lower atmospheric pressure. Venting inflation gas would reduce theinternal balloon pressure and lower the envelope’s stress to acceptable levels (see Figure 5-23).3.6.8 Atmospheric TrajectoryIn this section we present the simulated DARE balloon trajectory in support of the Emplacementof Surface Network mission scenario. We chose the tetrahedron configuration shown on Figure3-21 to demonstrate ESN mission scenario. While this may not be the configuration that wouldbe preferred for the actual mission, it does demonstrate the capabilities and benefits of the DAREarchitecture with a realistic problem.In simulating the trajectory we assumed no specific knowledge of the global atmospheric winds.We did assume a general knowledge of the global winds – which comes from the Global60

Circulation Models (GCM), for example. General knowledge means that we know the generaldirection of the flow for a particular season and dust loading of the atmosphere – let’s say, weknow that there are easterly zonal jets in the mid-latitudes of the both hemispheres at aroundsolstice, - but we do not know the detail of the flow, it’s exact extent, strength, etc. We assumethat we can measure relative winds at the balloon and at the BGS wing, and that we can deducethe local winds in this way.We assumed that during a day the BGS control decisions are made by the onboard computer, andonce every day or two there is an opportunity to download commands from Earth via thetelecommunication orbiter in the Mars orbit.The simulated DARE trajectory is shown on Figure 3-28. The trajectory is plotted over a filledcontour plot of surface topography. The contours are plotted with 1 km interval. The color scaleof the height contours is shown on the legend in the figure. The simulated trajectory is shown bythe black curve, sites of the surface stations are shown by red triangles and a square. The highestlocations on the map are the Tharsis rise and Olympus Mons (latitudes from –30° to 20°,longitudes from –140°E to -90°E). The lowest location on the map is the Hellas Basin (latitudesfrom –30° to -50°, longitudes from 50°E to 100°E).The simulation starts on July 13, 2013 at the latitude of 50º and longitude of 55º. The datecorresponds to the arrival date for the lowest energy Earth-Mars Type I trajectory of the 2013opportunity. The season is almost equinox (L s =356º). The geographic location for the start of thetrajectory is consistent with potential entry location for the 2013 opportunity.The configuration of the DARE platform for this simulation is described in Table 3-6.61

Table 3-6. Configuration of the DARE platform for ESN scenarioParameterValueSeason, L s 356ºBalloon shapeprolatespheroid 1:2Balloon altitude, km 8 - 10BGS tether length, km 8Wing area, m 2 6Wing mass, kg 13Wing length, m 7Wing lift coefficient, C L 1Wing drag coefficient, C DW 0.02Balloon radius, m 17.3Balloon drag coefficient, C DB * 0.2Tether density, kg/m 3 100Tether diameter, mm 0.3* The balloon drag coefficient shown here is not the actual drag coefficient of the prolatespheroid 1:2 (C d =0.022, see Section 4.4.4.1), but a parameter used with the BGS model in thetrajectory simulations. With this drag coefficient the drag force on a sphere of equivalent volumeis equal to the drag force on a prolate spheroid 1:2 (see Section 4.4.4.1).62

Figure 3-28. Simulated DARE trajectory for Emplacement of Surface Network mission scenario.63

The starting point of the trajectory is marked by the number 1. The first mission goal of theDARE platform after the start of the mission is to deliver a surface station to location 2, indicatedby a red triangle on Figure 3-28. This takes about 2 days at this season. After dropping a 5 kgsurface station at the site marked by the number 2 on Figure 3-28, the DARE platform rises byabout 400 m. About 160 g of the buoyant gas is vented to reduce the daily maximumsuperpressure and the associated stress on the envelope. The altitude of the balloon during themission, the maximum height of the BGS wing and the topography are shown on Figure 3-29.Figure 3-29. DARE balloon altitude, the lowest altitude of the BGS wing and surface topographyas a function of timeAfter delivering the first surface station, the DARE platform is directed to fly south, to deliverthe second station to the site at latitude of 10º and longitude of 120º (marked by the number 5 onFigure 3-28). The immediate target is the northern slope of the Alba Patera (number 3 on Figure3-28), to avoid flying over the high topography of Alba Patera. The DARE platform is nowflying over higher terrain typical of the southern hemisphere and the lowest altitude of the BGSwing is raised by 1 km (see Figure 3-29). Note that optimizing the BGS performance mayrequire reeling up the BGS wing for a portion of a day (as described in Section 4.4.5.1). Thesechanges of the BGS altitude are not shown on Figure 3-29.After flying through the wide pass between Olympus Mons and Alba Patera (marked by 4), thelowest altitude of the BGS wing is lowered again and the location for the second station (markedby the number 5) is targeted. The targeting strategy is to push the DARE platform south until it64

eaches the latitude of the target and then maintain the latitude of the target as the zonal windstake the platform in the eastward direction to the target.The second surface station is dropped on day 21 of the mission (5 on Figure 3-28 and Figure3-29). The lightened DARE platform rises by about 400 m and buoyant gas is vented to lowerthe stress on the envelope to the nominal level.At this point, the mission managers may decide to push further south, cross the equator and reachthe sight of the 3 rd station drop at latitude of –15º and longitude of 10º. However, in a couple ofdays it would become apparent that the southward velocity of the DARE platform is notsufficient to avoid impacting Tharsis uplift (a complex of ancient volcanoes occupying latitudesfrom 15º to –30º and longitudes from –130º to -90º). The reason for that is a strong northwardflow in the equatorial regions. The decision would be made to turn the balloon north and trycrossing the equator at a later date and from a more favorable location. At the point marked 6 onFigure 3-28 the DARE platform starts targeting the pass between the Olympus Mons and Tharsis(marked by 7 on Figure 3-28). Alternatively, the platform may be directed to fly north ofOlympus Mons. In both cases the goal is to avoid high topography while moving eastward in theprevailing flow.The DARE platform flies over the highlands in the northern flanks of Tharsis (where the BGSmaximum tether length needs to be shortened to 4 km to avoid high topography, see Figure 3-30(see topography scale for this figure on Figure 3-9) and Figure 3-29) avoiding the hightopography of Tharsis, lowers the BGS wing by 8 km to take advantage of the lower surfacetopography, and is directed south – towards the sight of the drop of the 3 rd surface station.Figure 3-30. High-resolution version of the portion of topography map shown in Figure 3-28,showing trajectory of the DARE platform during passage between Olympus Mons and AscraeusMons (the northernmost volcano in the Tharsis complex)65

The segments of the trajectory shown on Figure 3-30 take the DARE platform over severalchannel and valley systems originating on the Tharsis uplift. High-resolution observations of theorigins of the valley networks are one of the proposed science objectives of the DAREarchitecture (see Section 3.3.3.3). Hence this simulation also illustrates potential trajectories forthese types of observations.The 3 rd station is dropped on August 31, 2013 or day 48 of the mission (9 on Figure 3-28, Figure3-29). Station 3 is dropped near the pre-selected location, missing the targeted site by about 5º.This is done to avoid returning to the location of the 3 rd station after reaching the southern midlatitudes.Alternatively, the site of the 3 rd station can be targeted more accurately by directing theDARE platform north after delivering the surface station at the fourth site at latitude of –45º andlongitude -135º. However, the error of 5º in targeting a surface station should not have anysignificant effect on the operation of the network, so it may be advantageous to drop the stationclose to the pre-selected site when the opportunity presents itself, rather than to wait for a bettertargeted opportunity.After delivering the 3 rd surface station the DARE platform again rises by about 400 m and ventssome buoyant gas to reduce the stress on the envelope from the internal superpressure (9). Next,the final station drop site at latitude of –45º and longitude -135º is targeted. On the way there theDARE platform passes over high topography on the southern flanks of Tharsis, where the BGSwing has to reel-up completely, to avoid impacting the high topography (10 on Figure 3-28,Figure 3-29). The DARE platform drifts in prevailing easterly winds for a couple of days untilthe BGS wing can be lowered and the platform can be controlled again.The 4 th station is delivered on target on September 14, 2013, Earth day 62 of the primarymission. With the primary mission of delivering four surface stations to the nodes of tetrahedronsurface station the platform can continue with a secondary mission – performing observations ofthe magnetic anomalies, surface topography and morphology, and the state of the atmosphere inthe Southern hemisphere. It may be quite possible to transfer the DARE platform back toNorthern hemisphere for an observational campaign there (see Figure 4-26).Figure 3-31. Simulated 60-day trajectory of a free-floating balloon at the same season as the ESNsimulation.66

During the spring in the Southern hemisphere (about 320 Earth days from the end of the primarymission) the DARE platform can move into the Southern Polar Regions to observe processesshaping the polar cap.For comparison, Figure 3-31 shows a simulated 60-day trajectory for the season and startinglocation of the ESN mission scenario. Only the Northern hemisphere of Mars from equator to 80ºlatitude is shown. As can be seen, without control the balloon remains in the mid-latitudes of theNorthern hemisphere within a 10º latitudinal corridor about its initial latitude of 50º.The trajectory shown in Figure 3-28 is one example of a possible trajectory of the ESN missionscenario. Other ESN scenarios and trajectories can be devised. For example, the DARE platformmay spend more time in the Northern hemisphere at the beginning of the mission, takingadvantage of the summer season to observe the Northern Polar Regions. The crossing of theequator and the deployment of the equatorial surface stations may be undertaken in southern fall,when the meridional flow over the equator should be more favorable for crossing from North toSouth (at least according to the GCM simulations).Figure 3-32. DARE platform 30-day trajectory starting on April 22, 2014 (L s =120º) in theNorthern hemisphereFor example, Figure 3-32 shows a simulated 30-day trajectory of a DARE platform travelingfrom the northern to the southern hemisphere during northern summer (L s =120º). As can be seen,the trajectory differs quite a bit from the trajectory in Figure 3-28 simulated for northern spring67

(L s =356-26º). DARE balloon parameters for the simulated trajectory in Figure 3-32 are shown inTable 3-7.Table 3-7. Balloons parameters for the trajectory in Figure 3-32ParameterValueSeason, L s 120ºBalloon shapesphereBalloon altitude, km 10BGS tether length, km 6Wing area, m 2 3Wing mass, kg 11Wing length, m 5Wing lift coefficient, C L 1Wing drag coefficient, C DW 0.02Balloon radius, m 17.2Balloon drag coefficient, C DB 0.5Tether density, kg/m 3 100Tether diameter, mm 0.33.7 Performance Benefits of the DARE ArchitectureDARE architecture offers a number of benefits for planetary exploration over the moretraditional approaches (landers, rovers, orbiters, probes). Some of the observations enabled bythe DARE architecture are unique to the architecture and cannot be performed in any other way.For other observation the DARE architecture offers a less expensive or more capable alternative.DARE architecture has the potential to save billions of dollars to the Mars and Space ExplorationPrograms and accomplish programmatic goals more effectively. Some of the unique and morecapable observations and applications of the DARE architecture are described below.3.7.1 Unique ObservationsA unique capability enabled by DARE platforms is the pinpoint delivery of surface probes topre-selected surface sites. Current capability for probe targeting is limited to landing error ellipse68

with the dimensions of 20 by 120 km. Reducing this error down to the scale of kilometerscurrently is an active area of research. Pinpoint landing directly from space requires addition ofactive propulsion capabilities to the probes, which makes them heavy and complex, and not verydifferent from expensive landers. In addition, the choice of landing sites is limited for aspacecraft approaching a planet, due to the geometry of approach. Hence, deploying severalsurface probes to widely separated surface sites can be a very expensive undertaking. A singleDARE platform can carry several probes for delivery and fly from one selected surface site to theother delivering probes with pinpoint accuracy, as demonstrated in this study of Emplacement ofSurface Stations conceptual mission scenario.Another unique capability of the DARE platforms is the ability to quickly respond to adiscovery. If an orbiter discovers a surface site that has a high probability of having harbored lifein the past or even shows signs of biologic activity today, – a DARE platform that is present atthe planet can be immediately dispatched to that site. The DARE platform would provide highresolutionimagery and spectral observations of the site and deploy surface mini-labs to samplechemical and mineralogical composition at the site. Scientists would not have to wait for 10years to request funding for a new mission to the discovered sites, for a mission proposal to getselected, for the hardware to be build, launched, travel through space and landed or orbitedaround Mars. In this way DARE architecture accelerates the pace of discovery and savessubstantial space exploration resources.Imaging walls of narrow canyons and cliffs – sites where we might see Martian geologicalhistory preserved in rock layers or seepage of subsurface water – are hard to image from an orbitdue to challenging geometry of observations. Orbital cameras can be pointed just slightly offnadir and thus provide greatly distorted imagery of cliffs and canyon walls. On the contrary,imaging cameras of the DARE platforms can be as close as 1 km to the surface and image thecliffs and canyon walls almost at horizon-looking angles. Flying along and above VallesMarineris may actually enable having the cameras below the edges of the cliffs and image thewalls of the largest canyon of the Solar System with an unprecedented clarity.DARE platforms can assist Mars Exploration in a very unique way by serving as a planetarytransport for delivery of samples from remote sites on Mars to the location of the vehiclereturning to Earth (see Section 3.5.3).3.7.2 Alternatives to Traditional ApproachesBeing close to the surface, DARE platforms enable high-resolution imaging (3 cm/pixel) of thesurface features with relatively small (10 cm) and light cameras (5 kg). Achieving the same levelof resolution from orbit would require enormous telescopes with apertures of 100’s of metersweighing 100’s of tons. Landers and rovers can provide an even higher level of resolution, but ata very limited number of sites that support rover mobility or suitable for landing. In contrast,DARE platforms can fly over and image any kind of terrain, the only constraint being hightopography (10 km above reference areoid). Gliders and airplanes can provide imagingcapabilities similar to that of DARE platforms, but they only survive in the Martian atmospherefor at most one to several hours, which limits the number of targets they can image. DAREplatforms will float for years in the atmosphere of Mars ands will have access to a virtuallyunlimited number of imaging targets. Imaging cameras can also be carried by powered low-69

flying airships, however massive propulsion systems require large airships, which require a lot ofpower to overcome the aerodynamic drag forces. As a result, powered airships would be verylarge and massive, and require an excessive amount of power for limited propulsion (at the samelevel of control afforded by the DARE BGS).Measurements of the Martian crustal magnetic field are hard to do from an orbiter and airplanescarrying magnetic field observing instruments would only last for an hour in the Martianatmosphere, providing just a glimpse at the global picture. A balloon platform is an optimalplatform for these observations at Mars. DARE platforms can be controlled to overfly areas ofcrustal magnetic anomalies enabling characterization of the spatial frequency of the magneticanomaly pattern, the depth, thickness and lateral extent of the sources. The relations of themagnetic features to topography would be established and the subsurface water reservoirs will bemapped.DARE platforms enable long-term in situ observation of the atmosphere on global scale –something that cannot be done with any other platform. Space probes can provide in situatmospheric profiles, but only over one location and for a single moment in time. Sending morethan a handful of such probes would be prohibitively expensive. At the same time we are findingout that the Martian atmosphere varies over extended spatial and temporal scales andunderstanding it requires observations at several sites and for more than a year (for example,atmospheric water vapor, CO 2 and dust cycles). Airplanes are powered airships that have a verylimited lifetime and are very massive and power demanding, respectively. Remote sensinginstruments can provide some information about the state of the atmosphere and its constituents,but in many cases lack resolution and accuracy for definitive measurements. For example, it wassuggested before to use infrared solar occultation spectrometer onboard a spacecraft to measureconcentrations of trace gases in the Martian atmosphere. In this approach the instrument lookssideways through Mars' atmosphere toward the setting or rising Sun for an extremely sensitivereading of what chemicals are in the thin air that the sunlight passes through before hitting theinstrument. However, this technique enables these observations only over specific locales (in theatmosphere above the terminator) and the measurements are necessarily averaged over a thickslice of the atmosphere. A DARE platform carrying a suite of Tunable Laser Spectrometers (seeSection 3.3.3.1) onboard would measure the gases right at the location of the platform - and moreaccurately, would not be limited by the time of day or location to make the measurements.DARE can measure atmospheric gases in polar day conditions, when the above-described remotesensing technique could not be applied due to the absence of sunsets and sunrises.In situ atmospheric wind measurements are very important for verifying atmospheric and climatemodels and for assistance with landing and aerobraking procedures. Wind measurements can bemade from orbit (wind Lidar), but weather forecasters on Earth still rely on daily in situmeasurements by radiosondes. The DARE platform can provide first global in situ measurementsof atmospheric winds and wind profiles (employing lightweight dropsondes).DARE platforms can validate observations from orbit by flying over the observed region andproviding imagery or spectral observations at a different resolution and from a different vantagepoint. In this way orbital observations can be tied to “ground truth” and interpreted moreaccurately. Validation of orbital observations with a surface probe or an airplane would require a70

prohibitive number of such platforms given the extent of the existing orbital observationsdatabases.71

4 Balloon Guidance Research4.1 IntroductionIn this section we discuss the Balloon Guidance System (BGS) research that was performed inthe Phase II of the DARE architecture development effort. We discuss the physics and operationof the BGS, the development of the advanced mathematical model of the BGS, efforts towardsoptimization of the BGS performance and the development of the atmospheric trajectorysimulation software.4.2 Balloon Guidance System Operation and PhysicsThe Balloon Guidance System (BGS) of the DARE platforms consists of a wing hanging on itsside below the balloon on a very long (several km) tether. Due to the difference in windsbetween the altitudes of the balloon and the wing, the wing experiences relative winds, thatcreate a sideways lifting force, that can be used “to drag” the balloon across the winds. Becausethe density of the atmosphere is higher at the wing altitude, the wing can be much smaller thanthe balloon. Figure 4-1 shows the single- and dual-wing BGS systems currently being developedby GAC. The dual-wing BGS increases the amount of control force (compared to a single-wingBGS) that can be exerted on a balloon in a strong winds environment.Figure 4-1. Single- and dual-wing BGS.A scaled version of the single-wing BGS was successfully tested in 2001. The development ofthe full scale BGS mechanical wing assembly, winch testbed and tether for Earth applicationswere completed in 2002. Global Aerospace Corporation is currently in discussions with NASAon the completion of the development and flight-testing of the prototype BGS.72

Figure 4-2. Simplified BGS Vector Diagram.Figure 4-2 shows a vector diagram illustrating the dominant forces during operation of the BGS.The view is looking down from above the system and is not to scale. The BGS (represented as anairfoil section) is at a much lower altitude than the balloon (represented by the circle). Forillustration purposes, the BGS wing is shown much larger than it would be in proportion to theballoon. The upper portion of the figure shows expanded views of some small details.Many simplifying assumptions are inherent in this illustration. The more complex models thatwere developed in Phase II do not share these limitations. For example, the tether is shown hereas a straight line. In reality, due to the variation in relative wind and atmospheric density alongits length, the drag forces on the tether will cause it to have a gentle curvature. The morecomplete models include this effect by breaking the tether down into shorter segments overwhich conditions are treated as being constant. Here, the tether drag force is assumed to act atthe BGS. This is not a bad assumption since the drag on the lower 10% or so of the tetherdominates the rest of the tether because the atmospheric density is greatest and the relative windis also greatest here. The wing is assumed to be exactly vertical so the lift and drag forces act in ahorizontal plane. In actual operation, the wing will hang with some tilt to the side andbackwards. The more complete models resolve the forces in three dimensions.73

The meanings of the various vectors is as follows:Table 4-1. Notation for Figure 4-2.V 10V 3V BTrue Wind Velocity at Balloon altitude (~10 km on Mars)True Wind Velocity at BGS altitude (~3 km on Mars)True Velocity of Balloon = True Velocity of all parts of the systemV DRIFT Drift Velocity of the balloon due to action of the BGS = V B - V 10V REL Relative Wind Velocity at the BGS = V B – V 3V DF Vector Difference between Winds at Balloon and at BGS = V 3 - V 10(used in an even simpler analysis, but not used here)V CT Cross-Track Velocity Component of V DRIFT (perpendicular to V 10 )VBT Back-Track Velocity Component of V DRIFT (parallel to V 10 )F L Lift Force on BGS (acts horizontally and is perpendicular to V REL )F D Drag Force on BGS (acts horizontally and is parallel to V REL )F RResultant force on BGS = F L + F D ~ Drag force on balloonF DRIFT(Not illustrated) Drag Force on balloon due to V DRIFT ~ F RThe system is assumed to be in equilibrium, so the vector sum of the forces must equal zero. Inthe vertical direction, the buoyancy force exactly equals the weight of the system. These verticalforces are not shown. The aerodynamic drag force on the balloon is equal to and opposite to theresultant aerodynamic force on the BGS (including the tether drag force). With thissimplification, one can calculate the drift velocity of the balloon (inset box in figure) in whichA B is the projected area of the balloon (as viewed from the side), C D is the drag coefficient of theballoon (for flow from the side), and ρ is the atmospheric density at the altitude of the balloon.The angle of attack, α, of the BGS wing is controlled by adjusting the incidence angle of therudder (not shown), and is usually arranged to produce close to the maximum lift coefficient forthe wing as this typically produces maximum useful control effect for the system. A smallamount of iteration is required to determine the drift velocity vector that balances the forces,making the system self-consistent. The diagram is drawn to illustrate a self-consistent solution.4.2.1 BGS performance in low Reynolds number regimeIn recognition that operation in the Martian environment would result in fairly low Reynoldsnumbers (on the order of a few thousand), we investigated existing literature on low Reynoldsnumber wings. Although there have been some investigations in this regime, there are far fewersources of data or modeling results than for the higher Reynolds numbers typical of normal74

aircraft flight. It is clear that for these low Reynolds numbers, drag coefficients are significantlyhigher and maximum lift coefficients are a little lower than for higher Reynolds numberoperation, as studied by Sunada et al. (2002) 15 . The reduced L/D is quite a challenge for systems,which must generate their lift aerodynamically while providing power to overcome the drag 16 .However, for the proposed BGS system, the weight is supported by buoyancy. The “lift” fromthe wing is directed close to horizontal and predominantly across the flight path of the balloon.The drag acts mostly to slow down the balloon, and is relatively unimportant to the operation ofthe proposed BGS system. In fact the drag on the tether is typically much greater than the dragon the BGS wing. The forces generated by the BGS are very low (of order 1 N). However, thedrag on even a large balloon moving at only 1 m/s in the Martian atmosphere is also very small.Figure 4-3. Examples of the low Re wings profiles and corresponding lift and drag curves 15 .The airfoil sections appropriate to operation at Reynolds numbers near 1000 are quite a bitdifferent from those for higher Reynolds numbers. Typically, the airfoil sections are muchthinner (about 2%) resembling thin plates rather than streamlined airfoils. Some corrugatedshapes have been tested, emulating wings of insects, and have been found to give similar results15 Sunada, S., Yasuda, T., Yasuda, K., and Kawachi, K., "Comparison of Wing Characteristics at an Ultralow ReynoldsNumber," Journal of Aircraft. Vol. 39, No. 2, March-April 2002, pp. 331-33816 Kunz, P.J., Strawn, R.C., "Analysis and Design of Rotors at Ultra-Low Reynolds Numbers," 40th AIAA Aerospace SciencesMeeting & Exhibit, 14-17 January 2002, Reno, NV, AIAA 2002-009975

to flat plates under steady conditions, and may give better performance for flappingapplications 15 (see Figure 4-3). There is some indication based on Navier-Stokes solvers thatmaximum lift coefficient actually improves again as the Reynolds number decreases below a fewthousand, although the drag continues to rise, being dominated by viscous effects 17 . Datameasured on small wings towed in a water tank show lift coefficients in excess of 1.0 (as high as1.4) for a strongly cambered flat plate at a Reynolds number of a few thousand.In our discussion with JPL engineers concerns were expressed regarding the damping of thesmall force generated by the BGS by the tether. However, it is difficult to imagine a mechanismto swallow up steady state force acting on the tether despite its length of several kilometers. Ofcourse, this is one reason to test such a system. However, the system is operating in essentiallysteady state, with dynamic effects being due to changes in the wind patterns. The only"damping" is the drag on the tether. A tether with a length of 5000 m and a diameter of 0.2 mmhas a projected area of 1 m 2 , significantly smaller than the balloon, but of the same magnitude asthe wing. But the atmospheric density decreases appreciably with altitude, so the effect area, if itwere all considered acting at the wing, would be on the order of 20% of this area. The overallcombined L/D is likely to be no better than 1. But again, the drag serves mostly to slow down theballoon a little. The lift force still moves the system across the prevailing winds. It is a smallforce, but it acts persistently, and over time. The integral produces a result that is useful and thedriving energy comes freely from the gradient of the winds within the atmosphere.As much as our predictions give us confidence that the system will perform as predicted, we arefirm believers in experimental confirmation. We also advocate taking an incremental approach,starting with relatively simple tests, and progressing to more complex tests only after the simpleones have worked successfully. For instance, before building a full-scale wing for earthapplication, we developed a quarter scale model with moments of inertia and mass dynamicallyscaled and tested the aerodynamic performance of just the wing-boom-rudder assemblysuspended from a tethered blimp (which simply acted as a sky hook) in natural winds at ElMirage dry lake. In addition to confirming that the system behaves as expected, we also exploredpost-stall behavior, investigated dynamic behavior (spiral mode has no meaning here, but it wasvery difficult to excite Phugoid and short period modes, and there were no Dutch Roll-likeyawing or rolling instabilities coupling into the tether), operated the system with both stable andunstable center of mass locations, verified successful canard and conventional configurations,measured the lift curve including a maximum lift coefficient of 1.2 at a Reynolds number ofabout 50,000 corresponding to the full-scale system (on Earth).4.2.2 Dynamic scaling of the Mars BGSScaling laws developed during a previously funded activity were modified and applied to theMars environment in order to develop sizing of a scale model for testing on the Earth that wouldsimulate a full-scale BGS operating in Martian conditions. A full-scale wing with area of 2 m 2and aspect ratio of 10 was selected as a nominal design. The wingspan is 4.47 m and the chord0.447 m. When scaled to Earth test conditions, a wing span of 37 cm and wing chord of 3.7 cm17 Kroo, I., Prinz, F., "The Mesicopter: A Meso-Scale Flight Vehicle," NIAC Phase II Technical Proposal, http://wwwrpl.stanford.edu/files/report/Proposal_Phase2.pdf.76

was determined. This size wing would have the same Reynolds number and Froude number asthe full-scale system. This is a size that could reasonably be used in a wind tunnel, although thesmall size would require accurate machining.Table 4-2. Dynamic Scaling Parameters4.3 Balloon Guidance System (BGS) Numerical ModelThe model of the BGS consists of a number of non-linear equations that are solvedsimultaneously. The solution balances horizontal and vertical forces acting on the systemballoon-tether-BGS wing. The vertical forces arising from the operation of the BGS are balancedby the gravity forces acting on the BGS wing and tether. The altitude of the balloon is keptconstant while the system of equations is being solved. The buoyancy forces acting on theballoon were excluded from the model for simplicity. Inclusion of these forces would make theballoon change altitude in response to BGS operation. These altitude changes are small for theBGS wings considered for Mars, so the error is minimal. If the vertical wind structure does notchange significantly with altitude (wind shear is constant with altitude), the exclusion of the77

uoyancy forces does not introduce an error for any altitude change. On Venus, the altitudechanges may be significant and must be taken into account in balloon design (see Section 4.3.2).The model calculates aerodynamic forces for the components of the DARE platform. Drag andlift force vectors are calculated for the BGS wing using the formulas appropriate for the wingdimensions and area. The lift coefficient of the wing is an input parameter. The tether is modeledas a number of linked elements. Aerodynamic drag on each link of the tether is calculated as foran infinite cylinder using atmospheric parameters at the altitude of the tether link to calculateReynolds number and relative winds. The spatial orientation of the tether links is calculatedusing the tension forces acting on the edges of a tether link, drag and gravity. The only forceacting on the balloon is the aerodynamic drag.The system of equations is solved iteratively, adjusting the spatial orientation of the BGS wingand the tether links at each time step until a self-consistent solution that balanced all the forces isfound. In the model of the Dual-Wing BGS, the roll angle of the wing is varied and the attitudeof the DBGS is adjusted until a solution that balances all the forces and provides maximumcontrol force in the direction perpendicular to the direction of the horizontal winds at balloonaltitude is found.When these models are incorporated into the trajectory simulation code with Mars GRAM 2001or Venus GRAM 2005 environments they may have difficulties converging. This is because bothMars GRAM and Venus GRAM have an option to add a random component to the wind fields intheir databases. As the BGS numerical model goes through iterations, it receives updates fromthe environment models on the environment parameters at the locations of the BGS modelelements. Since these updates vary randomly, this may “confuse” the solver of the BGS modeland prevent convergence. Ways to resolve this problem include turning off the randomcomponent in Mars GRAM and Venus GRAM, or “caching” the environmental parameters fromMars or Venus GRAM the first time BGS model requests these parameters, and using just theseparameters in subsequent iterations during a single time step of the trajectory simulation model.Interestingly, Mars GRAM and Venus GRAM behave like “quantum objects” in these situations– they change their state unpredictably each time we measure their state!4.3.1 Mars BGS Performance Numerical ModelingHaving developed and tested the numerical model of the BGS we have undertaken a study of theBGS performance. We have looked at BGS performance in different wind regimes characterizedby strong or weak vertical wind gradients, varied tether length and balloon size. The modelsolves for the self-consistent solution that maximizes the cross-track velocity of the balloon (thevelocity in the direction perpendicular to the direction of the prevailing winds).An example of a model solution is illustrated in the figures below. Table 4-3 summarizes themodel input parameters. This model run takes place at about equinox (Ls=6°). Thecorresponding date is August 13, 2013 – within a month of the 2013 Mars arrival window ofopportunity for the Network Emplacement Mission. The winds at the altitude of the balloon (10km) are shown on Figure 4-4.78

Table 4-3. Input parameters for the BGS model run at Ls=6ºInput ParameterValueLatitude, deg 40Longitude, deg 0Season, Ls 6Balloon altitude, km 10BGS tether length, km 8Wing area, m 2 3Wing mass, kg 11Wing length, m 5Wing lift coefficient, C L 1Wing drag coefficient, C DW 0.02Balloon radius, m 17.2Balloon drag coefficient, C DB 0.4Tether density, kg/m 3 100Tether diameter, mm 0.3Number of tether segments to model 10Ratio of tether segments lengths 0.979

Figure 4-4. Global wind field at the 10 km altitude on Ls=6°.Two strong westerly jets in both hemispheres characterize this season at Mars. In both jets thewinds rapidly increase with altitude. The winds over the equatorial region are much weaker andthe vertical gradients are less well pronounced.The location of the model run is on the southern edge of the zonal jet in the northern hemisphere.The action of the BGS in this case changes the velocity of the DARE platform by 2.5 m/s. TheBGS velocity vector is close to the direction to southwest. The solution for the tetherconfiguration and the relative winds are shown in Figure 4-5. Note the different scales in the Uand V directions. The scales in U and V directions are also different from the vertical scales toshow the tether shape. The deflection of the tether from the straight down configuration in the Udirection is due to the drag of the tether in the strong relative U wind. The deflection from thestraight down configuration in the V direction is primarily due to the generated sideways liftingforce of the BGS.80

Figure 4-5. Tether shape in the a) U direction (EW), b) V direction (NS), c) and the relativewinds.The DARE platform continues to be embedded into the zonal flow (ground speed U=45.7 m/s)with a slight southward drift more than 50% of which are due to the action of the BGS (groundspeed V=-2.7 m/s, BGS U=-1.84 m/s, BGS V=-1.83 m/s).Table 4-4 below gives the output parameters of the BGS model:Table 4-4. Output parameters of the BGS model for Ls=6°.Parameter nameValuesU (EW) V (NS) ZGravity, m/s 2 3.7278Density at balloon altitude, kg/m 3 0.005839364BGS mass, kg 11Tether length, m 8000Number of tether segments 10Tether link length ratio 0.9Tether density, kg/m 3 100Tether diameter, m3.00E-04Wing length, m 5Wing area, m 2 381

BGS wing lift coefficient 1BGS wing drag coefficient 0.02U velocity of BGS, m/s -1.839286422V velocity of BGS, m/s -1.827793497W velocity of BGS, m/s 0True velocity vector 45.72775971 -2.736354226Drag coefficient of tether segment 1 4.138914291Drag coefficient of tether segment 2 4.150929524Drag coefficient of tether segment 3 4.172215475Drag coefficient of tether segment 4 4.212239657Drag coefficient of tether segment 5 4.287804024Drag coefficient of tether segment 6 4.435670864Drag coefficient of tether segment 7 4.745451509Drag coefficient of tether segment 8 5.473364753Drag coefficient of tether segment 9 7.678246338Drag coefficient of tether segment 10 18.62117366BGS wing orientation vector, m -0.025929179 -0.674272236 -4.954259241Tether segment 1 Reynolds number 5.686313686Tether segment 2 Reynolds number 5.653819843Tether segment 3 Reynolds number 5.597008685Tether segment 4 Reynolds number 5.492727969Tether segment 5 Reynolds number 5.304459234Tether segment 6 Reynolds number 4.965725085Tether segment 7 Reynolds number 4.362583293Tether segment 8 Reynolds number 3.342323887Tether segment 9 Reynolds number 1.832341205Tether segment 10 Reynolds number 0.427511452Force at BGS wing, N -0.210710327 -5.479391607Resultants force at tether segment 1, N -0.237531106 -5.478879377Resultants force at tether segment 2, N -0.288120449 -5.477842328Resultants force at tether segment 3, N -0.383044969 -5.475721966Resultants force at tether segment 4, N -0.559313183 -5.471341322Resultants force at tether segment 5, N -0.880475867 -5.462271516Resultants force at tether segment 6, N -1.444564828 -5.443861564Resultants force at tether segment 7, N -2.363599828 -5.408870286Resultants force at tether segment 8, N -3.634330051 -5.352032036Resultants force at tether segment 9, N -4.801944095 -5.273493654Resultants force at tether segment 10, N -5.12131593 -5.063280208Force at balloon, N 5.17677373 5.144426257In Table 4-4 U and V velocity of the BGS are the control velocities generated by the BGS.These velocities arise due to the lifting force of the BGS wing and due to the drag forces of theBSG wing and the tether. Vertical control velocity of the BGS W is always 0 – it is assumed thatthe vertical forces are balanced. The true velocity vector is the vector of the velocity of theDARE platform relative to the ground. This velocity is the resultant of the flow velocities and ofthe control velocities generated by the BGS. Drag coefficients and Re for each tether segment are82

given for the middle-point of the tether segment. The tether segments are counted from the BGSwing. Tether segments of different lengths can be seen in Figure 4-5. The BGS wing orientationvector gives the components of the vector running through the center of the wing along itslength. As can be seen, the BGS wing is almost vertical in this case (the vertical projection of thewing is 4.95 m long, while the wing is 5 m long – the plane of the BGS wing is tilted by about 8ºto the vertical). The resultant forces are given for the mid-point of each tether segment. The forceat the BGS wing is the force due to BGS wing lift and drag. The force at the balloon is the dragforce due to the balloon being dragged through the air by the forces generated by the BGS wing.The resultant forces are calculated by consequently adding the forces at each tether segment tothe force at the BGS wing. As can be seen in this example the tether drag increases thecomponent of the force in the U direction with increasing altitude, while in the V direction thetether drag is in the opposite direction than the BGS wing lifting force in the V direction. Theresultant forces are equal in magnitude and opposite in direction to the balloon force. In thisexample there is actually a small residual force (0.05 N, 0.08 N) due to limitations of thenumerical model trying to minimize the residual force. This residual force is small (about 1%)compared to the forces acting on the BGS wing, balloon and the tether - ~5 N – and so can beconsidered non-existent.The Reynolds number Re for the balloon for the configuration described above is 5.2*10 4 ,corresponding to the drag coefficient of 0.4 for a sphere. This value of Re is about a factor of 2 to4 smaller than the value of Re for which the drag coefficient for a sphere reduces to 0.2 (the socalled“drag crisis”). Operating the balloon in the regime beyond the drag crisis (i.e. withRe>2*10 5 ) would improve the BGS performance by about 40%. Changing the balloon’s shape tothat having a smaller C d than a sphere (see analysis in Section 4.4.4) or reducing the diameter ofthe tether can improve BGS performance. Reducing the tether diameter is not always feasiblesince the strength of the tether depends on its diameter, and certain strength is required to supportthe weight of the tether itself and of the BGS. Changing the shape of the balloon increases themass of the balloon for constant balloon volume. However, even for a small increase in theballoon’s mass (10%-20%) the reduction in C d - and the corresponding improvement in the BGSperformance - is quite significant (see Section 4.4.4). Hence, in our analysis of the ESN missionscenario we assume a non-spherical DARE balloon.The example above describes the conditions that are very well suited for the operation of theBGS – winds at the altitudes of the balloon and the BGS are almost parallel, large relative windat the altitude of the wing. However, these atmospheric conditions are, in general, limited to thezonal jets in the mid-latitudes of both hemispheres. Let’s consider the BGS performance in morechallenging conditions.Table 4-5 gives the BGS model input parameters for the analysis of the BGS performance in theequatorial regions characterized by weak zonal and meridional winds that are not parallel atdifferent altitudes.83

Table 4-5. Input parameters for the BGS model run Ls=89ºInput ParameterValueLatitude, deg 0Longitude, deg 180Season, Ls 89Balloon altitude, km 8BGS tether length, km 6Wing area, m 2 3Wing mass, kg 11Wing length, m 5Wing lift coefficient, C L 1Wing drag coefficient, C DW 0.02Balloon radius, m 14.6Balloon drag coefficient, C DB 0.5Tether density, kg/m 3 100Tether diameter, mm 0.3Number of tether segments to model 10Ratio of tether segments lengths 0.9The wind field plot is shown on Figure 4-6.84

Figure 4-6. Global wind field at the 8 km altitude on Ls=89°.The model solution is illustrated on Figure 4-7. In this case the relative wind at the wing altitudeis much smaller than before – just about 1 m/s in the zonal direction. Interestingly, the total BGSvelocity is of the same order of magnitude as the relative wind at the altitude of the wing – 0.9m/s. Most of this BGS control velocity comes from the tether drag. The cross-track velocitygenerated by the BGS (measured relative to the vector of the relative wind at the wing altitude) isjust 0.02 m/s. The tether deflection shown on Figure 4-7 is all due to the tether drag (notedifferent spatial scales on all three plots). This analysis illustrates the worst case for the BGSperformance – in low relative wind at the wing altitude the cross-track component of the BGSvelocity is negligibly small. However, the conditions described in this analysis do not persistlong at Mars. Simulations for a different longitude at the equator at the same time (different localtime, not shown) show an increase in relative wind and a corresponding increase in the crosstrackcomponent of the BGS velocity (1 m/s for relative wind of 9 m/s). Hence, our analysisindicates that the BGS will enable balloon steering capabilities for a wide range of atmosphericconditions that can be encountered at Mars.85

Figure 4-7. Tether shape in the a) U direction (EW), b) V direction (NS), c) and the relativewinds.Table 4-6 below gives the output parameters of the BGS model in this case:Table 4-6. Output parameters of the BGS model for Ls=89°.Parameter nameValueU (EW) V (NS) ZGravity, m/s 2 3.7278Density at balloon altitude, kg/m 3 0.007178408BGS mass, kg 11Tether length, m 6000Number of tether segments 10Tether link length ratio 0.9Tether density, kg/m 3 100Tether diameter, m3.00E-04Wing length, m 5Wing area, m 2 3BGS wing lift coefficient 1BGS wing drag coefficient 0.02U velocity of BGS, m/s 0.828411975V velocity of BGS, m/s -0.141421621W velocity of BGS, m/s 086

True velocity vector 0.454818858 1.374718711Drag coefficient of tether segment 1 23.64114933Drag coefficient of tether segment 2 23.10855799Drag coefficient of tether segment 3 22.33482908Drag coefficient of tether segment 4 21.2017876Drag coefficient of tether segment 5 19.66200878Drag coefficient of tether segment 6 17.66772981Drag coefficient of tether segment 7 15.30734269Drag coefficient of tether segment 8 12.81216993Drag coefficient of tether segment 9 10.47152736Drag coefficient of tether segment 10 19.27653054BGS wing orientation vector, m 3.78E-04 0.00174433 4.999999681Tether segment 1 Reynolds number 0.293529866Tether segment 2 Reynolds number 0.304200115Tether segment 3 Reynolds number 0.320897437Tether segment 4 Reynolds number 0.348269385Tether segment 5 Reynolds number 0.392249515Tether segment 6 Reynolds number 0.464713536Tether segment 7 Reynolds number 0.584334503Tether segment 8 Reynolds number 0.778942627Tether segment 9 Reynolds number 1.084850697Tether segment 10 Reynolds number 0.404724379Force at BGS wing, N 0.003101352 0.014305527Resultants force at tether segment 1, N 0.003411965 0.014250085Resultants force at tether segment 2, N 0.004034589 0.014154611Resultants force at tether segment 3, N 0.005313732 0.01399229Resultants force at tether segment 4, N 0.008049554 0.013728155Resultants force at tether segment 5, N 0.01421832 0.013283088Resultants force at tether segment 6, N 0.029209775 0.012435342Resultants force at tether segment 7, N 0.069216889 0.009946645Resultants force at tether segment 8, N 0.187641594 -9.41E-04Resultants force at tether segment 9, N 0.57733369 -0.052847604Resultants force at tether segment 10, N 0.80291636 -0.091222272Force at balloon, N -0.836670383 0.1428314484.3.2 Venus BGS Performance Numerical ModelingWe have studied BGS performance at Venus with our numerical model. Without detailed balloondesign effort we estimate balloon sizes of the Venus DARE platforms by assuming a constantfloating mass of 70 kg for the balloon flight system (consistent with the previous Venus balloonanalyses), altitudes of 55 and 60 km and a range of envelope surface densities from 250 to 350g/m 2 (also consistent with previous studies). Altitudes of 55 to 60 km are chosen because of therelatively benign atmospheric conditions at these altitudes at Venus (T=300-260 K, P=0.53-0.24bar, respectively). Venus balloon films for these are heavier than the films for Martian balloons(350 g/m 2 vs 20 g/m 2 ) because Venus balloons have to withstand much larger superpressure87

stress due to higher atmospheric pressure at Venus. In addition, protection from atmosphericsulfuric acid at Venus dictates using heavy Teflon fabric for balloon envelopes.Dependence of the payload mass and the balloon radius on envelope surface density and balloonaltitude is shown in Figure 4-8.Figure 4-8. Dependence of the Venus DARE platform payload mass and balloon radius onaltitude and envelope film density.As can be seen from Figure 4-8, Venus balloons can carry 30 to 40 kg payloads at altitudes of 55to 60 km at Venus. These balloons are relatively small with diameters between 5 and 7 meters.The two Vega balloons deployed in the Venus atmosphere in 1985 floated at altitudes between50 and 55 km. The balloons were 3.4 m in diameter and their envelope surface density was 289g/m 2 .Atmospheric circulation at altitudes of 55 to 60 km at Venus is characterized by strong zonalflow (Figure 4-9), reaching wind speeds of 90 m/s at the altitude of 60 km.88

Figure 4-9. Venus zonal winds profiles from Venera and Pioneer Venus probes.The wind shear (the change of the wind speed with altitude) can reach values of several metersper second per kilometer of altitude at the altitude of 55 to 60 km (see Figure 4-8).Meridional winds are much weaker than the zonal winds and increase in amplitude with altitude(Figure 4-10).89

Figure 4-10. Meridional winds profiles for the Pioneer Venus probes 18We used Venus-GRAM 2005 as the atmospheric environment for the numerical modeling of theBGS performance at Venus. We ran simulations of Single-Wing and Dual-Wing BGS. Theresults of the analysis are given in Table 4-7. There, U bgs and V bgs are the BGS control velocitiesin U and V directions, h is the balloon altitude, R b is the balloon radius, L t is the length of thetether, A w is the area of the BGS wing, L w is the length of the BSG wing, M w is the mass of theBGS wing, D t is the diameter of the BGS tether, ρ t is the density of the BSG tether, Lat is theVenus latitude of the simulation, C d is the balloon drag coefficient, SBGS and DBGS are singleanddual-wing BGS, respectively. Figure 4-11 shows BGS control velocities V from Table 4-7plotted as a function of BGS tether length and the balloon floating altitude.18 P. Gierasch et al., “Circulation of the Venus atmosphere”, in Venus II, The University of Arizona Press, 199790

Table 4-7. Venus BGS modeling summaryrun # Ubgs Vbgs h, km Rb, m Lt, km Aw, m 2 Lw, m Mw, kg Dt, mm ρt, kg/m 3 Lat Cd BGS1 3.37 1.58 55 2.63 7 3 5 10 1 100 30 0.2 SBGS2 4.11 0.85 55 2.63 10 3 5 10 1 100 30 0.2 SBGS3 2.69 1.8 55 2.63 5 3 5 10 1 100 30 0.2 SBGS4 2.1 1.54 55 2.63 4 3 5 10 1 100 30 0.2 SBGS5 3.1 1.78 55 2.63 5 3 5 10 1 100 30 0.2 SBGS6 2.74 1.8 60 3.3 6 3 5 10 1 100 30 0.2 SBGS7 1.77 1.47 60 3.3 7 3 5 10 1 100 30 0.2 SBGS8 3.36 1.6 60 3.3 8 3 5 10 1 100 30 0.2 SBGS9 2.5 0.94 60 3.3 6 1 2.2 5 1 100 30 0.2 SBGS10 1.56 0.76 60 3.3 4 1 2.2 5 1 100 30 0.2 SBGS11 3.22 0.89 60 3.3 8 1 2.2 5 1 100 30 0.2 SBGS12 2.19 1.26 60 3.3 6 1 2.2 5 0.5 100 30 0.2 SBGS13 2.8 1.2 60 3.3 8 1 2.2 5 0.5 100 30 0.2 SBGS14 1.38 0.98 60 3.3 4 1 2.2 5 0.5 100 30 0.2 SBGS15 2.02 1.44 60 3.3 6 3 5 5 0.5 100 30 0.2 SBGS16 3.72 1.87 60 3.3 7 3 5 5 0.5 100 30 0.2 DBGS17 2.8 1.7 60 3.3 8 3 5 10 0.5 100 30 0.2 SBGS18 2.69 2.12 55 2.63 5 3 5 10 0.5 100 30 0.2 SBGS19 3 2.22 55 2.63 6 3 5 10 0.5 100 30 0.2 SBGS21 7.3 4.7 55 2.63 6 3 5 10 0.5 100 30 0.2 DBGS91

Figure 4-11. Dependence of the Venus BGS V velocity on the tether length and balloon altitude.Table 4-7 and Figure 4-11 show that BGS performance is maximized with the use of the dual-BGS wing. This is due to the BGS system generating very strong lifting forces that lift and tiltthe single-wing BGS into higher altitudes, reducing relative winds and the horizontalcomponents of the lift. The dual-wing BGS rolls to turn its lifting force vector downward, whichprevents the wing from flying up. Note, that there is little difference between SBGS and DBGSperformance at Mars, because the lifting forces are much smaller and the BGS wing weightkeeps it from tilting and flying up. The aerodynamic forces generated by the BGS at Venus areseveral orders of magnitude larger than at Mars (see below), which makes DBGS a moreefficient system.Table 4-7 and Figure 4-11 also indicate that BGS performance improves for: lower altitudes (dueto smaller balloon size and drag area), smaller tether diameter (due to reduced tether drag). Thereis also seems to be an optimal tether length at 6 to 7 km for the studied balloon system setup. Theoptimal tether length may arise due to opposing trends of increasing tether drag and increasingBGS relative winds with increasing tether length.Figure 4-12 shows tether configuration for run 21 – the run that produced the maximum BGS Vcontrol velocity.92

Figure 4-12. Tether configuration for run 21 in Table 4-7.Relative velocity at the BGS wing altitude was 9.7 m/s in run 21. Table 4-8 shows the output ofthe BGS model for the run 21. The notation is the same as described for Table 4-4.Table 4-8. Output of the BGS model for Venus BGS run 21.Parameter nameValueU (EW)Gravity, m/s 2 8.6328Density at balloon altitude, kg/m 3 0.9207BGS mass, kg 10Tether length, m 6000Number of tether segments 5Tether link length ratio 0.8Tether density, kg/m 3 100Tether diameter, m5.00E-04Wing length, m 5Wing area, m 2 3BGS wing lift coefficient 1BGS wing drag coefficient 0.02U velocity of BGS, m/s 7.338627439V velocity of BGS, m/s 4.671638734V (NS)93

W velocity of BGS, m/s 0True velocity vector 72.66137256 4.671638734Drag coefficient of tether segment 1 1.1940330428106625Drag coefficient of tether segment 2 1.2044011388019151Drag coefficient of tether segment 3 1.2401195710882627Drag coefficient of tether segment 4 1.2893304056733026Drag coefficient of tether segment 5 1.311306726617151DBGS wing roll angle, rad -0.871300713Wing Re1.5E+05Tether segment 1 Reynolds number 369.9869947Tether segment 2 Reynolds number 342.1960825Tether segment 3 Reynolds number 268.756305Tether segment 4 Reynolds number 203.1930672Tether segment 5 Reynolds number 182.0611337Balloon Re2.8E+06Force at BGS wing, N 93.68056366 160.0836517Resultants force at tether segment 1, N 103.7594425 154.321861Resultants force at tether segment 2, N 118.7988644 144.6960219Resultants force at tether segment 3, N 135.8323944 130.8150977Resultants force at tether segment 4, N 147.8786974 111.5277606Resultants force at tether segment 5, N 127.7273198 81.30733106Force at balloon, N -127.727362 -81.30895004Results of the modeling presented in Table 4-8 indicate that the balloons and the DAREplatforms and wings of the BGS at Venus will be operating at high Reynolds numbers (2.8*10 6and 1.5*10 5 , respectively). For run 21 the roll angle of the DBGS wing is 50º up from thedirection directly down. Hence, the force generated by the DBGS is divided almost equallybetween the downward and sideways components.Our numerical model of the BGS performance does not balance vertical forces. Vertical forcesgenerated by the BGS will push the whole system up or down in the atmosphere, until the newbuoyant equilibrium is reached. For simulations in the Martian atmosphere, where the BGSforces are small, the vertical component of the BGS force has very little effect on the altitude ofthe balloon. At Venus, the forces are large enough to change the altitude of the balloon byseveral kilometers. When the balloon is pulled by the BGS wing deeper into denser atmosphere,at some depth the ambient pressure will overwhelm the internal pressure of the superpressureballoon and the balloon will collapse, causing an unrecoverable descent to the surface. Let’sassume that we can tolerate downward altitude excursions of 2 km with the initial altitude of 55km. The change of ambient pressure over this 2 km is about 34% of the pressure at 55 km (from0.53 to 0.71 bar, respectively). Hence, to tolerate the 2 km downward altitude excursion, theVenus balloon has to be designed for 34% superpressure. Our analysis above assumed a 5%superpressure. Increasing superpressure from 5 to 34% requires an addition of about 2 kg ofbuoyant gas (hydrogen). Atmospheric density at 53 km is 1.15 kg/m 3 . We can estimate the forceneeded to push the DARE balloon downward by 2 km in the following way:94

Fdρ V = M + ,gwhere ρ is atmospheric density, V is balloon volume (76.2 m 3 ), M is the floating mass of theflight system (70 kg), F d is the downward force and g is Venus gravity (8.6 m/s 2 ). The requireddownward force is about 160 N. This is very close to the downward force generated by theDBGS in run 21 above (153 N). Hence, DBGS creating a 7.6 m/s control velocity in the Vdirection will also push the balloon down by 2 km. If the wing generates higher lift forces (due tohigher relative winds, for example), it will push the balloon deeper, where it will losesuperpressure and collapse. Hence, the DBGS wing performance needs to be carefully monitoredduring operation and the downward component of the lifting force limited to safe values.4.4 Balloon Guidance System Performance OptimizationWe have studied Martian atmospheric environment in order to optimize the performance of theBalloon Guidance System. We analyzed seasonal and diurnal density variations in theatmosphere and the atmospheric wind structure. This analysis was done using the Mars-GRAMmodel of Martian atmosphere.4.4.1 MARS GRAM 2001We used Mars Global Reference Atmospheric Model version 2001 19 – Mars-GRAM 2001 - inour analysis of Martian environment. Mars-GRAM is an engineering-oriented model of theatmosphere of Mars. Mars-GRAM is based on input data tables of output from the NASA AmesMars General Circulation Model (MGCM) 20 and the University of Arizona Mars ThermosphericGeneral Circulation Model (MTGCM) 21 . Mars-GRAM outputs a number of parameters requiredin our analysis, such as profiles of the pressure and temperature, winds, radiative fluxes, etc. Allthese data are available for a range of atmospheric dust opacities, which is crucial for correctlysimulating the Martian environment.4.4.2 Analysis of Martian Atmospheric EnvironmentThe variation of the height of the balloon’s floating altitude (constant atmospheric density level,see Section 5.4.1) also affects the performance of the BGS. The performance of the BGS isroughly proportional to the square root of the ratio of densities at the height of the wing of the19 C. G. Justus, D. L. Johnson, “Mars Global Reference Atmospheric Model 2001 Version (Mars-GRAM 2001): Users Guide”,NASA/TM-2001-210961, 200120 Haberle, R. M., Pollack, J. B., Barnes, J. R., et al.: “Mars Atmospheric Dynamics as Simulated by the NASA Ames GeneralCirculation Model 1. The Zonal-Mean Circulation.” Journal of Geophysical Research, Vol. 98, No. E2, pp. 3093-3123,1993.21 Bougher, S. W., Roble, R. G., Ridley, E. C., et al.: “The Mars Thermosphere: 2. General Circulation with Coupled Dynamicsand Composition.” Journal of Geophysical Research, Vol. 95, No. B9, pp. 14,811-14,827, 1990.95

BGS and the height of the balloon. The larger the ratio - the larger is the control force generatedby the BGS.We analyzed variations in the ratio of densities at the BGS wing and balloon altitudes assuming a5 km tether and balloon floating level corresponding to the density level of 0.0058 kg/m 3 (8-10km altitude depending on season and location). We have looked at three seasons - summer in theNorth (Ls=90°), equinox (Ls=180°) and summer in the south (Ls=270°). The ratio varies fromabout 1.4 to about 2 for different locations and seasons. The ratio is the largest and the BGSperformance is the best over the poles during winter seasons (North Pole - data points 600-700,Ls=270° and South Pole - data points 0-100, Ls=90°), because the atmosphere is cold and "moredense". However, these seasons and locations correspond to polar night and thus would require apower source that is not dependent on the sun. Both poles at equinox show elevated values of thedensity ratio (1.9 at South Pole and 1.7 at North Pole), and thus may be the best locations tooperate the BGS. The lowest values of the ratio are observed over the equator at all seasons (1.4to 1.6) and over summer poles (1.4 – 1.5).Since the BGS control velocities are roughly proportional to the square root of the density ratios,they will vary by about 20% with season and geographic (assuming the same relative winds andother parameters).4.4.3 Analysis of Martian WindsThe performance of the BGS is strongly dependent on the structure of the atmospheric winds.Variations of the wind directions with altitude can strongly affect BGS performance. Thesevariations affect the directions of the aerodynamic drag forces acting on the BGS tether and inthis way affect BGS performance. For an example of the BGS performance in winds that blow indifferent directions at different altitudes see Figure 4-7 and accompanying text in Section 4.3.1.To find out how often and to what extent the vertical wind structure will affect the BGSperformance we analyzed the relationship between the directions of the wind vectors in ahorizontal plain at different altitudes in the atmosphere.Figure 4-13 shows an example of such an analysis for the Northern summer season (L s =90°) fortwo altitudes: 5 and 10 km. These altitudes correspond to potential altitudes of the BGS wing andthe balloon, respectively. The plot shows the histogram of the wind differences between the twoaltitudes on a global scale. This plot shows that the winds at 5 and 10 km are mainly alignedparallel to each other – 50% of the wind directions differ by less than 10° degrees. There ishowever a long “tail” in this distribution of wind direction difference angles that indicates thatthe winds can be misaligned by as much as 50°. This analysis indicates that conditions that arefavorable for BGS operation are common at Mars.96

Figure 4-13. Histogram of the wind direction differences between 5 and 10 km altitudes in theMartian atmosphereThe BGS performance depends most strongly on the difference between the wind speeds at thealtitude of the DARE balloon and the BGS wing. The difference between wind speeds definedthe relative wind at the BGS wing altitude, which creates the lifting and drag forces used tocontrol the balloon system. Below is an example of the extensive analysis that we haveundertaken. The example is for the season of Ls=150°. We are looking at a balloon at constantdensity level of 0.005 kg/m 3 (atmospheric heights from ~9.5 to 10.5 km above the reference levelacross Mars), and tether length of 6 km.97

Figure 4-14. Histogram of wind speed differences between 10 and 4 km for Ls=150°.Figure 4-14 shows the histogram of wind differences between the altitudes of the balloon and ofthe wing. Not surprisingly, the distribution is bimodal, corresponding to two wind patternsexisting in the atmosphere at this season: a strong easterly jet in the southern mid-latitudes (winddifferences between 10 and 40 m/s) and more benign winds in the tropical regions and northernmid-latitudes (wind differences between –15 and 15 m/s). From the point of view of BGSoptimization, it would be beneficial to be in the jet in the south, where wind differences wouldvary from 10 to 40 m/s. The corresponding control velocities could be from 2 to 8 m/s.The winds do not change monotonically with altitude, so that a shorter tether or a lower balloonaltitude would yield a better BGS performance.Other environmental parameters, such as the Reynolds numbers, also influence BGSperformance.98

Figure 4-15. Histogram of the wing Reynolds numbers.Figure 4-15 shows the histogram of the Reynolds numbers for the BGS wing. The Reynoldsnumber defines the aerodynamic regime in which the BGS wing operates. As can be seen onMars the wing will operate at a very low Reynolds number (of the order 1000). This requiresspecial considerations for the design of the BGS wing (see Section 4.2.1).4.4.4 BGS Performance with Non-Spherical BalloonsLooking for ways to improve the performance of the Balloon Guidance System (BGS) we turnedto non-spherical balloons. Our analysis with the BGS numerical model indicates that the balloonsat Mars would be operating at Reynolds numbers below 10 5 , thus the balloon drag coefficientwould be equal to about 0.5. Non-spherical – for example, elliptical balloons, - have lower dragcoefficients. The reduction in drag comes at a price of increasing the balloon area and mass forthe same volume. We have also looked at increasing the area of the BGS wing as a way toimprove performance.4.4.4.1 Balloon shapes - prolate spheroidsConsider a prolate spheroid as a balloon (Figure 4-16). We are going to estimate the dragcoefficient of such a balloon when operating a BGS, and also estimate the mass penalty foremploying such a balloon.99

Figure 4-16. 3D views of a prolate spheroida - semi-major axis (l/2)b – semi-minor axis (d/2)V4 π ab 32S=2⎛ arcsin e ⎞b= 2π b⎜b+ a ⎟ , where e = 1−2⎝ e ⎠ aSuch a balloon would be passively oriented with the semi-major axis along the relative windswith fins or similar appendages. The mass of these appendages is not considered in the followinganalysis.For a prolate spheroid with the aspect ratio of 1:2 (l/d=2, a/b=2):V434343323= π R = πab= π 2b, where R – radius of the sphere of equivalent volume.It follows:Rb = ≈ 0. 8Rand a=1.6R,32d=2b=1.6Rl=2a=3.2R⎛ arcsin(0.87) ⎞2S = 2π⋅ 0.8R⎜0.8R+ 1.6R⎟ = 2πR( 0.64R+ 1.55R) ≈ 4.4πR⎝0.87 ⎠- a 10% increaserelative to equivalent sphere surface area.100

Consider DARE Mars balloon:R b =15.8 mρ=0.006 kg/m 3 (density at balloons altitude)η=10 -5 Pa-s (viscosity)V= 1 m/s (relative velocity)Then:2RVRe = ρ = 1.9·10 4ηC d_balloon =0.5For a prolate spheroid of the same volume:VlRel= ρ = 3.2 Re = 6.1·10 4 , C d_wet =0.05 (see Figure 4-17, Re l – Reynolds number calculatedηusing the length of the body l, C d_wet is drag coefficient based on the “wetted area”=S).Figure 4-17. Aerodynamic drag for rotationally-symmetric bodies 2222 S. F. Hoerner, “Fluid-dynamic drag”, 1958.101

The C d of the sphere of equivalent volume (R=15.8 m) can be calculated by equating the dragforce on the prolate spheroid and the sphere:Fd _ sphere1C221ρ V A = Fd_ spheroid= C2=d ⊥d _ wetρV2SHence:2Cd_ wetS Cd_ wet4. 4πRCd= == 4.4·0.05=0.22, i.e. or a factor of 2 reduction in the drag force2A⊥πRrelative to the equivalent volume sphere.Similarly, for a prolate spheroid with the aspect ratio of 1:3 (l/d=3, a/b=3):V434343323= π R = πab= π 3b, where R – radius of the sphere of equivalent volume.It follows:Rb = ≈ 0. 7Rand a=2.1R,33d=2b=1.4Rl=2a=4.2R⎛arcsin(0.94) ⎞2S = 2π⋅ 0.7R⎜0.7R+ 2.1R⎟ = 2πR( 0.49R+ 1.93R) ≈ 4.8πR- a 20% increase⎝0.94 ⎠relative to equivalent sphere surface area.For the same balloon setup as above:VlRel = ρ = 4.2 Re = 8·10 4 , C d_wet =0.02 (see Figure 4-17)η2Cd_ wetS Cd_ wet4. 8πRCd= == 4.8·0.02=0.1, or a factor of 5 reduction in the drag force2A⊥πRrelative to the equivalent volume sphere.In summary, a prolate spheroid with aspect ratio 1:2 generates a drag force that is a factor of 2smaller and has surface area that is 10% larger than that of the equivalent volume sphere. Aprolate spheroid with aspect ratio 1:3 generates a drag force that is a factor of 5 smaller and hassurface area that is 20% larger than that of the equivalent volume sphere.102

In our design we are going to assume a DARE balloon that is a prolate spheroid with 1:2 aspectratio.4.4.4.2 Increased area of the wingWe have also looked at ways to improve BGS performance by increasing the area of the BGSwing.Doubling the wing area to 6 m 2 has a smaller (compared to shape change), but significant effect– ~50% increase in the SBGS cross-track velocity.BGS performance studyBGS-0.4-0.9-1.4BGS V Fall 1BGS V Fall 4BGS V Fall 7BGS V Fall 8-1.9-2.49/13 9/14 9/15 9/16 9/17 9/18Date, dayFigure 4-18. BGS V winds for different DARE platformsFigure 4-18 shows a simulation of daily variation of the BGS V velocities simulated for 4different DARE platforms for Ls=21º (9/13/2013) - Early Northern hemisphere spring, for astarting location with the coordinates of 30º latitude and 260º longitude. Mars GRAM winds areused in this simulation. The four cases differ in the size and shape of the balloon, and in the sizeof the BGS wing:Fall 1: Balloon: R b =17.2 m, A w =3 m 2 , L t =6 km, C d =0.5Fall 4: same as Fall 1, smaller balloon – R b =15.8 m, L t =7 km, C d =0.5Fall 7: same as Fall 4, prolate spheroid, equivalent C d =0.22Fall 8: same as Fall 7, larger wing - A w =6 m 2 103

Here A w – BGS wing area, L t – length of the tether, C d – balloon drag coefficient. The names ofthe simulations (Fall 1, Fall 4, …) refer to the Earth date of the start of the simulation (fall of2013). The balloon’s altitude is about 10 km above the Mars areoid (reference level).The control objective of the simulation was to push the balloon southward by maximizing theBGS V. The magnitude of the BGS V (northward) velocity in this simulation changes from 0 toabout –2 m/s. The magnitude of the V velocity changes with the changing local andcorresponding relative winds. The maximum value of the BGS V velocity is achieved when therelative wind at the altitude of the BGS reaches its maximum. The BGS V is equal to 0 m/s whenthe relative winds are unfavorable for southward movement and the BGS is reel-up so as not todrag the whole system in the wrong direction (see more on BGS control techniques in Section4.4.5).The simulation shown in Figure 4-18 shows that employing a larger wing, and a smaller balloonshaped as a prolate spheroid can improve the BGS performance by a factor of about 2.4.4.5 Guidance TechniquesWhile simulating DARE trajectories at Mars we developed several techniques that optimize theperformance of the BGS. These are described in detail below.4.4.5.1 BGS reeling up and downPerformance of the BGS is strongly dependent on the strength of the relative wind at the BGSwing altitude. The general assumption while operating the BSG is that the winds increase withaltitude and, hence, maximizing the altitude difference between the balloon and the BGS wing(essentially, increasing the length of the BGS tether) will maximize the performance of the BGS.This assumption is correct most of the time at Mars, especially in the mid-latitude regions thatexhibit strong seasonal zonal atmospheric jets. However, in the tropical regions and over theequator, the vertical wind structure might be more complex and the optimal performance of theBGS is not correlated with the length of the tether.For example, Figure 4-19 shows the vertical profile of total winds for a range of latitudes in theNorthern hemisphere for 3/22/2014. As can be seen from that figure, the maximum relative windfor the BGS wing will be achieved with the BGS wing at altitudes of 5 to 6 km (for the DAREballoon at 10 km altitude).Thus one of the possible techniques to optimize the BGS performance would be to reel the BGSwing up or down to find the best altitude. The tether may carry small wind sensors that wouldprovide information on relative winds above the BGS wing altitude. If the maximum of relativewinds is found at the current BGS wing altitude, the wing will be lowered down until theperformance of the BGS starts to worsen or until reaching the preset limiting altitude abovetopography. If the maximum relative wind is detected above the BGS wing, the wing will bereeled-up to that altitude.104

Figure 4-19. Vertical wind profile in the Northern hemisphere on 3/22/2014.Another situation, when reeling up the BGS wing improves the BGS performance, occurs whenthe relative winds at the BGS wing altitude have certain strength and blow in a certain direction.Figure 4-20 shows the vector diagram of the daily change of the relative winds at the BGS wingaltitude for 4 days in April and May of 2014 for a DARE balloon traveling in the tropical andequatorial regions of Mars.Figure 4-20. Vector diagram of the daily relative winds at the BGS wing altitude105

The circular curves in Figure 4-20 show the end points of the wind vectors of the relative windsat the BGS wing altitude. The starting point is the origin. One vector is shown in red. Each curverepresents end points of the wind vectors for 25 consecutive hours during a Martian day (sol).For example, for 4/22 there are moments when there is no relative wind (the dark blue curvepasses through the origin), there are moments when the relative winds are towards north (Ucomponent equal to 0 m/s, V component equal to 5 m/s), and moments when the winds aretowards east (U component equal to 13 m/s, V component equal to 0 m/s). Remarkably, for atime period of almost 20 days the winds behave in a very similar way each day.If the control objective of the DARE balloon is to travel south, then, when the relative winddirection is close to being northward, the BGS is not generating any useful control (since theuseful control is directed perpendicular to the relative wind vector). On the contrary, the drag ofthe tether and of the wing will be directed northward and hence prevent the DARE platform fromreaching its target in the south. In a situation like this reeling the BGS wing completely up mayprove quite beneficial by reducing the drag of the system in the “wrong” direction.Reeling up of the BGS wing should take place over a relatively short period of time, sinceperiods of the “bad” winds persist for hours during a day. Assuming that a 13 kg BGS wingneeds to be reeled-up in 1 hour over length of 10 km we obtain the power requirement of 130 Wfor BGS reel-up. This requirement should be added to the power budget and the DARE platformpower system sized with the BGS wing reel-up in mind.4.4.5.2 BGS as a drag deviceWhen the relative winds blow in the “right” direction (i.e. southward when the control objectiveis to move south) it may be beneficial to use the BGS wing as a simple drag device. The BGSwing can be turned almost perpendicular to the flow and drag the whole DARE platform in theright direction. We included this BGS capability in our trajectory simulation code, assuming thatthe BGS generates drag force as a flat plate with the area equal to the area of the wing. Thetrajectory shown in the next Section is simulated using the techniques described in this Section.4.5 Atmospheric Trajectories SimulationsIn this section we discuss simulations of guided balloon’s trajectories in the atmosphere of Mars.We discuss the numerical tool used to simulate the trajectories and give examples of thesimulated trajectories. Examples of atmospheric trajectories can also be found in Section 3.6.8.4.5.1 Trajectory Simulation ToolTo simulate the atmospheric trajectories of the DARE platforms we use the software that wassuccessfully used in Phase I of the study. In Phase II we modified the software to use the outputof the Mars-GRAM 2001 model (see Section 4.4.1) as the new simulation environment.To simulate an atmospheric trajectory we need information about the winds at a particularlocation. To simulate the balloon flight, we follow the constant atmospheric density levels in oursimulations. This means that the altitude of the balloon varies as it floats above Mars (seeSection 5.4.1). When the DARE platform drops a dropsonde, the altitude of the balloon changes106

correspondingly. To keep track of the height of the balloon the data on the pressure andtemperature atmospheric profiles are needed. This information is provided by the Mars-GRAMmodel for a wide range of atmospheric conditions that can be encountered on Mars. In addition,during a simulation we keep track of the height of the BGS system, which should be above thelocal topography. If the height of the BGS system is below the height of the local topographythat means that the BGS system has impacted the topography. The BGS will probably bedestroyed by such an impact and the DARE platform will loose its steering capabilities withoutthe BGS. Even though a hypothetical mission can continue without the BGS (with reducedobjectives), here we consider impacting topography as an end of mission. The Mars-GRAMmodel also provides data on the local topography.We have incorporated the new advanced models of the Single Wing BGS into the trajectorysimulation model. In Phase I, due to complexities of incorporating the full BGS model into thetrajectory simulation code we used an approximation, in which the control velocity due to theaction of the BGS was constant. This simplification is justified if the vertical structure of theatmospheric winds does not change significantly for the duration of the balloon mission. Thisapproximation was sufficient for the Phase I work and produced some important results thatjustify the use of the BGS to control balloons at Mars. In Phase II of the effort we fullyincorporated the BGS model into the trajectory simulation code. The control of the DAREplatform is determined at each time step using the environmental conditions provided by theMars-GRAM model at the locations of the balloon, BGS and along the tether. The threedimensionalpositions and attitudes of the BGS wing and segments of the tether are dynamicallyadjusted as the DARE platform is being propagated through the model atmosphere. The availabletrajectory control capabilities are continuously recalculated using the BGS model and trajectorycontrol commands are issued by the control algorithm accordingly.One of the features of the Mars-GRAM model is the addition of the random component to thewind and density fields to mimic the natural variability of the atmosphere on short time scales.Our trajectory simulations are performed with this feature of the Mars-GRAM turned off since itinterferes with the convergence of the BGS model. BGS code can be modified to be morecompatible with the random winds in the Mars-GRAM in the future.4.5.2 Atmospheric TrajectoriesAs part of the effort to define the Emplacement of Surface Networks conceptual missionscenario, we have simulated and analyzed a number of atmospheric balloon trajectories. Figure4-21 shows a 30-day simulated trajectory that illustrates possible guided balloons paths in theMartian atmosphere. The trajectory is plotted over the contour map of the Martian topography.The height color scale is shown in the legend in Figure 3-28.The season of the simulation is northern summer (Ls=89°, February, 13, 2014). This date isabout 200 sols (Martian days) later than the starting date of the trajectory shown on Figure 3-28.During the summer months a cross equatorial flow develops bringing warm air from the summerhemisphere to the winter hemisphere. The DARE platform floats at 9.6 to 10 km altitude for theduration of the mission.DARE platform parameters for this simulation were:107

Table 4-9. DARE BGS parameters for the trajectory on Figure 4-21.Input ParameterValueSeason, Ls 89Balloon altitude, km 9.6-10BGS tether length, km 6Wing area, m 2 3Wing mass, kg 11Wing length, m 5Wing lift coefficient, C L 1Wing drag coefficient, C DW 0.02Balloon radius, m 17.2Balloon drag coefficient, C DB 0.5Tether density, kg/m 3 100Tether diameter, mm 0.3The simulation starts at a latitude of 10° and longitude of 180°E, with the balloon at the altitudeof about 10 km above the reference level. At the start of the simulation the balloon is in thewesterly flow of the northern hemisphere. The action of the BGS makes it drifts southward. Atthe longitude of about 50°E after 6 days of flight the balloon crosses the equator. It continues todrift slowly in the weak equatorial winds until being picked up by the strong easterly flow ataround latitude of -20°. The wavy appearance of the simulated trajectory is due to diurnalatmospheric tides. The balloon covers 40° of latitude over the 30-day simulation.The platform in this simulation has weaker control capabilities than the one in the simulation onFigure 3-28. The difference is that the balloon in this simulation is spherical and no reel-up andother guidance techniques are being employed.Another sample simulation for the seasons of Ls=89° is shown Figure 4-22. It differs from theprevious one only by the starting position.108

Figure 4-21. Simulated trajectory, Ls=89º.Figure 4-22. Another simulated trajectory for Ls=89°.109

The following simulated trajectory is an example of trajectory that can be employed for mappingthe crustal magnetic anomalies in the Southern Hemisphere (Section 3.5.1). The DARE platformconfiguration is the same as in ESN mission scenario. The tether length is 5 to 6 km. The seasonis southern spring (Ls=200°). Figure 4-23 shows a 30-day trajectory plotted over the Martiantopography contour map. The trajectory starts over the Hellas basin (-45°S, 60°E). The platformis directed to float south up to a latitude of 80° and then fly back north. The platform is notguided to target any specific sites. The simulation ends near the lower right-hand corner of theplot.Figure 4-23. Simulated DARE trajectory for observations of crustal magnetic anomalies,Ls=200°.Figure 4-24 compares the simulated trajectory of a free-floating balloon (red) and of a DAREplatform (black) plotted over the map of crustal magnetic anomalies in the Southern hemisphere.As can be seen, the free-floating balloon is confined to a very narrow latitudinal corridor over a30-day mission, while the guided DARE platform is capable of exploring a much larger area. Bymoving towards the South Pole the DARE platform can circle Mars faster and thus completes 5passes over the field of crustal anomalies while the free-floating balloon completes only twopasses.110

Figure 4-24. Simulated trajectory of a free-floating balloon (red) and a DARE platform (black)plotted over the map of crustal magnetic anomalies.Figure 4-25 shows the altitude of the DARE platform during the simulation shown on Figure4-23. The sudden dip in altitude between 9/29/14 and 4/3/14 is due to the platform moving overthe edge of the receding southern CO 2 polar cap. The platform crosses the edge of the cap at thelatitude of about –69°S on 9/29/14, flows south up to a latitude of -80°S, turns back north andcrosses the edge of the cap again on 4/3/14, but this time at the latitude of –67°S.Figure 4-25. Altitude of the DARE platform for the simulation on Figure 4-23.111

Figure 4-26 shows an example of a trajectory that takes the DARE platform from the Southern toNorthern hemisphere. The DARE platform configuration is as for ESN mission scenariosimulation. The season of the start of the simulation is early northern spring Ls~23º, September16, 2013. The trajectory is simulated for 24.5 Earth days. The time of the start of the simulationis near the end of the ESN simulation shown on Figure 3-28.Figure 4-26. Simulated trajectory of a DARE platform crossing equator from southern tonorthern hemisphere in early northern spring, Ls=23º.The simulation starts in the atmosphere at the site marked 1 on Figure 4-26. The DARE platformis at the altitude of about 9200 m. At the start of the simulation the BGS wing is reeled upcompletely, so the platform behaves as a free-floating balloon. The BGS wing is reeled-up to flyover high topography south of Tharsis (-30ºS, -100ºE). After flying over the high topography inthe strong zonal flow prevailing in the Southern hemisphere at this season, the DARE platformlowers the BGS wing by 5 km to steer northward (2). As the platform floats further north overlower topography, it lowers the BGS wing even more – to 8 km length, and gains more control.It crosses the equator and than gets into weak equatorial flow that slowly takes it westward. Asthe platform drifts north pushed by the BGS, it enters stronger easterly flow at the latitude ofabout 20ºN. At this point the DARE platform is at the latitude where it clears most of the hightopography of the Northern hemisphere and is in sufficiently strong flow that enables enoughcontrol to steer the platform to practically anywhere in the Northern hemisphere.112

5 DARE Platform Design Description5.1 IntroductionThis section discusses the DARE platform design development effort of Phase II of the study.We discuss enabling and enhancing technologies relevant to the design and present the roadmapfor technology development. We conclude with the description of design trade studies thathelped to shape the design and the roadmap.5.2 Enabling Technologies Development RoadmapIn this section we describe the roadmap for the development of the DARE architecture. Theroadmap lists Enhancing and Enabling technologies relevant to the DARE architecture. Enablingtechnologies are viewed as those technologies that must be present for any level of DAREmission – such as advanced balloon materials, for example. Enhancing technologies will allowfor more intensive DARE missions. Enhancing technologies include energy storage, platformstructure, etc. It is assumed that the commercial sector will advance enhancing technologies,resulting in increased performance of the DARE platform. Reduction in the mass of the DAREplatform is the primary driver for the development of enabling and enhancing technologies.Various DARE system elements referred to in the following sections are defined in Figure 5-1.Figure 5-1. DARE system definitions.113

5.2.1 Technology HorizonsTo facilitate the development of the roadmap we assign technologies to three technologyhorizons: Current, Near term and Far term.The Current technology horizon is taken to be technologies with a Technology Readiness Level(TRL) of 8-9, requiring 0-3 years to get to TRL 9 (NASA TRL definitions are given in Figure5-2). The Current technology horizon includes the following assumptions: 28-29% efficient solarcells and energy storage devices with 150 Whr/kg densities. The envelope will be spherical andmade of polyethylene with a Mylar coating to protect against the elements.Figure 5-2. NASA Technology Readiness Levels (TRL) definitions.The Near-term technology horizon includes technologies with a TRL of 3-6, requiring 3-10 yearsof development to get to TRL 9. In the design work we will use technologies from the Near termtechnology horizon. It is envisioned that near-term technologies will allow for modest increasesin photovoltaic efficiency and significant increases in energy storage density as chemical fuelcells become readily available. Structural materials are projected to make only modest increasesin strength-to-weight ratios and toughness. The balloon envelope is seen as a gas barriercontained in a low weight scrim all enclosed in a non-structural protective layer. The Neartechnology horizon is the focus of the design work with the other horizons providing guidancefor future work and acting as inputs for the BGS model.The Far-term technology horizon includes technologies with a TRL of 1-3, requiring more than10 years of development to get to TRL 9. While these technologies are difficult at best toforecast they are desirable for future missions. It is envisioned that carbon nanotubes (CNT) willallow for extremely light structural components such as the balloon scrim and the gondola114

structure. Power is produced continuously by quantum dots thus eliminating the need for energystorage other than a minimal amount of primary cell capacity for system initiation.5.2.2 DARE Mission RequirementsThe most general DARE mission requirements that drive the requirements for the enabling andthe enhancing technologies are flight duration and the mass of the balloon flight system. Theserequirements are summarized in Table 5-1.Table 5-1. DARE missions requirementsTechnology Horizon Current (3-5 years) Near (10-15 years) Far (25-40 years)Flight duration, days 100 700 5-10 yearsBalloon flight system mass, kg 80 140 199The requirements for mission duration come from the requirements for global planetary coverageand for observations of several seasons at Mars. Relatively slow speed of the DARE platform(~10 m/s) and the specifics of the platform’s unique path guidance approach may require monthsof flight time to reach particular spots on the planet. Variability of the Martian climate requirespresence of this unique platform for several Martian years.The balloon flight system mass requirements are driven by the desire to maximize the mass ofthe useful payload delivered to Mars and to drive the requirements for the enabling andenhancing technologies beyond the limits of what is currently possible. The next sectiondiscusses the assumptions that went into defining the balloon flight system mass requirements.5.2.3 Balloon Flight System MassAs a starting point for the DARE design analysis and technology projections we define the total“useful” mass that can be injected into the Martian atmosphere from Earth. We call this mass a“balloon flight system mass”. The balloon flight system mass is the balloon mass plus mass ofthe buoyant gas plus the mass of any suspended weight (see Figure 5-1):m = m + m + mfbgswhere suspended mass includes the BGS mass, the science payload mass and the support mass(power, telecommunications, structure, etc.):ms= mBGS+ msci+ m supFrom the analysis of the MABS system 2 the mass of the balloon flight system with the mainparachute and inflation hardware for a Delta 7325 launch vehicle with 616 kg capability wasabout 211 kg. The balloon flight system mass was 83 kg. We assume no advances in the cruisestage and entry vehicle, and only project advances in balloon system technologies (i.e. reductionin the inflation tank weight, reduction in the parachute weight, balloon envelope weight, etc.). Inthis way, for the constant mass of the entry vehicle “content” of 211 kg, we project the balloonflight system masses as shown in Table 5-2:115

Table 5-2. DARE balloon flight system masses for the three technology horizonsTechnology Horizon Current (3-5 years) Near (10-15 years) Far (25-40 years)Balloon flight systemmass, kg80 140 199Balloon film arealdensity, kg/m 2 0.024 0.012 0.006Balloon mass, kg 51(6.5 km floatingaltitude, D=27 m)50(8-10 km floatingaltitude, D=35 m)28(8-10 km floatingaltitude, D=39 m)Buoyant gas mass, kg 12 (He) 8.3 (H 2 ) 12 (H 2 )Payload, kg 16 81 159The masses of the balloons and of the buoyant gas in Table 5-2 are calculated based on theanalysis of Section 5.4.1 and assumptions on balloon envelope density.Defining balloon flight system masses for the Near and Far technology horizons allows us tostudy balloon sizes and system mass budgets for a range of possible balloon altitudes in theatmosphere of Mars.5.2.4 RoadmapThe roadmap describing the evolution of the DARE architecture technologies over the threetechnology horizons is displayed below in Table 5-3. Each of these technologies and requiredadvancement projections are discussed in more detail in the following section.New materials for development of stronger and lighter fabrics will play a crucial role in thedevelopment of the DARE architecture. These composite materials will enable lightweightballoon envelopes, reducing the mass of the balloons and enabling larger balloon payloads, andlighter parachutes, reducing the total mass of the Entry, Descent and Inflation system.Simultaneous with the development of the new materials, Entry, Descent and Inflationtechnologies will mature, enabling safe deployment and inflation of balloons in the planetaryatmospheres. The hardware will become lighter due to employment of cryogenic systems for gasstorage and stronger, and lighter materials in gas storage tank construction.Balloon Guidance Technology is essential to the success of the DARE architecture. Onceprototype testing of the first generation Single-Wing BGS system is successful, the developmentof advanced BGS designs (inflatable, Dual-Wing) can proceed. The BGS system will belightweight and will be able to achieve high lift forces. Guidance techniques will be developed tooptimize the performance of the BGS enabling navigation to pre-selected locations on the planet,targeted observations and hazard avoidance. Autonomous navigation and control techniques116

together with trajectory prediction capabilities will be developed to support operation of theDARE platforms.Lightweight and small science sensors with novel capabilities will be developed and enable newtypes of observations of the planetary atmospheric and surface phenomena. New, expendableexploration vehicles, including surface probes, atmospheric sondes, droppable rovers, will beused to deliver these science sensors to their targets from DARE platforms.Solar power generation and energy storage technologies will develop to offer higher power andenergy densities. Advancement in these technologies will enable a reduction in mass of theDARE platforms, and also extends the spatial range and lifetime of the platforms.117

Table 5-3. DARE architecture conceptual technology roadmap.Technology Area Current (about 2010) Near Term (2015-2020) Far Term (2030-2045)Entry, Descent, Inflation (EDI) Chute 500 m 2 – 50 kgEntry Vehicle – 340 kgChute 500 m 2 – 25 kgEntry Vehicle – 340 kgChute 500 m 2 – 0 kgEntry Vehicle - 340 kgGas StorageInflation Hardware MassGas MassStorage RatioGaseous He74 kg11.7 kg (He, MABS 2 )16%Liquid H 215.4 kg8.3 kg (H 2 )54%Liquid H 212 kg12 kg (H 2 )100%Make-Up Gas System not needed not needed with envelopeElectrolysis systemmanufacturing advancesAdvanced Envelope MaterialsEnvelope Density3.5-µm Mylar55 denier Kevlar6-µm-PE0.020 kg/m 2 1-µm Mylar38 denier Kevlar3-µm-PE0.012 kg/m 2 1-µm Mylar1 denier CNT1-µm-PE0.006 kg/m 2Film thermo-optical (TO)propertiesAluminized top, white paintbottomAluminized top, white paint bottom Variable-optical-propertiesfilmPrimary StructureSpecific strength-to-weight ratioComposite GraphiteSSTWR = 16Short CNT CompositeSSTWR = 300Long CNT CompositeSSTWR = 612SSTWR (7075Al = 1)Power Generation Solar Array 25% GaAs2.4 kg/m 2 33% Multi Junction2.0 kg/m 2 42% Multi Junction/Quantum Dots (QD)1.6 kg/m 2Supplemental PowerPrimary Cell Batteries350 - W/kg0.35 - kWhr/kgTritiated ASi30 - W/kg1000 - kWhr/kgQuantum Dots (QD) ArraysTBD - W/kgTBD - kWhr/kgEnergy StorageEnergy DensityLi-ion Polymers150 - W-h/kg150 - W/kgFuel Cell400 - W-h/kg120 - W/kgLi-Air/S/O 2>2000 - W-h/kg100 - W/kgBGS10 kg Single-wing3 kg PBO-tether3 kg Inflatable SW1 kg PBO-tether2 kg Dual-wing0 kg CNT-tetherSensor Technologies MAP/MABS cameras 1-3 kg visible and IR cameras, 0.1-1 kg in situ OSL age118

Light-weight dropsondesAutonomous Navigation25 kg Netlander6.5 kg Deep Space 2MER Sun, Phobos/Deimos,star sensorsMER hazard recognitionspectrometers1 kg vector magnetometer0.1-1 kg in situ atmospheric, soilchemistry0.1 kg deployable microscopicimagers3 kg seismometer5 kgsurface penetratorsatmospheric profilersSurface feature recognition1 day trajectory forecastingBGS control estimation/strategiesdating5 kg, 10 km range LIBSspectroscopy5-10 kg radar0.1 in situ atmospheric, soilchemistry1 kg in situ mineralogy0.1-1 kgsurface penetratorsatmospheric profilersglidersdroppable rovers10 day trajectoryforecastingBGS advanced controlstrategies119

5.3 Enabling and Enhancing TechnologiesIn this section we provide details on enhancing and enabling technologies relevant to the DAREarchitecture development.5.3.1 Entry, Descent and Inflation (EDI)The EDI sequence involves the period of the mission from atmosphere insertion until the DAREplatform is operating nominally. It is described in detail in Section 3.6.2.The deployment and inflation portions of the EDI sequence include several technological issuesthat need to be solved including envelope sailing, oscillatory movements, and deployment shock.Current research speculates that the correct distribution of mass between parachute, gondola, andenvelope, can control the sailing phenomena. Without proper analysis the envelope is likely to betwisted and “sail” horizontally as compared with the vertically descending system. Oscillatorymotions in the system need to be damped or prevented. Lengthening of the tether connecting theballoon and gondola can mitigate some oscillatory movements. More sophisticated active controlmay be necessary. These issues are a subject of current research at NASA JPL.5.3.1.1 Deployment ShockSeveral approaches to aerial deployment exist. We have studied one of the approaches in whichthe balloon is folded and rolled onto a Styrofoam cylinder. During aerial deployment the gondolaand inflation hardware pull the balloon downwards, while the descent parachute pulls the balloonupward, quickly unrolling the balloon (see Figure 5-3).Figure 5-3. Fold then roll aerial deployment 23 .23 C. White and S. Day, “Analysis and Experimentation for the Aerial Deployment of Planetary Balloons”, 2003120

As the balloon goes taut at the end of this deployment sequence a strong shock force is generatedinside the balloon in the middle of the unfolded envelope. The force depends on the “separationvelocity” of the two ends of the envelope, which in turn depends on the length of the deflatedballoon. The longer the balloon – the higher is the separation velocity and the stronger is theforce.We estimate that for our point design of 34.4 m diameter balloon the separation velocity wouldbe about 24 m/s and the shock force would be from 3 to 6 kN (the uncertainty reflects differencebetween the experimental and modeling data).The strong force can be absorbed by an internal tether or a reinforced inflation tube inside theballoon (see Figure 5-4). Making the balloon material stronger to accommodate largedeployment stresses would result in a much larger mass increase (~10 kg), whereas the tether canbe very light (1 kg). The deployment shock would propagate towards the ends of the balloonwhere it would be absorbed by rip stitches.5.3.1.2 Main ParachuteFigure 5-4. Inflating balloon and inner tether/reinforced inflation tube.In the technology horizon matrix the parachute is shown as decreasing in mass as the horizonsprogress from 50 kg for the Current term, to 25 kg for the Near term and 0 kg for Far termtechnologies. The decrease in mass for the Near term technology is projected due to employmentof fabrics incorporating carbon-nano-tubes (CNT). We envision short strands of the CNTs beingincorporated into the fabrics making them very strong and able to withstand the requiredpressures during atmospheric deceleration of the entry vehicle during EDI. Further reduction ofthe mass of the main parachute in Far term technology horizon comes from the use of the longstrands CNTs in the fabric of the main parachute or from the use of the under-inflated balloonenvelope (also made out of the long strands CNTs) for deceleration. Reduction in the mainparachute mass can also come from the reduction in size, however this leads to shortening of thedescent and inflation times.121

5.3.2 Cryogenic Inflation Gas StorageUpon examination of the mass breakdown of previous Mars balloon missions it was evident thatthe inflation-gas tanks and hardware were very large contributors to the system mass (74 kg outof 200 kg). A small study was conducted to determine if mass could be removed from thissubsystem. Various storage methods were examined including:- metal hydrides (H 2 ),- micro-spheres,- room temperature compression,- cryo-compression,- liquid H 2 and others 24 .These are some of the technologies invoked by hydrogen-powered automobile researchers. Wehave analyzed the feasibility of storing the inflation gas in a liquid phase as opposed to theconventional gaseous storage method. It is believed that a liquid system would decrease therequired storage tank volume and therefore result in a decreased inflation system mass.One of the options considered was based on development by Schein et al. 25 . Schein andcolleagues designed and tested a cryogenic helium make-up gas system for extended durationballoon flights. The system had a storage ratio of 54%. We consider this work as the basis for theNear term technology cryogenic inflation gas storage system mass estimate, although theoutgassing rates of the Schein et al. system are lower than those required by the EDI.Using the storage ratio for the Shein et al. system we estimate the mass of the system at 15.4 kg(for 8.3 kg of gas). For the Far term technology we assume that cryogenic storage and CNTstorage tanks would enable storage ratios on the order of 100%.The cryogenically cooled gas would need to be heated up very quickly before inflation cancommence. The necessary heat to raise the enthalpy of the fluid to that required for atmosphericballoon flight is substantial but could be produced by simple combustion of hydrogen andoxygen. It is envisioned that a heat exchanger (a “boiler”) would transfer the energy ofcombustion to the liquefied lifting gas. For the conceptual system with the following parameters:• 8.3 kg - H 2• Inflation period of 200 s24 See for example http://www.fuelcellstore.com/information/hydrogen_storage.html25 Schein, M., Thomson, D., and Lachenmeier, T., "An Integrated Cryogenic Helium System for Extended Duration BalloonFlights," AIAA International Balloon Technology Conference, Norfolk, VA, June 28-July 1, 1999, Collection of TechnicalPapers (A99-33301 08-01).122

• ∆h = 2429 kJ/kg – required H 2 enthalpy change (assuming h store =463.7 kJ/kg, T store =33 K,P store = 1.3 MPa, h atm = 2893.0 kJ/kg, T atm = 205 K, P atm = 254 Pa; subscript “store” refersto values during storage, subscript “atm” refers to values upon expansion).the heating rate is approximately 100 kW and the total work done to the fluid is approximately20.2 MJ. Combustion of 0.14 kg H 2 and 1.13 kg of O 2 will provide this required amount of workin an ideal system. Assuming 50% efficiency in converting oxygen and hydrogen into heat, atotal of 2.5 kg of combustion gases is needed. This is a very small mass and volume of gas,which could be stored in relatively small tanks. We estimate that with current technology(Structural Composites Industries (SCI) 26 ) the oxygen and hydrogen for combustion can bestored in 2 cylindrical tanks (models AC-5166 and AII-449C) with a total mass of 6.9 kg or in 8tanks (all AC-5046) with the total mass of 4.3 kg. It is reasonable to assume that Near termtechnology would enable reduction of mass and in the number of tanks so that the total massdoes not exceed 2 kg and H 2 and O 2 are stored in two separate tanks.In the Far term the amount of buoyant gas increases to 12 kg and the mass of the combustiongases increases proportionally – to 0.4 kg of H 2 and 3.0 kg of O 2 . Assuming development ofmaterials that use CNT in their structure, the gas storage tanks can be made extremely light,virtually weightless. With these assumptions the mass budget of the inflation system for the threetechnology horizons is summarized in Table 5-4.Table 5-4. Inflation system mass budgetTechnology HorizonCurrent(3-5 years)Near(10-15 years)Far(25-40 years)Storage typePressurizedtanksCryogenicCryogenicBuoyant gas storage and inflationhardware mass, kg75 15.4 12Combustion gases mass, kg 0 2.5 3.5Combustion gases storage mass, kg 0 2 0“Boiler” mass, kg 0 1.5 1.5Total mass, kg 75 22 17Specific technologies required to advance this technological area are:- mass efficient cryogenic storage systems,26 C. A. Braun, “Innovative Features Providing Proven Solutions for Integration of Composite Pressure Vessels into AerospaceSystems”, AIAA 93-2099, AIAA/SAE/ASME/ASEE 29th Joint Propulsion Conference and Exhibit, June 28-30, 1993,Monterey, CA.123

- cryogenic storage with high outgassing rates, and- high efficiency hydrogen combustion chamber and heat exchanger.5.3.3 Make-Up Gas SystemFor the Current technology horizon the make-up gas systems may not necessary due to the shortproposed mission durations (tens of days). Estimated gas leak rates are slow enough to enablemissions with duration of up to 100 days for Current technology horizon (see Section 5.4.7 foranalysis of the gas leak rates for Martian balloons). For the Near term technology and forincreased required duration of the Martian missions we assume that improvements in themanufacturing process enable 10-fold reduction in the buoyant gas leak rates. The use of thecomposite balloon envelope material enables significantly reducing gas leakage through pinholesand due to diffusion. As the missions are predicted to last multiple years in the Far termtechnology horizon, it becomes necessary to provide make-up gas to account for gas lost due toleakage through manufacturing defects of the gore seals in the envelope. The amount of gas toreplenish is estimated at 1.6 kg/year (see Table 5-10 in Section 5.4.7) or 8 kg for a 5-yearmission.We looked at different options for the gas replenishment system. One of the options consideredwas based on development by Schein et al. 25 However, as it turns out, the system is not efficientfor the amounts of gas that needs to be stored in a long duration DARE mission (8 kg). Theamount of gas lost to cool the system turns out to be larger than the amount needed forreplenishment, which makes the whole system very large. We have rejected this approach for thereplenishment system. (Note that cryogenic storage of inflation gas requires a much smalleramount of gas for cooling during the trip to Mars because the storage tank has a much lowertemperature in space than in a planetary atmosphere).One of the possible approaches to a buoyant gas replenishment system includes condensingwater from the atmosphere of Mars and extracting hydrogen from the water via electrolysis. Wehave not studied this approach in detail. However, crude estimates indicate that this may be aviable option even for the low atmospheric water content observed at Mars.There is about 10 -4 kg of atmospheric water per kilogram of the atmosphere in the Northernhemisphere of Mars and about 10 -5 kg – in the Southern hemisphere 27 . The difference in thewater content between the hemispheres is due to sublimation of the water ice from the Northernpolar cap during northern summer (the polar cap in the southern hemisphere is CO 2 ). Assumingaverage relative wind speed of 5 m/s for the DARE platform, collection area of 1 m 2 for theatmospheric water vapor collection device, continuous operation and 10% combined efficiencyfor water vapor collection, the total amount of water that can be collected in one year in theNorthern hemisphere is:−42Mw≈ 10 ⋅ ρa⋅5m/ s ⋅88775sec/sol ⋅ 669.6sols⋅1m⋅ 0.10 ≈ 17. 8kg27 H. H. Kieffer et al., “Mars”, p.996, 1992.124

where ρ a is atmospheric density (0.006 kg/m 3 for 8 to 10 km balloon floating altitude). Hydrogenmass is 1/9 of the mass of the collected water or about 2 kg. This amount of hydrogen issufficient to replenish estimated buoyant gas loss. The amount of hydrogen collected in theSouthern hemisphere will be an order of magnitude smaller. To increase the hydrogen productionthe collection device can be made to have larger area or its efficiency in collecting water andconverting it to hydrogen improved.The specific technology required to advance this technological area is a high efficiencyelectrolysis system.5.3.4 Advanced Envelope MaterialsComposite materials combining Mylar for substrate stiffness and polyethylene for toughness andpinhole resistance were proposed in the past for Mars balloon envelopes 2 . It is possible that someadvanced material that can provide simultaneously good gas permeability resistance, fracturetoughness, and load bearing properties will be developed. However such a material is impossibleto speculate about, therefore we have chosen to speculate on the advancements of currentplanetary balloon envelope materials. Sample of the composite material is shown in Figure 5-5.Specific technologies required to advance this technological area are:- ultra-thin Mylar films (minimal pin-holing);- ultra-thin PE films (consistent thickness)- long chain carbon-nano-tube fibersFigure 5-5. Sample of the composite balloon envelope material 2125

5.3.4.1 MylarCurrent technology allows for the reliable production of 3.5 µm Mylar. Thinner films would beexcessively pin-holed as a result of manufacturing processes. We envision that Near termtechnology will allow for the production of 1 µm Mylar film. Ultra thin Mylar film is beingdeveloped for space applications, including solar sails and reflectors 28 . It is thought that a furtherreduction in film thickness is impractical from a manufacturing and assembly standpoint.5.3.4.2 PolyethyleneA polyethylene layer 6 µm thick is envisioned as the current technology level. Near and far termtechnologies will allow for 3 and 1- µm film thickness, respectively. The polyethylene layer willprovide little strength and minimal performance as a gas barrier but it adds a level of redundancythat is important. The high fracture toughness will supplement the Mylar’s brittle nature and thelow quantity of pinholes will further retard the diffusion of the lifting gas through the material.5.3.4.3 ScrimThe woven scrim that supports the loads of the envelope, including the envelope super pressure,in the current technology horizon is 55-denier Kevlar (Denier is one unit of measure for thelinear mass density of fibers. Denier is defined as the mass in grams per 9000 meters). Near termtechnology will provide 38-denier PBO scrim. The future technology horizon utilizes 1-denierCNT thread.5.3.4.4 Film Thermal-Optical (TO) PropertiesFilm TO properties are important for limiting the super pressure that must be contained by theenvelope. It is thought that the best combination of properties results from the aluminization ofthe top half of the envelope while the bottom half is painted white. This reduces the heat lossfrom the envelope in the very cold atmosphere of the polar region and prevents condensation ofthe atmospheric CO 2 on the surface of the envelope. The condensation would not damage theenvelope, but it would probably create a CO 2 snow cloud or fog around the DARE platform,which could have a detrimental effect on observational capabilities of the instruments in thegondola. CO 2 in a solid phase may also condense on the platform making it heavier and causinga loss of altitude. For these reasons condensation should be avoided.In the far term technology horizon the envelope is envisioned as being covered with a variableTO film that would allow for the active control of lifting gas temperature. A balloon made out ofsuch materials would be able to change the reflectivity of its envelope depending on theenvironment and in this way to enjoy a reduction in the envelope’s stress due to the reduceddiurnal variations of the film and buoyant gas temperatures. This would enable a reduction in themass of the envelope due to less stringent envelope strength requirements.28 http://www.stg.srs.com/atd/Solar_Sail.htm; M. Leipold et al., “Interstellar Heliopause Probe: System Design of a Solar SailMission to 200 AU” , submitted to AIAA Journal of Spacecraft & Rockets.126

5.3.5 Primary StructureCurrent technology levels allow for structural materials with a 16-fold strength-to-weight benefitover conventional aluminum (graphite composites). Near and Far term technology horizons willutilize CNT composites with significantly higher strength to weight ratios – 300 and 612,respectively 29 . The use of CNTs will provide significant reduction in the masses of the primarystructure, main parachute, balloon envelope and buoyant gas storage.5.3.6 Power GenerationAll technology horizons will utilize photovoltaic arrays and only small increases in efficiency areenvisioned. Current GaAs multi-junction photovoltaic arrays (“solar cells”) can have 25%efficiency in converting solar power to electric current. The mass densities of such arrays(including the masses of the cells and the substrate) are about 2.4 kg/m 2 30. Advancedtechnologies may provide decreases in array mass but increases in power output are limited. Inthe Near term the efficiency of the multi-junction arrays can increase to 33%, with thecorresponding decrease in mass density to 2 kg/m 2 . In the Far term the efficiency of the arrayscan further increase to 42% and mass density decrease to 1.6 kg/m 2 with the combination ofmulti-junction arrays and quantum dots (QD). A quantum dot, also called a semiconductornanocrystal, is a semiconductor crystal whose size is on the order of just a few nanometers. Theycontain anywhere from 100 to 1,000 electrons and range from 2 to 10 nanometers, or 10 to 50atoms, in diameter 31 . A schematic drawing in Figure 5-6 shows the arrangement of atoms in aQD.The NASA Glenn Research Center (GRC) has been investigating the synthesis of quantum dotsfor use in solar cells 32 . Using quantum dots in a solar cell allows harvesting of a much largerportion of the available solar spectrum. In a paper published in the May 2004 issue of PhysicalReview Letters a team from Los Alamos National Laboratory found that quantum dots produceas many as three electrons from one high-energy photon of sunlight 33 . When today's photovoltaicsolar cells absorb a photon of sunlight, the energy gets converted to at most one electron, and the29 Frankland, S.J.V., et.al., “The Stress–Strain Behavior Of Polymer–Nanotube Composites From Molecular DynamicsSimulation,” Composite Science and Technology 63 (2003) 1655-1661; Demczyk, B.G., et al., “Direct MechanicalMeasurement Of The Tensile Strength And Elastic Modulus Of Multiwalled Carbon Nanotubes,” Materials Science andEngineering A334 (2002) 173–17830 G. Landis, T. Kerslake, P. Jenkins and D. Scheiman, “Mars Solar Power”, NASA/TM 2004-213367, AIAA–2004–5555.31 http://en.wikipedia.org/wiki/Quantum_dots32 http://www.grc.nasa.gov/WWW/RT2001/5000/5410bailey1.html33 R. D. Schaller and V. I. Klimov, “High Efficiency Carrier Multiplication in PbSe Nanocrystals: Implications for Solar EnergyConversion”, Phys. Rev. Lett. 92, 186601, 2004.127

est is lost as heat. Theoretical studies predict a potential efficiency of 63%, which isapproximately a factor of 2 better than any state-of-the-art devices available today 34 .5.3.7 Supplemental PowerFigure 5-6. Quantum dot schematic drawing.For system initialization and instances requiring extra power above what can be provided by thesolar array and secondary batteries it may be necessary to carry a modicum of supplementalpower storage. Conventional primary cell batteries provide energy densities of 350 Whr/kg andpower densities of 350 W/kg. For the Near term technology horizon we envision tritiatedamorphous silicon (ASi) cells as the source of supplemental power 35 . These are betavoltaic cellsthat directly generate electricity from the constant radioactive decay of the embedded tritium.Tritium is being consumed over the lifetime of the cell, but the halftime of the tritium decay (12years) is much longer than any mission duration, so that the decrease in the cell efficiency overmission lifetime should not be noticeable. Tritiated ASi cells are being studied at the NationalRenewable Energy Laboratory (NREL). These cells will provide energy densities of 1 MWhr/kgand power densities of 30 W/kg. In the Far term technology horizon QD photovoltaic arrays thatcan generate power at multiple wavelengths may further increase energy and power densities forthe supplemental power devices 36 . However, there are no estimates of the possible energy andpower densities for these devices.34 Luque, A.; and Marti, A., “Increasing the Efficiency of Ideal Solar Cells by Photon Induced Transitions at IntermediateLevels”. Phys. Rev. Lett., vol. 78, no. 26, 1997, pp. 5014-5017.35 http://www.physorg.com/news4081.html36 http://www.evidenttech.com/applications/quantum-dot-solar-cells.php128

5.3.8 Energy StorageEnergy storage is seen as an enhancing technology that will likely be developed to advancedlevels through commercial research. Current technology allows for energy densities of 150Whr/kg and power densities of 150 W/kg with Li-Ion Polymer batteries. It is envisioned that inthe Near term technology horizon regenerative fuel cells with energy densities of 400 Whr/kgand power densities of 120 W/kg will become available. For the Far term technology horizon weenvision Li-Air or Li-S technologies that have theoretical energy densities in excess of 2000Wh/kg and power densities of 100 W/kg. Companies that are actively researching thesetechnologies include Excellatron 37 .5.3.9 Balloon Guidance SystemThe BGS is an enabling technology that allows for guided planetary balloon missions. A detaileddescription of the operations of the BGS is given in Chapter 4. The unique challenges ofoperating a BGS in an extraterrestrial atmosphere require some advanced technologies. At thecurrent technology levels we see a relatively heavy (10 kg) single wing BGS of moderate size (1-3 m 2 ) with a tether made of PBO. Technological advancement for the Near term technologycome from employing a large (6 m 2 ) and light (3 kg) inflatable single-wing BGS with the PBOtether. The Near term BGS houses scientific instruments, camera actuators and attitudecontrollers in its boom, together with the batteries and solar arrays that may be flexible and coverboth sides of the BGS inflatable wing surface. In the Far term technology horizon we envision alight (potentially inflatable) and large (>10 m 2 ) dual-wing BGS wing, and light (1 kg for 10 kmlength) and thin tether made with the use of CNT materials.5.3.9.1 Single WingThe conventional single wing system looks very much like a keel on a sailboat (see Figure 5-7).It has no ability to orient its lift vector. In strong relative winds the wing generates strong liftingforces that tilt and literally lift the BGS wing to higher altitude, which negatively impacts BGSperformance by reducing the magnitude of the force available for trajectory control. The dualwing system discussed in Section 5.3.9.3 is designed to alleviate this problem.37 http://www.excellatron.com/129

Figure 5-7. Single-Wing BGSWith Current level of technology development, a small single-wing BGS at Mars is as efficientas a dual-wing BGS because of the small forces and large weights of the single-wing BGS.5.3.9.2 Inflatable Single Wing BGSOne of the ways to decrease the mass of the BGS wing is to make it inflatable. This sectiondiscussed the concept for inflatable BGS wing.ILC Dover has studied inflatable airfoils for use in gun-launched unmanned aerial vehicles.Figure 5-8 and Figure 5-9 depict early research into inflatable wings at ILC Dover.Relatively high stiffness is achieved by maximizing the inflated section’s area moment of inertia.Current work in inflatable wings is being conducted at NASA Dryden. The small packed volumeof an inflatable aircraft makes it suitable for sensing operations from high altitude platforms andpossibly planetary exploration. In order to achieve an optimal airfoil shape the ILC Dover wingsare composed of several small cells and covered with a “smooth” skin. The cellular design alsoincreases the wing’s bending stiffness.130

Figure 5-8. ILC Dover inflatable wing experiment.Figure 5-9. ILC Dover gun-launched conceptFrom the BGS trajectory modeling it was decided that a DARE BGS would need a large surfacearea, a low aspect ratio, and a high coefficient of lift. Other requirements imposed by the natureof the DARE mission include lightweight design, small packed-volume, and long missionlifetime. The packed volume limitation is directly related to the entry capsule chosen for theconceptual DARE mission. This limitation coupled with the goal of maximizing the wing surface131

led to the inflatable BGS design. It is thought that an inflatable BGS can best meet all of therequirements described for the DARE-Martian-BGS.Figure 5-10 shows the conceptual Martian inflatable DARE BGS. The drawing on the left showsthe external view of the BGS, while the drawing on the right shows the internal structure of theBGS wing sections.Figure 5-10. Inflatable DARE BGS.The inflatable BGS is designed with 8 distinct wingspan-sections. This allows for easyconstruction and deployment redundancy (mission could continue with deflation or improperinflation of 1 or more sections). The wing sections employ the cellular design as described byILC Dover. Previously we discussed low Re number airfoils and the shapes that provide optimalcoefficients of lift. With this research in mind the conceptual inflatable wing has a planar crosssectional area with no camber and no thickness variation. As an added benefit it may be shownthat the sectional design of the wing is modular and scalable.Inflating and maintaining the airfoil’s shape is necessary for operation throughout the entiremission. Initial inflation is requires small amount of gas that can be stored in a small container onthe BGS boom. The BGS wing can also be inflated with ambient air – this would require a smallcompressor positioned on the BGS boom.132

After the initial inflation the airfoil must maintain its shape for the duration of the mission. Thiscould be achieved by providing periodic make-up gas injections using the compressor systemdescribed above. Alternatively the wing could be constructed of a rigidizing material. Several ofthese materials are under investigation for use in lightweight aerospace applications. There areshape, temperature and UV rigidizable materials currently available. This solution would ensurethat the BGS wing is optimally shaped without the need for a make-up gas supply.The boom of the BGS folds into three sections for storage in the entry capsule. The boomsupports the solar arrays, gimbaled camera platform and the inflatable rudder (Figure 5-11). Thesolar arrays are sized according to the solar power study and fold around the boom for storageduring launch and the cruise stage. The science payload is supported on a gimbaled platform thatpoints all of the cameras simultaneously. The position of the camera platform is such that thecenter of mass, fore-to-aft is below the tether. The rudder employs the same inflatable design asthe main wing and is supported below the boom in the “clean” undisturbed airflow.Figure 5-11. Lower portion of the DARE BGS showing the boom with the rudder, solar panelsand the gimbaled platform with the imaging cameras.5.3.9.3 Dual WingThe dual-wing system provides improved balloon guidance control, especially in strong windconditions. It allows for controlling the orientation of the lift vector in order to pull the BGS backdown into denser air. As the system must support bending moment loads not present in the singlewing system it is expected to have a greater mass per wing area than the single wing BGS.Advanced construction materials will allow for ultra-lightweight wings. Lightweight wings will133

have a greater tendency to slip upwards as the lift force approaches the mass of the wing. Thiswill necessitate the use of a dual wing system.5.3.9.4 WinchFigure 5-12. Dual-Wing BGS.The winch at the gondola enables deploying the BGS wing at the start of the mission and reelingthe BGS wing up and down, as may be required by trajectory guidance techniques (see Section4.4.5.1).There is a need to control the descent of the BGS as it is lowered into operating position severalkilometers below the gondola. The tether winch will provide this control. Currently there are twomethods being proposed. The first, electromechanical, method would convert the energy ofdescent into electrical energy and dissipate it through electric lights and possibly a lowtemperature radiator. Alternatively the electrical energy could be used to charge or “top-off”batteries that have been dormant for the journey from Earth to Mars. The second “thermomechanical”method would convert all of the descent energy into thermal energy. This couldheat a thermal sink carried on the gondola or possibly use the Martian atmosphere as a sink.5.3.9.5 TetherThe tether connecting the BGS wing to the gondola is small but important in terms of its mass,and is a significant contributor to the aerodynamic drag of the system. The Far term technologyhorizon envisions CNT materials to reduce the mass and the diameter of the tether. In the134

Current and Near term technology horizons PBO tethers that have masses in the order ofkilograms are utilized.Heavier BGS wings require tethers with larger diameters and higher strength to support theweight of the mass at the end of the tether and the weight of tether itself. Given the physicalcharacteristics of the tether material, the required cross-sectional area of the tether is calculatedfrom the following relationship:At( M + M ) g ( M + A L ρ )w tw t t tg= SF= SF,σσwhere SF is the safety factor, M w is mass of the BGS wing, M t is the mass of the tether, g isgravity, L t is tether length, ρ t is the density of the tether and σ is ultimate strength of the material.This equation can be solved for A t . For Near term BGS wing M w =13 kg (including the masses ofthe cameras, actuators, batteries, etc.), SF=3, L t = 10 km, and PBO σ=5800 MPa and ρ t =1560kg/m 3 the estimated required tether diameter and mass areD t =0.2 mmM t =0.4 kgThe volume of the tether is about 250 cm 3 . For comparison the volume of a soda can is 355 cm 3 .Hence even a very long tether occupies a relatively small volume. With the introduction of theCNT-based materials in the Far term technology horizon the tether will remain small and lightwhile supporting a heavier BGS wing.5.3.10 Sensor TechnologiesDevelopment of lightweight and more capable sensors is one of the enabling technology areasthat would enable revolutionary science from DARE platforms.As an example of the current state-of-the-art balloon sensor technologies we list conceptualdevelopment of the narrow and wide field cameras, and thermal emission spectrometers in theMAP and MABS projects. These cameras were projected to have masses on the order of akilogram and apertures of the order of centimeters.For the Near term technology horizon we envision the need for slightly larger, but still lightvisible and infrared cameras and spectrometers that would have masses from 1-3 kg. Thesecameras would enable ultra-high-resolution imaging of the surface, surpassing the resolvingcapabilities of the HiRises instrument on the MRO by an order of magnitude. Section 5.4.4discusses sizing of the optical instruments for the Near term DARE platform payload.Lightweight proton magnetometers (1 kg) would be used to make accurate measurements of themagnitude and direction of the magnetic field of the Martian magnetic anomalies. Reducing themass of the magnetometers would enable carrying several of these instruments positioned alongthe BGS tether – enabling measurements of the vertical profiles of the magnetic field.135

Small and very light (0.1-1 kg) droppable sondes would enable in situ analysis of the atmosphereand of the Martian soil’s chemistry. Reducing the masses of these sondes would enable carryingmany of these instruments and deploy them over different sites on the planet. Atmosphericsondes would be able to measure profiles of atmospheric temperature, pressure, dust distribution,water content, winds – and maybe other parameters. The probes would transmit the data to theDARE platform above via a small radio. Probes landing on the surface would be made light toreduce impact energies. These surface probes would run a sequence of short tests to determinechemical composition of the soil and transmit the data to the DARE platform. The batteries andcommunication packages of the surface landing probes will be miniaturized to enable operationover just a short period of time and transmit data over relatively short distances (10 – 20 km) tothe overflying DARE platform. Some of the surface landing probes may contain a very small andlight (0.1 kg) microscopic imager that would image the surface at just one spot at the landing siteat high magnification. Deploying many of these probes over a variety of the sites would enablemicroscoping imaging capabilities similar to that of the existing Martian rovers, but with theadded benefit of the global geographic coverage. Deploying these microscopic imagers over amultitude of widespread Martian surface sites would enable studying the small-scale morphologyof the Martian surface on unprecedented scale.A lightweight (3 kg) seismometer would be developed for incorporation into deployable surfacestations for seismological and meteorological networks.In the Far term technology development horizon we envision development of disposable in situlightweight (0.1-1 kg) age dating instruments, for example based on the Optical SimulatedLuminescence (OSL) technique (see Section 3.4.3). These sensors would be incorporated intosurface landing probes. While on the surface, the sensors would analyze the soils and rocks todetermine their age. The surface landing probes equipped with these sensors would be deployedover different sites to determine, for the first time, the ages of various geologic formations atMars with high accuracy.The DARE platforms would carry advanced remote instruments for determining the elementalcomposition of the soils and rocks. One possible technology that would make these instrumentspossible is the Laser-Induced Breakdown Spectroscopy (LIBS, see Section 3.4.2). Theinstruments, housed in the DARE platform’s gondola would shine a laser beam at a rock sampleat the surface underneath the platform, ablate its surface and analyze the spectrum of the vapors.The range of the LIBS technique operation would need to be extended to 10 km to enableobservations of the surface from the altitude of the DARE platform. The instrument would be analternative to the use of disposable droppable analyzers described above for the Near termtechnology. If the range of LIBS instrument operations can only be extended to 1 km, theinstrument can be housed at the BGS wing structure. The instrument mass would need to bereduced down to 5 kg to reduce the stress on the BGS tether.Weights of the deployable in situ surface and atmospheric probes would be reduced to 100 g toenable carrying hundreds of such probes and sampling hundreds of sites at Mars.The DARE platforms would carry a lightweight (5-10 kg) radar for high-resolution observationsof the Martian subsurface. Closeness of the DARE platform to the surface would enable imagingthe structure of the subsurface at Mars with high vertical and horizontal spatial resolutions.136

Small lightweight (1 kg) in situ probes capable of analyzing the mineralogical composition of thesurface would be developed in the Far term. These probes would operate like the surfacechemistry probes described above, except they would analyze mineralogy of the surface rocks,which is a much more complex task.5.3.11 Light-weight DropsondesA number of the sensors described in the previous Section are deployed via dropsondes. Weenvision the need for dropsondes for atmospheric profiling and for dropsondes that can survivethe surface impact and operate for a short time on the surface.As an example of the Current technology state we list ESA’s Netlander 10 and NASA’s DeepSpace 2 (DS 2) probes 11 . Both of these systems deploy from space and require entry heatprotection in the form of a heat shield on the entry vehicle. We envision dropsondes with similarfunctionality, but with reduced total system mass and with improved energy generating andstoring capabilities. The mass reduction would come from removing the heat protection and fromminiaturization.In the Near term technology horizon we envision surface probes that are nodes of theseismological and meteorological network. These are relatively light (5 kg) and can survive onthe surface of Mars for about 1 Martian year. They are deployed from a DARE platform abovethe selected site. Impact velocities are comparable to that of the probes entering from space(Netlander, DS 2) because of the rarified Martian atmosphere. Impact velocities can be reducedsomewhat by employing relatively small parachutes (1 m 2 ) – see Section 3.4.4 on the analysis ofthe impact velocities. The impact energy is dissipated by the lower part of the dropped probepropagating into the subsurface in a way similar to that of the DS 2 probes. The aft section of thelanded station remains on the surface and deploys solar panels for solar power generation andcontains the meteorological payload. The forward section penetrates by several ten’s ofcentimeters into the subsurface and provides surface contact for the seismometer and serves as aprobe for a possible soil experiment. A small temperature sensor in the penetrator would enableanalyzing heat transport into the subsurface. We see the main technological challenge for thissystem in developing lightweight power generation and storage capabilities. A large fraction ofthe Netlander mass is power generation and storage systems. Solar power generation capabilitieslimited deployment of the Netlanders to ±30º latitude (see Figure 3-19). To enable operation ofthe network nodes in higher latitudes would require development of alternative power generationcapabilities. Surface stations would communicate directly with the Mars communications orbiter.Smaller surface probes would house the in situ surface sensors described in the previous Section.For these sensors the power generation and storage requirements, and communicationrequirements are less stringent than for the network surface stations. The atmospheric profilerswould house the sensors, batteries and telecommunication payloads. The landing dropsondeswould provide impact protection for the science payload for high impact velocities. The landingsondes would penetrate into the subsurface to dissipate the impact energy and to bring thesensors closer to the target.For the Far term technology horizon we see reduction in the mass of the atmospheric profilersand subsurface penetrators down to the masses of 0.1-1 kg. In addition to atmospheric profilers137

and surface stations the DARE platforms would carry “gliders” - probes with limited path controlcapabilities that would enable more accurate targeting of the surface targets; and small droppablerovers that would be able to operate for a couple of days in the vicinity of the landing site andapply their very simple science payloads (for example, a microscope) to 2 or 3 targets.5.3.12 Autonomous Navigation and ControlTo enable the revolutionary science observations described in Section 3.3 the DARE platformswould need to be able to make decisions regarding observation targets or trajectory control “onthe fly”, e.g. without human operator. The platforms will be able to receive commands fromEarth via Mars communication orbiter over limited periods of time during a day. Sometimescommunications will be impossible for several days in a row – for example during Earth-Marsoppositions and Mars solar conjunctions when the sun prevents communicating with Mars. Inaddition, long Mars communication times (hours) make Earth-based operation of the Marsexploration assets inefficient. The DARE architecture would employ advanced autonomousnavigation and control techniques to perform science and engineering activities without help of ahuman operator.The field of autonomous navigation and control for planetary exploration platforms is rapidlydeveloping. As an example of the current level of technology development, Mars explorationrovers (MER) are able to drive semi-autonomously for several feet on the surface of Mars, afterreceiving carefully estimated path guidance commands from Earth. The rovers can recognizeterrain obstacles and hazards and stop. MERs are semi-autonomous because the future drivingpath is still calculated by a human operator. The location of the MERs is determined fromobservations of the Sun, stars and Martian satellites Phobos and Deimos with the onboardcameras. Since the rovers only travel several meters a day, orbital imagery and surface featurerecognition (by human operator) are also used to pinpoint the location of the rovers.The DARE platform moves much faster than the MERs (30 m/s vs 30 m/day), so the controldecisions will need to be made much faster – every hour or maybe every 15 minutes. The DAREplatforms cannot stop at a target and wait for the communication opportunity to receiveinstructions on the follow up action. In the Near term technology horizon we envision DAREplatforms carrying in its onboard computer the coordinates of hundreds or maybe thousands ofobservational targets with sets of action settings for each target. The DARE platform wouldimage the surface and the sky with the medium resolution context camera to pinpoint its locationvia surface feature recognition algorithms and by tracking sun, stars and Phobos and Deimos.Communications with the orbiter will also be used to estimate the location of the platform. Thesurface feature recognition algorithms would use high-resolution digital elevation maps (similarto that developed from Mars Orbital Laser Altimeter – MOLA - data) stored on the onboardcomputer. When the platform floats in the vicinity of a target of opportunity, the onboardcomputer would estimate the control requirements to overfly the target, compare it to estimatedBSG capabilities in the current environment, evaluate whether this excursion to a target wouldinterfere with the other trajectory objectives (i.e. topography collision, long term goal of crossingequator, etc.) - and so forth. If the target appears to be within reach, the computer will employ acontrol algorithm to reach the target (similar to the ones we used in our trajectory simulation)and issue control commands. Control algorithms would benefit from forecasting atmospherictrajectories. For the Near term technology horizon we envision that 1 day forecasting would138

ecome available through global circulation numerical modeling and observation of Martianwinds and BGS performance models. The biggest uncertainty in atmospheric trajectoryforecasting, as on Earth, is the uncertainty in wind measurements. An orbiter with a Dopplerwind Lidar instrument, combined with remote observations of global temperature fields, in situmeasurements of the environmental parameters via Mars surface stations and a DARE platformshould greatly improve our ability to predict Martian winds and atmospheric trajectories.In the Far term technology horizon the autonomous navigation and control capabilities willinclude 10 day trajectory forecasting and BGS advanced control strategies. Long-term trajectoryforecasting will improve accuracy and precision of targeted observations, increase the safetymargin of the mission, and increase the scientific output of the mission. The advanced controlstrategies may employ complex searches for the best possible trajectories through ensembles offorecasted atmospheric trajectories (similar to current weather prediction techniques).5.4 Design Trades5.4.1 Balloon AltitudeOn of the important design parameters is the operational altitude of the DARE platform. Moreaccurately, the design parameter of importance is the atmospheric density at balloon altitude,since the atmospheric density at which the balloon is neutrally buoyant determines the geometricaltitude of a superpressure balloon.The choice of the balloon altitude is driven by several conflicting requirements:- the balloon is required to clear most of the Martian topography (see Figure 3-9), which favorshigher altitudes;- high-resolution observations of surface features require lower altitudes;- ground speed requirements (if they exist) may call for a certain floating altitude, asatmospheric winds in general increase with altitude;- operational stability of the balloon platform, which may be required for instrument pointingand BGS performance, requires avoiding regions of boundary layer atmospheric turbulence;- lower floating altitudes require smaller and thus lighter balloons;- higher altitudes are favored for increasing the control capabilities of the BGS by enablinglonger tether lengths and higher relative velocities at the BSG wing altitudes;- lower altitudes and smaller balloons improve BGS performance due to smaller aerodynamicdrag forces acting on a balloon.As can be seen from the list above, these requirements cannot be simultaneously satisfied.139

Figure 5-13. Northern polar region MOLA topographyIn our analysis we consider altitudes from 0 to 30 km above the reference level (Martian areoid).At the altitude of 0 km the balloon can only operate in the Northern Polar Regions (northward of60°, see Figure 5-13). At this altitude the balloon will be 4-5 km above the surface, and 2 kmabove the highest point of the polar ice cap. This could be an acceptable set up for a missionstudying the Northern Polar cap and the layered polar terrain.Figure 5-14 shows the maximum possible payload masses and balloon radius as a function ofaltitude above the reference level for a Martian balloon and the three technology horizonsdiscussed in 5.2.1. Solid green, red and blue curves mark the payload masses (left axis) for theCurrent, Near and Far technology horizons, while the dashed curves mark the radius of theballoon (right axis) for the same technology horizons.140

Figure 5-14. Maximum payloads and balloon radius for DARE platform design.Note that for the Near term and Far term technology horizons balloon sizes become comparableto that of the large NASA Long Duration Balloons being launched on Earth for astrophysicsmissions.Altitudes of the constant density levels vary significantly throughout a year at Mars. Hence, weneed to establish an understanding of the variability (temporal and spatial) of the constantatmospheric density levels in the atmosphere. A superpressure balloon of constant volume willfollow the constant atmospheric density level in the atmosphere.We used the Mars-GRAM 2001 model of the Martian atmosphere to study the structure andvariability of the Martian atmosphere (see Section 4.4.1 for the description of Mars GRAM).Consider, for example, an atmospheric density level of 0.0058 kg/m3. This density levelcorresponds to an altitude of roughly 10-km above the reference level at Mars. At this altitude aMartian super-pressure balloon will float about 6 km above the Southern highlands and about 14km above the Northern lowlands. (This large difference in floating height above the surface hintsat usefulness of varying the tether length depending on geographic location on Mars. A winchsystem can be implemented to vary the tether length).141

Figure 5-15 shows daily limits of variations of the height of the density level of 0.0058 kg/m 3across the planet at southern solstice (the beginning of summer in the Southern hemisphere -Ls=270°). The height is given in terms of difference from the 10-km altitude above the referencelevel, so that -1 km means that the density level is at 9 km.Figure 5-15. Variations of the height of a constant density level across MarsFigure 5-15 shows that the atmosphere is the warmest (constant atmospheric density level is atthe highest altitude) over the equator and the coldest over the North Pole (where it is polar night).The height of the constant density level varies by about 0.5 km over the summer hemisphere anddecreases by about 3 km over the North Pole.The daily variability of the height of the constant density level (representing 1-σ variability aboutthe mean), indicated on the plot by the orange and blue curves, is less profound than the spatialvariability. The height of the constant density level varies the most over the equatorial regions,where it is about 300 m at this season.The analysis shown on Figure 5-15 can be extended to all seasons. Figure 5-16 shows thevariations of the height of the atmospheric density level 0.0058 kg/m 3 across the planetthroughout a Martian year.142

Figure 5-16. Annual limits of the height variability for a constant density level at MarsThe plot was constructed by combining the minimum and maximum heights (at 1-σ level) fordifferent seasons. As can be seen, the height varies by about 2 km for all locations throughout ayear. This is variability at 1-σ level - it can be expected to be observed with the probability of60%. For 99% probability (3-σ) the variability is about 3 times larger - 6 km (±3 km about 10 kmaltitude).The height of the density levels also changes with the amount of the dust in the atmosphere. Asthe atmospheric dust absorbs solar radiation and heats up the atmosphere, the density of theatmosphere changes. Figure 5-17 shows variations of the floating altitudes of the DARE platformpayloads as a function of atmospheric dust content. The level of atmospheric dust optical depth τ(tau) of 0.1 corresponds to very clear atmosphere with low dust content. Dust levels of τ=0.3and τ=3 correspond to medium and high atmospheric dustiness, respectively. Atmosphericoptical depth τ=3 is only observed during global dust storms.As can be seen from Figure 5-17, atmospheric density does not change significantly for amoderate increase in the atmospheric dust content (from τ=0.1 to τ=0.3 For a very dustyatmosphere (τ=3) the floating altitude of a balloon would decrease quite significantly - by about3 to 4 kilometers, depending on the altitude.143

Figure 5-17. Floating altitudes for various payloads as a function of atmospheric dust loadingThe worst case scenario for a balloon at Mars is to be over the southern highlands during a globaldust storm – high topography of the southern highlands (2-3 km) and the drop in the balloonfloating altitude may end the mission prematurely. Hence the main requirement for the DAREballoon floating altitude becomes the ability to maintain a safe altitude even during a dust stormwhile over the highlands of the Southern hemisphere. We assume that a safe altitude is 3 kmabove the surface. Thus we chose the nominal floating altitude of 10 km for the DARE balloondesign effort. At this altitude the DARE platform would carry an 87 kg payload and have aballoon that has a radius of 17.2 m (Figure 5-14). In a dust storm the altitude of the DAREplatform would decrease from 10 to 6 km, leaving 3 km above the high topography of southernhemisphere highlands.5.4.2 Martian EnvironmentIn this section we discuss the effects of the Martian environment on the design of the DAREplatform and present various design trades.One of the important parameters of the planetary balloon design is the amount of buoyant gas.This amount must be sufficient to keep the balloon fully inflated at all times. This means that thebuoyant gas has to have higher pressure than the ambient atmosphere even at the lowest possiblegas temperature. In other words the balloon has to maintain its superpressure. When a144

superpressure balloon loses its superpressure, its volume decreases, the balloon looses positivebuoyancy and drops to the surface. Hence, one of the important properties of the environment isthe lowest buoyant gas temperature that it can create.With a very high degree of accuracy the lowest buoyant gas temperature will be found duringnight and over surface spots with the lowest temperature. Previous Mars balloon studies(MABS 2 , MAP 3 ) looked at various worst-case conditions using various models. In particular,MABS used results of a one-dimensional model that provided surface temperature. We usedMars-GRAM 2001 model to repeat this analysis, since Mars GRAM is a more advanced tool andshould provide better results.In addition, new observations and data since the time of MABS and MAP analysis provide newunderstanding of the Mars environment. For example, observations with the MGS TESinstrument indicate that the lowest brightness temperature, corresponding to the tops of the polarclouds can be as low as 115 K (see Figure 5-18). The gas temperature of a balloon flying oversuch a cloud can drop much lower than in any environment without the clouds. This situationwas not considered by any of the previous Mars balloon studies, since clouds are not expected tobe encountered outside the Polar Regions, and none of the previous studies looked at balloonsoperating in Polar Regions.The temperature of the balloon film of the DARE platform flying over (or inside) a cloud woulddrop below the condensation temperature of the CO 2 and the CO 2 frost will start to condense onthe envelope. While we don't believe this may cause damage to the envelope, it may add mass tothe whole system. Possible ways to counter the condensation is to avoid flying over the polesduring polar nights (when clouds are more likely to be encountered), to heat the buoyant gas, touse passive means to keep the gas temperature high (use low emitting materials as film coating tokeep the heat absorbed during the day). At any rate, this situation should be modeled in moredetail. However, it can be considered exotic, i.e. never encountered outside particular short timeperiod and limited locations.145

Figure 5-18. Clouds over North Pole on Mars 38 .We now turn to estimating the lowest buoyant gas temperatures experienced by balloons at Marsfor the less exotic case without the polar clouds.38 A. Ivanov, D. Muhleman, “Observations of the Polar Night Clouds on Mars with the Mars Orbiter Laser Altimeter (MOLA)”,Mars Polar Science 2000, paper 4097.146

The minimum required mass of the buoyant gas is obtained by setting the night superpressure tozero: ∆P night =0. In this case the mass of the required buoyant gas can be expressed as:m gas =M f *R a T a /R g T ghere m gas is the mass of the floating gas (helium), T a – temperature of the atmosphere at thealtitude of the balloon, T g – temperature of the gas, M f – floating mass (defined for each of thetechnology horizons), R a and R g are the atmosphere and buoyant gas constants, respectively. Ascan be seen, the worst case for balloon design – the maximum gas mass – is found for maximumTa/Tg.Extensive analysis using Mars GRAM model and several balloon designs with different thermoopticalproperties of the balloon film (see Table 5-5) yielded some interesting results.Table 5-5. Thermo-optical properties for various DARE balloon designs used in the analysisCase# α top α bottom ε top ε bottom COMMENTS1 0.25 0.25 0.9 0.9 MABS design used in peak superpressureanalysis. Low film temperature at night.2 0.25 0.25 0.06 0.9 MABS design – white paint bottom andpolished metal top – used in carbon dioxidecondensation analysis. Uses incorrect valuefor metal absorptivity. Increases the filmtemperature at night3 0.03 0.03 0.86 0.86 Hypothetical high reflectivity materials –possibly made of multiple reflective layers.Low film temperature at night.4 0.25 0.25 0.06 0.06 Fully metallized balloon. Low filmtemperature at night (unless radiative coolingtime is longer than the night). Increasedradiative cooling time. Highest peak daytemperature.5 0.25 0.25 0.9 0.06 Metallized bottom, white paint top. Lowestnight temperature. Lowest peak daytemperature.6 0.09 0.25 0.06 0.9 MABS design – aluminized top, white paintbottom. Uses thermo-optical properties forpolished aluminum and white paint.The Ta/Tg ratios were calculated for a number of cases using the Mars GRAM data andcalculating Tg from a simple balloon model balancing atmospheric infrared fluxes at night and147

gas temperature (no convective effects, temperature of the balloon film is assumed to be equal tothe gas temperature at all times):Tg⎛ ⎞⎜εtF ⎟u+ Fd⎜ ε ⎟b= ⎜ ⎟⎜ ⎛ ε ⎞1 ⎟⎜⎜ +tσ⎟⎟⎝ ⎝ εb ⎠ ⎠1/ 4,where σ is Stephan-Boltsman constant, ε t and ε b are emissivities of the top and the bottom of theballoon material, respectively, F u and F d are upwelling and downwelling radiative fluxes,respectively.The buoyant gas temperature is strongly dependent on the upwelling infrared flux, which in turndepends on the temperature and radiative properties of the Martian surface. Absorptive, emissiveand thermal inertia of the surface material affect upwelling infrared flux. The lowest surfacetemperatures are found at night at the sites with low thermal inertia and high emissivity.However, sites with low infrared flux also have lower atmospheric temperatures, so it not clear apriori where we would find maximum values of the Ta/Tg ratios. We searched through datagenerated with Mars GRAM model to find the maximum values of the Ta/Tg ratios.Figure 5-19 illustrates variability of the Ta/Tg with season and latitude.Figure 5-19. Variations of Ta/Tg with season and latitude.Figure 5-19 shows three sets of Ta/Tg values at midnight at the altitude of 6 km for seasons ofsouthern summer (Ls=250), northern summer (Ls=90) and equinox (Ls=0) – as a function of148

latitude. As can be seen, large values of Ta/Tg are confined neither to equatorial regions (highestTa), nor to polar regions (lowest Tg).Figure 5-20 shows variation of the maximum Ta/Tg with altitude for all latitudes for the seasonof Ls=90º, for three dust levels in the atmosphere (0.1, 0.3 and 1) and for the design case withε top /ε bottom =1 (corresponding to design cases 1, 3 and 4 in Table 5-5). These conditions representthe worst case for the buoyant gas mass determination.Figure 5-20. Variation of maximum Ta/Tg with altitude for all latitudes for Ls=90°.The conclusion that follows from this analysis is that the value of Ta/Tg=1.4 is the minimumvalue that provides for non-zero superpressure at night for all density levels and dust conditions.It is interesting to note that while the value of Ta/Tg does vary with altitude, it varies just slightly(within 10-20%), so that a constant value of this parameter is a very good approximation.For a balloon with aluminized top (design #6, results not shown) the minimum required value forTa/Tg is reduced to 1.3, which is not significantly different from 1.4 in this analysis.Surprisingly, the balloon design problem is simplified by this result - for a balloon at any altitude(from 0 to 20 km) the same amount of buoyant gas can be used in the design.We have also determined the extremes of the film temperature for several balloon designs.Figure 5-21 shows the results of this analysis. The lowest temperatures on this plot are over the149

polar caps (lowest surface temperature) and are compared to the CO 2 condensation line. Note,that even balloons with aluminized top will experience frost condensation in the Polar Regions ifthey are floating at the altitudes below about 5 km. In the presence of polar clouds even suchballoons will experience frost condensation at all altitudes.Figure 5-21. Extremes of the DARE balloon film temperature for the 6 designs in Table 5-5.The best DARE balloon design will have high night and low day temperatures of the film. Ascan be seen from Figure 5-21 this is not possible with a single design: increasing night filmtemperature by changing the thermo-optical properties of the portion of the film also increasesday film temperature. This circumstance suggests a new enabling technology: an advancedballoon film coating with changing thermo-optical properties that vary according to day-nightconditions. We have not studied the benefits of such a technology on the DARE balloon design.Using the Mars-GRAM 2001 model we have also calculated "the hottest" cases - when the gastemperature will be the highest and the stress on the envelope due to the superpressure will bethe highest too.Figure 5-22 shows an example of the analysis that estimates the peak hoop stresses on theenvelope as a function of the maximum payloads for DARE balloons at different altitudes andfor different technology horizons. The solid lines show these functional dependencies beforeballast drop, and the dotted lines show these dependencies after a ballast drop of a certain mass(see the legend).150

Figure 5-22. Peak stress and maximum payloads for design #6.The altitudes of the DARE balloon on these curves change from the lowest (0 km) at the upperright end of the curve to the highest (different for balloons at different technology horizons, seebelow) at the lower left end of the curve.The horizontal solid and dashed blue lines mark the ultimate hoop stress and the hoop stress witha 1.5 safety factor of the composite balloon material proposed for a Mars balloon in the MABSstudy. These curves are shown here for reference. In our design work we are going to define thereference advanced balloon envelope material based on this stress analysis.Figure 5-23 shows the altitude change following a ballast release of a certain mass and thecorresponding amount of gas that needs to be vented to return the value of the superpressure tothe value that existed before the drop (superpressure - and hence the hoop stress - increases aftera ballast drop due to the balloon rising to a higher altitude and being in a lower pressureenvironment). The solid curves give the altitude change (left axis), while the dotted lines give thevented gas mass (right axis). The wavy appearance of the vented gas mass curves for Near andFar term technology horizons is an artifact of the method used to estimate the vented gas mass.For the Near term technology horizon the release of 10 kg of ballast (a dropsonde, a surfacelander) results in altitude increase of about 0.9 km according to this analysis. Venting of about0.33 kg of buoyant gas removes the excess of superpressure and reduces envelope stress.151

Figure 5-23. Altitude change and the mass of the vented gas for design #6.The analysis presented in this Section gives us a clearer picture of the balloon behavior during amission and helps in developing conceptual design.5.4.3 Entry, Descent and Inflation (EDI) analysisThe EDI sequence involves the period of the mission from atmosphere insertion until the DAREplatform is operating nominally. It is described in detail in Section 3.6.2.We have developed an EDI model to study possible vertical trajectories of the entry vehicle ofthe DARE platform during deployment. Using inputs such as atmospheric density profile,inflation rate, float altitude, parachute size, etc. the model calculates the position of the craft as afunction of the time from the moment of the main chute deployment.5.4.3.1 Model AssumptionsThe model assumes that the entry vehicle has a similar geometric as the Pathfinder entry capsule.The vehicle has an entry velocity of 7.026 km/s and a flight path angle of –15 deg. In selectingan entry velocity we examined the V infinity values for possible mission opportunities. A nominal152

ange of values for typical mission opportunities is between 3-5 km/s. The higher value of 5 km/sresults in an entry speed of 7.026 km/s assuming an entry interface altitude of 125 km altitude.Flight path angle for Mars entry are usually between –10 and –15°. A conservative flight pathangle of –15° was selected since lower angles will result in EDI completion at higher altitudes.The entry vehicle has an assumed total mass of 340 kg. With these parameters we calculate thedescent trajectory of the craft with HyperPASS, the code developed by GAC for planetary assisttrajectory analysis. It is assumed that the main chute will not be deployed until the dynamicpressure (1/2*ρV 2 ) has reached 700 Pa, which is consistent with successful Mars landingmissions. Using the output from HyperPASS we can determine the point when the entry vehiclehas reached a dynamic pressure of 700 Pa. The altitude and velocity of this point are then used asthe initial conditions in the EDI model.5.4.3.2 Model DescriptionThe model continually sums the forces acting on the system and integrates over a small time stepto find velocity and altitude of the descending system. The trajectory is characterized by fourdistinct phases.5.4.3.2.1 Trajectory Phase IThe first Phase is triggered when the craft reaches a dynamic pressure of 700 Pa, at which timethe forces include the aerodynamic drag force on the entry capsule and the weight of theparachute and entry, inflation and balloon system.5.4.3.2.2 Trajectory Phase IIThe second Phase is triggered when the craft slows to a user specified dynamic pressure. Thispressure defines the capability of the envelope; higher pressures would cause tears and rupturesin the envelope during descent. In the Mars Aerobot Balloon Study (MABS) report a value of 5Pa was given as the dynamic pressure that the envelope could safely sustain. We use the samevalue in our analysis. This value is conservative, since we assume the use of advanced materialfor the parachute that is stronger and lighter then the materials used in the MABS analysis.During Phase II the balloon envelope is deployed from its canister and inflation begins. Inflationrate is adjustable for trade studies. The forces summed by the model during Phase II are the dragof the entry vehicle, drag of the parachute, drag of the semi-inflated envelope, buoyancy of theinflating envelope and the weight of the entry, inflation and balloon system.5.4.3.2.3 Trajectory Phase IIIPhase III of the trajectory is triggered when the sum of the forces created by the envelope drag,envelope buoyancy, and entry vehicle drag is equal to the drag created by the parachute. At thispoint it is assumed that the main chute is cut away and inflation continues. The forces of interestin Phase III are the buoyancy and drag of the envelope, the drag of the entry capsule, and theweight of the entry, inflation and balloon system.5.4.3.2.4 Trajectory Phase IV153

Phase IV is triggered by the completion of the inflation of the envelope. When the inflation iscomplete the remainder of the entry system mass is dropped. This includes the jettisoning of theinflation system and the heat shield and any other system not essential to the DARE mission.The forces of interest in Phase IV are the drag of the envelope, the buoyancy of the envelope,and the weight of the balloon system.5.4.3.3 Modeling ResultsThe model performed well and allowed for parametric analysis of a DARE EDI system. Thevariables studied were: parachute area, inflation time, dynamic pressure at inflation initiation,float altitude, and density profile. All variables were studied individually in order to understandthe trades involved. Each variable was manipulated without consideration of mass, for instancethe area of the parachute was altered while the parachute mass was constant. This is a crudeapproach but allows us to make important characterizations. The next four sections detail a tradeperformed using the “averaged” density profile. In each section we discuss the effect of varying aparticular parameter while the other three remain constant. The four parameters studied are:parachute area, inflation time, dynamic pressure at inflation initiation, and float altitude. Eachsection contains a chart showing the effect of varying the parameter of interest on the systemminimum altitude. In each chart is a small yellow text-box that displays the values of theparameters that were not changed. The values shown in the yellow text-box are initial designvalues. They represent the first design point.5.4.3.3.1 Parachute AreaThe parachute directly affects the altitude at which inflation begins. A larger chute will slow theDARE system faster and allow the envelope to be inflated at a higher altitude. Figure 5-24 belowshows the relationship between minimum system altitude and parachute area. If the system isequipped with a chute that is less than approximately 300 m 2 it will crash to the ground beforefully inflating. The plot on Figure 5-24 levels off at the altitude of 10 km since this altitude waschosen as the floating altitude in this study.1200010000Z min (m)8000600040002000500 s Inflation Time10 km Float Altitude6 Pa Pressure Inflate00 200 400 600 800 1000A chute (m 2 )Figure 5-24. Variation in system minimum altitude with parachute size154

5.4.3.3.2 Inflation TimeThe inflation time is the duration of the envelope inflation period. Again we are assuming thatthe inflation hardware has constant mass for all inflation rates. On Figure 5-25 below we see thatwe can vary the minimum system altitude by adjusting the inflation rate. For a system with ashort inflation period the envelope is fully buoyant much sooner which prevents the system fromdipping low into the atmosphere.12000100008000Z min (m)600040002000400 m 2 Chute Area10 km Float Altitude6 Pa Pressure Inflate00 200 400 600 800T fill (s)Figure 5-25. Variation in system minimum altitude with inflation timeAs the inflation time is shortened a larger and more complex inflation system is necessary. Thestored gas must be heated to atmospheric temperature as it inflates the envelope and thisrepresents a significant challenge. A system that does not require a fast inflation rate will besmaller and possibly less complicated. For the three trades in this analysis the inflation time is setat 500 s. This is merely a guess and is not based on any study or calculation.5.4.3.3.3 Inflation Initiation Dynamic PressureThe inflation initiation pressure is the trigger for the inflation initiation. If the balloons envelopeis capable of sustaining a higher dynamic pressure, the envelope can be unfolded and theinflation started earlier, and full buoyancy can be achieved much sooner. If the dynamic inflationpressure is too low (below about 5 Pa) the system will not ever begin inflating before crashing tothe ground. The sharp change in slope at around 7 Pa inflation pressure is probably an artifact ofthe low resolution of the trade. Dynamic pressure requirements were varied in relatively largesteps. A higher fidelity analysis would result in a softer curve but the general shape wouldremain the same.155

1200010000Z min (m)800060004000500 s Inflation Time10 km Float Altitude400 m 2 Chute Area200000 5 10 15 20 25P Inflate (Pa)Figure 5-26. Variation in system minimum altitude with inflation initiation dynamic pressureFigure 5-26 above shows the minimum system altitude as a function of the dynamic pressure thatinitiates inflation.5.4.3.3.4 Float AltitudeIf the float altitude is increased without varying the system mass we see that the system will notdescend into the atmosphere as far during EDI. This is because we have assumed that theinflation period is constant. A larger (increased floating altitude) balloon will inflate rapidly andhave an increased amount of drag as it descends through the atmosphere. Figure 5-27 belowshows the relationship between the minimum system altitude and the final floating altitude of theDARE system. In Figure 5-27 we see a sharp knee in the curve at a float altitude of 4000 m.Below this altitude the minimum altitude is equal to the system float altitude – the floatingaltitude is the minimum altitude reached by the entry probe. Above this altitude the floatingaltitude is larger than the minimum altitude – after reaching the minimum altitude the entryprobe rises to the floating altitude. The altitude where the floating altitude is equal to theminimum altitude is determined by the model parameters: the rate of the inflation, dynamicpressure to initiate the inflation and the chute area. For the values of the parameters used togenerate the plot on Figure 5-27 this altitude is about 4 km.156

Zmin (m)80007000600050004000300020001000500 s Inflation Time6 Pa Pressure Inflate400 m 2 Chute Area00 2500 5000 7500 10000 12500 15000Z float (m)Figure 5-27. Variation in system minimum altitude with system float altitude5.4.3.3.5 Atmospheric Density ProfileThe density of the Martian atmosphere varies substantially geographically and seasonally. Tostudy the effects of different atmospheric conditions on the EDI trajectories we have chosenthree density profiles corresponding to three hypothetical atmospheres. The profiles correspondto a nominal (Mars COSPAR atmosphere model 39 ) atmosphere, a dense cold atmosphere and ararified warm atmosphere. In choosing the atmospheres for the EDI trade study we have toconsider two properties of the atmosphere – its total mass and its extent. The atmosphere with asmall total mass will not decelerate the space probe as efficiently, as the more massiveatmosphere. Similarly, the atmosphere that does not extent very far above the surface will bepenetrated by a space probe quicker, than the more extensive atmosphere. For the purposes ofchoosing the density profiles for EDI analysis we will assume that the extent of atmosphere isdefined by the atmospheric scale height (the measure that defines “how quickly” the densitydecreases with height). In our analysis of the EDI trajectories we are primarily interested infinding out if any atmospheric conditions or/and geographic locations needs to be avoided forEDI.We do not know a priori what parameter has a larger effect on the EDI trajectories – the mass ofthe atmosphere or its extent. For this reason we have to choose several representative profilesand generate the EDI trajectories with these profiles. Table 5-6 gives the parameters of theatmospheres that were considered for the EDI analysis. These atmospheres span a full range ofseasons and dust loadings observed at Mars. For simplicity, the density profiles are defined by39 David. E. Pitts et al., "The Mars Atmosphere: Observations and Model Profiles for Mars Missions", NASA Johnson SpaceCenter report JSC-24455, 1990.157

the 2 parameters - ρ(0) – density at the 0 height level (Mars MOLA areoid) and, H – scale heightin km. The density profile is given by: ρ(z)= ρ(0)*exp(-z/H), where z is in km.Table 5-6. Various atmospheres considered for EDI trade studies. Atmospheres 1 and 2 representlimiting cases.Case ρ(0), kg/m 3 H, km Atmospheric Commentscolumnmass, kg/m 21 0.0169 9.96 168 L s =270, τ=0.3, Northern Lowlands (UtopiaPlanitia), topography = -4.9 km.Deepest atmosphere.2 0.0113 11.7 132 L s =270, τ=3 (dust storm), Hellas Planitia,topography = -6.8 km.Entering in the southern hemisphere during duststorm over the deepest depression.3 0.0138 9.94 137 L s =150, τ=0.45 (normal dust cycle), HellasPlanitia, topography = -6.8 km.Atmospheric entry over the deepest depression inthe Southern hemisphere during lowest planetarysurface pressure condition.4 0.012 11.1 133 L s =150, τ=0.45 (normal dust cycle), UtopiaPlanitia, topography = -4.9 km.Entering in the northern hemisphere duringlowest planetary surface pressure condition.5 0.0117 11.8 138 L s =150, τ=0.45 (normal dust cycle), equatorLon=0, topography = -1.2 km.Entering over equator during lowest planetarysurface pressure condition.6 0.0153 9.4 143 L s =150, τ=0.45 (normal dust cycle), southernhighlands, lat=-40, lon=0, topography = 1 km.Entering over southern highlands during lowestplanetary surface pressure condition.7 0.0137 11.4 156 L s =90, τ=0.3 (normal dust cycle), UtopiaPlanitia, topography = -4.9 km.Entering over northern lowlands during summer8 0.0151 8.9 134 L s =270, τ=3, Northern Lowlands (UtopiaPlanitia), topography = -4.9 km.Same as 1, but in storm9 0.0141 10.5 148 L s =0, τ=0.65 (normal dust cycle), UtopiaPlanitia, topography = -4.9 km.Entering over northern lowlands during equinox10 0.0139 10.9 151 L s =0, τ=0.65 (normal dust cycle), HellasPlanitia, topography = -6.8 km.Entering over southern highlands depression inequinox158

Atmospheres 1 (cold dense) and 2 (warm rarified) were chosen for the EDI trade studies.Atmosphere 1 is dense, but “compact” – it has the smaller scale height H. Atmosphere 2 is warm(dust storm conditions) and is more extensive (large scale height H)In the chart below we examine EDI trajectories for the three different atmosphere profiles. Forthis study we again called upon the HyperPASS code to provide initial conditions for the EDImodel for each density profile. Using the velocity and altitude, during the entry vehicles descent,corresponding to a dynamic pressure of 700 Pa we ran the EDI model with the three differentdensity profiles. Figure 5-28 below shows that for the simply assumed system the minimumaltitude is very feasible. The system as described provides clearance for Martian relief in excessof 5000 m above the reference level (MOLA areoid). This design iteration gives us a greatamount of flexibility and accounts for unknown atmospheric conditions.Figure 5-28. Minimum altitude reached during EDI for atmospheric conditonsThis analysis also shows that total atmospheric density is the most important atmosphericparameter that determines the lowest altitude of the EDI trajectory. Hence, the “warm rare”atmospheric profile used in the EDI analysis represents the worst case for the EDI.5.4.4 Sizing of the Imaging Suite for the Conceptual PayloadThe payload for the chosen conceptual mission scenario (Surface Networks Emplacement)includes visible and infrared cameras for observations of the surface and the atmosphere.Simulations of the BGS performance and of the balloon trajectories at Mars indicate thatemplacement of a network of several surface stations spread across the planet may take a159

substantial amount of time (several months). The time that it takes to travel to the next surfacestation emplacement location can be spent by making high-resolution observations of the surfaceand of the atmosphere with optical and infrared instruments. In addition, the cameras would alsobe required for autonomous navigation and guidance.We have decided the imagery payload would include one high-resolution narrow angle camera(NAC), one lower-resolution wide-angle camera (WAC) and one thermal emission spectrometercamera (TES). NAC and TES would be used for observations of the targets on the surface and ofthe atmosphere, while WAC would be used to provide context for high-resolution imaging andfor navigation.In this analysis of the camera sizing for the DARE concept we have relied (partially) on theprevious analysis of the aerial imaging systems by Ken Klaasen of JPL 40 .Imaging system design considerations involve many parameters, such as spatial resolutionrequirements, areal coverage requirements, image quality (e.g., signal-to-noise ratio, SNR)requirements, spectral resolution requirements, flight characteristics of the aircraft, mass andpower limitations and data storage and communication capability. Here we will limit the analysisto consideration of spatial resolution, image quality and flight characteristics. Additionalparameters may be considered as the concept develops further.Higher spatial resolution and image quality requires larger optics, slower aerobot ground speedsand lower altitudes, while the mission cost constrains the total mass of the system. BGSperformance, in general, improves with the increasing height of the balloon above the surface.In a typical aerial imaging system optics is the major mass component. The detectors are usuallythe line or area arrays of 10 µm pixels. Typical performance requirements for an aerial imagingsystem are: less than 1 pixel of smear; diffraction-limited optical performance - f/8 or fasteroptics for 10-µm pixels; SNR >50.Figure 5-29 shows the analysis of the maximum allowable ground speed from the Klaasenpresentation 40 . Higher ground speeds require faster optics (smaller f-numbers) for a given groundpixel size. For a 5-cm ground resolution and a diffraction-limited f/8 telescope the maximumallowable ground speed is about 100 m/s. Analysis of the Martian winds for the DARE balloonaltitudes (up to about 12 km above the surface) indicate that the ground speed will rarely exceed100 m/s – and only over a limited number of locations (see Figure 5-30). Hence, an f/8 telescopeaboard the DARE platform will be adequate to provide diffraction-limited centimeter-levelspatial resolution.40 “Aerial Imaging System Design Considerations”, Kenneth P. Klaasen, Jet Propulsion Laboratory, BEES 2000 Workshop,December 4-6, 2000.160

Figure 5-29. Maximum allowable ground speed as a function of the telescope f-number andspatial resolution (Klaasen, 2000).Figure 5-30. Statistics of the wind speed magnitudes at about 12 km altitude at Mars for L s =270and low dust conditions.The next step in sizing the aerial optical system for the DARE platform is to define the systemmass increase with the increasing aperture of the telescope. Klaasen suggests that the mass of theoptical system increases as the 3 rd power of the aperture in meters (Klaasen, 2000). We havedeveloped our own scaling law, which is different from that used by Klaasen.161

Table 5-7. Spacecraft, balloon and commercial telescope characteristics.162

Figure 5-31. Dependence of the optical system mass on the apertureWe have analyzed a number of optical systems that have been flown on spacecraft (Cassini,Galileo, MOC, HiRise, etc.) and atmospheric balloons (BLAST, MSAMI-95). We have alsoconsidered smaller commercially available telescopes (Orion). The data used in the analysis arepresented in Table 5-7.The data from Table 5-7 are plotted on Figure 5-31 in log-log coordinates. Two sets of data areshown – those for the whole telescope (blue), and those for the optical tube only (red, from thedata on commercial telescopes). The data span almost three orders of magnitude of telescopemasses and more than an order of magnitude of telescope apertures. As can be seen, the completetelescopes are heavier as expected, because in addition to the optical tube they include thermalcontrol, pointing and other subsystems. We fit both sets of data with a linear function.Interestingly, both sets of data can be fit with the linear function of approximately the sameslope.The scaling low derived from the fit to the data for complete telescopes gives the followingdependence of the optical imaging system mass M (in kg) on the systems aperture A (in m):1.7M ≈ 316AThis dependence predicts lighter optical systems then the 3 rd power dependence suggested byKlaasen.163

Figure 5-32. Optical system aperture in meters as a function of desired surface resolution(diffraction limited).Figure 5-33. Optical system mass in kg versus desired surface spatial resolution (diffractionlimited).We can calculate the sizes and the masses of the optical systems for various altitudes and spatialresolutions using the derived scaling law. Figure 5-32 and Figure 5-33 show the calculateddependence of the diffraction limited system aperture and system mass on the desired surfaceresolution for several altitudes above the target. The diffraction-limited aperture was calculatedassuming 10-µm pixels for the detector:164

A = 1.22⋅ λh/ r ,where A is aperture in meters, λ is the wavelength (assumed 1 µm for visible cameras), h isheight, r is resolution.As can be seen from the figures, for an optical system at 10 km altitude looking at nadir target, toachieve spatial resolution in the range of 1 to 10 cm per pixels would require a camera with anaperture of 10s to 100 centimeters that would weigh from tens to several hundreds of kilograms.Reducing the distance to target almost by a half to 6 km does not reduce the apertures andmasses very much. The distance to target has to be reduced substantially – to 1 km – to bring theoptical system mass down to several kilograms.Given this analysis we have decided to move the optical instruments from the gondola to theBGS, where they can be housed in the boom. The nominal altitude of the BGS is 1 km above thesurface. The winch at the gondola would reel-up or down the BGS wing to consistently achievethis height. The power for the camera will come from the solar panel and the batteries on andinside the boom of the BGS wing. Pointing to the target would be achieved with a small motor(estimated mass – 1 kg). Attitude sensors would be housed together with the cameras and enableprecise pointing of the cameras. All three cameras will be co-located on the boom and pointed asa single unit.For a 3 cm/pixel resolution and 1 km distance to the target the visible camera aperture isestimated at 4.1 cm and the camera mass is estimated at 1.1 kg. The resultant camera is verysimilar to the cameras conceptualized for the Mars Aerial Platform (MAP) 3 and Mars AerobotSystem (MABS) 2 efforts. MAP Narrow Angle Camera (NAC) has similar aperture and mass (3cm and 0.75 kg respectively, see Table 5-7). We assume that all three cameras are similar to theMAP cameras (NAC, WAC and TES). The parameters of these cameras are multiplied by afactor of 1.5 to estimate the parameters for the DARE optical suit of instruments, to account forslightly larger size of the DARE cameras. The estimated parameters for the cameras aresummarized in Table 5-8.Table 5-8. Parameters of the DARE optical imaging systemsNAC WAC TESAperture, m 0.041 0.041 0.041Mass, kg 1.1 1.1 3.3Spatial resolution, m/pixel 0.03 3 3Length, m 0.2 0.2 0.25Power, W 6 6 7Duty cycle, sol fraction 0.25 0.25 0.5165

Moving the cameras from the gondola to the BGS significantly reduced the required systemsmass (by about 40 kg or 30% of the floating mass).5.4.5 Sizing the Solar Panels and the Batteries for DARE Conceptual PayloadWe size the solar panel and the batteries by matching the required power and energy levels,estimated in Section 3.6.5. We size solar panels at the gondola and at the BGS separately, toavoid transmitting power and energy over large distances (several km) from the gondola to theBGS. Negative factors that accompany transmitting power over large distances include largeresistance losses, increased tether mass, and unacceptable tether heating due to electricalresistance.Several factors affect the power generating capabilities of the solar panels at Mars. They includeaccumulation of the atmospheric dust on the solar panels, attenuation of the solar flux in thedusty atmosphere, shadowing of the solar panel at the gondola by the balloon, and the elevationangle of the sun. These issues are discussed in more detail in the sections below.5.4.5.1 Atmospheric dust accumulation on solar panelsAccumulation of the atmospheric dust on the solar panels is a well-known issue for Mars. Thelifetime of the Mars Exploration Rovers (MER) was originally expected to be limited by theaccumulation of dust on the solar panel and the accompanied reduction in power production.MERs were lucky in that their solar panels were apparently “cleaned” by passing dust devils,extending their life-time beyond the initial estimates. Winds of the order of 10 m/s are needed atMars to lift dust particles (0.1 – 10 µm). For a balloon at a 10 km altitude similar wind velocitiesare expected for dust lifting. In our simulations we see relative winds at the balloon altitude thatare below 10 m/s. These winds are not sufficient to lift dust from the solar panels of the DAREplatforms. While dust devils can extend from the surface to the heights of several kilometers atMars, it seems highly improbable that a DARE platform could fly through a dust devil top at theheight of 10 km.According to Arvidson et al. 41 both MER crafts experienced a solar array output degradation of0.2 % per sol. While the dust concentration at 10 km is smaller than at the surface, the density ofthe atmosphere is also smaller, so that the dust grains fall faster. Hence, the degradation ratecould be the same at 10 km altitude as at the ground level. Estimated degradation of the solarpanel power output is illustrated in Figure 5-34. It can be seen that, in one year, the solar arrayoutput is reduced to 50% of the beginning of life (BOL) value. For a DARE mission lastingseveral Martian years it would be impractical to design the system to have a solar array largeenough to provide adequate power despite the accumulation of dust. Therefore the DARE systemmust rely on an active dust removal system.41 “Localization and Physical Properties Experiments Conducted by Spirit at Gusev Crater.” Science Vol. 305 6 August 2004:821-824166

Solar Array Output Degredation Due to Dust Accumulation100%Solar Array Output (%)80%60%40%20%0%0 100 200 300 400 500 600Mission Time (sol)Figure 5-34. Degradation of the solar panel power output due to accumulation of dust.The active removal of dust could take on several forms 42 . It is thought that the Martian dustwould adhere so well to an array that a mechanical wiper system would not be effective withouta lubricant. The availability of a liquid suitable for use in the Martian atmosphere is questionableand the liquid itself would be a consumable. Another option is the use of an electromechanicalsystem in which the dust is loosed from the array by an electrostatic or ultrasonic method andthen cleared away with a mechanical wiper. Compressed air could also be used to blow the dustfrom the array. It is conceivable that a small amount of air could be compressed over a longperiod of time and released in bursts to clear away dust. While local air velocities could easilyexceed 10 m/s it would most likely be difficult to store enough air to provide the necessary airvelocity over the entire array from one nozzle. Therefore a manifold distributing the air toseveral nozzles spread across the array may be required.5.4.5.2 Solar flux attenuation by atmospheric dustThe ever-present atmospheric dust in the Martian atmosphere can potentially significantly reduceavailable solar power. Using Mars-GRAM we estimated attenuation of the solar flux in theatmosphere for several scenarios with differing dust content in the atmosphere. We have lookedat a range of solar zenith angles (angle between the directions towards the sun and towards thezenith) and at atmospheric dust opacities of 0.3, 0.7 and 3. The first two values correspond to theminimum and maximum dust opacities observed by Viking landers over a Martian year without a42 G. A. Landis, “Mars Dust Removal Technology”, AIAA Journal of Propulsion and Power Vol. 14, No. 1, 126-128, Jan. 1998.Paper IECEC 97-97340, Intersociety Energy Conversion Engineering Conference, July 27-August 1, 1997.http://powerweb.grc.nasa.gov/pvsee/publications/mars/removal.html167

dust storm, while the last value corresponds to the very high dust opacities observed during adust storm.Figure 5-35 shows the results of this analysis. Figure 5-35 shows the transmittance (the fractionthat gets transmitted) of the solar power at the surface of Mars for the three values of the dustopacity tau, and for a range of solar zenith angles from 0 (the sun is directly overhead, in zenith)to 90 degrees (the sun is on the horizon). Solar flux at altitude of 10 km is not readily availablefrom the Mars-GRAM, so we are using the values for the transmittance at the surface as a proxy.Transmittance at 10 km will be 10 to 30% smaller than at the surface.Figure 5-35. Transmittance of the solar power through Martian dusty atmosphere for varioussolar zenith angles.This analysis shows that for low atmospheric dustiness (τ=0.3-0.7) atmospheric attenuationbecomes important (reduces by a factor of 2) for sun zenith angles in excess of 75-80º. For thishigh zenith angle (low elevation angle) the solar power generation that is proportional to thecosine of the solar zenith angle is already significantly smaller than for lower zenith angles.Hence the effects of atmospheric attenuation are secondary in this case.During a dust storm (τ=3) the incoming solar flux is reduced by a factor of 2 even when the sunis in zenith (optimal power generation configuration). For higher zenith angles the solar powertransmittance is reduced even further. Operating in dust storm conditions will prevent the use ofimaging equipments, since the surface features will be indiscernible and thermal contrastsminimal. Hence the power draw of the DARE payload will be significantly smaller in a duststorm. In the following analysis (Section 5.4.5.4) we assume a loss factor of 20% in our analysisof power generation, corresponding to low dust conditions. In the event of a dust storm the168

power requirements for the payload would be reduced (since the cameras will not be able tooperate in the opaque atmosphere) and even reduced power generating capabilities of the solararray would be sufficient to power the payload.5.4.5.3 Solar panel shadowing by balloonA small examination of the effect of the length of the tether connecting the gondola to theballoon on array shadowing was conducted. The results shown in Figure 5-36 below are for a 17m diameter balloon. Based on the shadowing study it was decided that the length of the tetherwould be 50 m. With this tether length the gondola solar panel will be shadowed by the balloononly when the sun is closer than about 15º to the zenith.100%Tether Length Study% Illuminated (%)80%60%40%20%10501001502000%0 20 40 60 80 100Angle DARE_Sun (deg)Figure 5-36. Illumination of the gondola solar panel as a function of the balloon-gondola tetherlength.5.4.5.4 Energy analysis from sample simulated mission trajectoryWe developed a solar power model that takes trajectory data from a DARE trajectory simulationand estimates the amount of solar energy that can be collected and stored in the fuel cell perMartian solar day (sol). We used a simulated mission trajectory (Figure 3-28) for this analysis.The solar power model assumes cell efficiency of 30%. The model assumes the array will shutoff if it is more than 50% shadowed in order to prevent array damage. It is assumed that the169

charge-discharge efficiency of the fuel cell is 80%. This inefficiency is applied to the energyused to charge the storage system. We assume that the solar array is always clean (settledatmospheric dust is being removed by means noted in Section 5.4.5.1). Atmospheric attenuationis assumed to be constant for simplicity and equal to 20%. In the event of a global dust storm thescience package will function at a minimal level and the BGS will not be reeled-up for optimalcontrol. Hence the global dust storm conditions do not drive the power generation and energystorage design. Therefore it was reasoned that the system would be designed to accommodateoperation in the atmosphere with 80% transmittance (moderately dusty conditions, see Figure5-35).Solar flux at the top of the atmosphere (TOA) in the model is varied according to the changes ofthe Mars orbit. The flux is calculated with the following equation using the orbital radius inastronomical units:W 1Flux ≈ 1365 • AU2 2m RFor Mars the orbital radius R varies from 1.38 to 1.67 astronomical units over one Martian year.The input parameters for the power and energy analysis model are the date, solar zenith angle,and Martian orbit radius. The DARE-solar zenith angle is defined as the angle between the zenith(line between a location on planet surface and a point directly above it) of the location of theDARE system and line of sight to the sun. Therefore the zenith defines the angle between theperpendicular to the gondola solar array and the sun. An example of the DARE-solar zenithangle is shown in the plot below. A DARE-solar zenith angle of 0 degrees corresponds to a local(DARE) noon. An angle of 90 degrees describes the DARE sunset, while 180 degrees representsthe instance when the DARE is on the dark side of the planet directly opposite the sun.The model includes energy storage assumptions, reel-up modeling, and reeled BGS shadowing.The energy storage system is assumed to be a regenerative fuel cell with 80% charge/dischargeefficiency. This is incorporated as an 11% loss on power going into and out of the energy storagesystem. The reel-up portion of the trajectory required assumptions about the speed and mass ofthe reeled system. It was assumed that for optimum control the BGS should be able to be reeledthe entire length of the tether in 1 hour. The tether length can vary in order to avoid terrain;therefore the reeling velocity is variable. The system BGS laden with science instruments isassumed to have a mass of 13 kg. As the BGS is expected to continue to operate when reeled upit is assumed that balloon shadowing will affect the energy production. The assumptions thatdrive the shadowing algorithm for the gondola array are used for the BGS array. This is notcorrect but is a close approximation.Figure 5-37 shows the DARE platform’s solar array power capability along the mission’strajectory as a function of mission time. Solar array power capability is defined as a percentageof the maximum power level available from the array. The array power capability is 100% forthe sun zenith angle of 0º.170

Figure 5-37. DARE platform solar array power capability along the mission trajectory as afunction of mission time.Figure 5-37 also shows that the DARE platform can travel fast enough to move out of phase withthe planet diurnal cycle. For example, during the first 5 days of the mission an observer at theDARE platform would have observed 7 local noons and sunrises. This is because the platform ismoving fast East at the beginning of the mission (see Section 3.6.8), shortening the days – timeintervals between consecutive noons or midnights – as observed from the platform.For the purposes of the power and energy budget analyses it is useful to convert to “DAREdays”, which are defined as time intervals between consecutive midnights. Each DARE day has adifferent number of hours in it, but we are primarily concerned with daily, not hourly variationsin our analysis.171

Figure 5-38. DARE platform solar array power capability along the mission trajectory as afunction of “DARE time”.Figure 5-38 shows the data shown on Figure 5-37, but plotted as a function of “DARE time”. Inthis representation, maximum power capability coincides with the local noon.Figure 5-39. DARE day duration and the latitude of the platform.Figure 5-39 illustrate the changes in the hourly duration of a “DARE day”. The blue line showshours in a DARE day and relates to the left axis (hours), while the red curve shows thegeographic latitude of the platform and relates to the right axis. At the start of the mission (days0-20) the platform moves in the easterly direction with the strong zonal flow and this shortensthe day as would have been seen by an observer at the DARE platform. As the platform drifts172

across the equator (days 20-60, see also Section 3.6.8), the winds become weaker and theDARE’s diurnal cycle becomes equal to the Martian cycle. As the DARE platform floats in theeasterly zonal flow at mid-latitudes in the Southern hemisphere (days 60-80) the DARE dayduration shortens again.Figure 5-40 shows the daily energy consumption of the BGS reel-up mechanism along with thelatitude of the DARE platform. The variation in daily reel-up energy required coupled with thevariation in daily energy generated complicates the analysis. Energy consumption by the reelingof the BGS affects only the energy balance of the gondola. This energy is added to the energyrequired by the payload to arrive at the energy budget of the gondola payload.Figure 5-40. Energy consumption due to the BGS reeling.To estimate the sizes of the gondola and BGS solar arrays and batteries we vary the areas of thesolar arrays and the capacities of the batteries until we match the daily power and energy budgetrequirements for the gondola and the BGS. The batteries power up the BGS and gondolapayloads at night and during days when power generated by solar arrays is below the requiredlevel. Figure 5-41 compares the average energy load of the gondola and daily generated energyof the 0.41 m 2 gondola solar array. It also shows the latitude of the DARE platform and thelatitudinal difference between the latitude of the sun and the DARE platform (yellow curvemarked as DARE Solar Latitude). The decrease in the generated energy between “DARE days”20 and 50 is due to daily shadowing of the gondola solar array by the balloon. During this timeperiod the DARE platform crosses the equator. The latitudes of the sun and of the DAREplatform are almost the same during this time, meaning that the sun is almost at zenith at noonwhen observed from the DARE platform.Figure 5-42 compares the average energy load of the BGS and daily generated energy of the 0.79m 2 BGS solar array. The energy requirement for the BGS is higher than for the gondola payloadbecause of the higher power consumption of the cameras at the BGS. BGS array energy173

generation daily variability is slightly different from that of the gondola array, as it is onlyshadowed when completely reeled up to the gondola, not when the sun is high in the sky.Figure 5-41. Daily energy generated by 0.41 m 2 gondola solar array.Figure 5-42. Daily energy generated by 0.79 m 2 BGS solar array.Figure 5-43 compares gondola solar array energy generation, gondola energy load and the stateof charge of the gondola battery with the maximum energy load of 700 W-hr. Such a batterywould weigh 1.75 kg according to the Near term fuel cell development projection (see Table174

5-3). The battery of this size is sufficient to provide enough power for the gondola payload forthe duration of the ESN mission. Solar array generated power is sufficient to keep the batteryfully charged most of the time. Twice during the mission (day 5 and day 35) the batterydischarges to the depth of about 50%. These moments correspond to instances of higher energyload due to frequent reel-ups (compare to Figure 5-40).Figure 5-43. Gondola battery charge (maximum energy 700 W-hr)The battery discharges completely by the end of the “nominal” mission due to decreased powergeneration at higher latitudes in the fall hemisphere and increased energy load due to BGS reelup.The levels of the battery discharge can be controlled by limiting BGS reel-ups, but this wouldhave negative effects on trajectory control capabilities.In this analysis the DARE platform is designed to generate enough power and energy to supportits payload until it is removed by about 60º of latitude from the latitude of the sun (see DARESolar Latitude curve on Figure 5-41 and Figure 5-42). Hence, the current design will supportDARE operations at higher latitudes at favorable seasons – up to latitude of 85º at the peak of asummer (solar latitude of 25º). Operations at the poles would require increasing the sizes of thesolar panel arrays and battery capacities.Figure 5-44 compares BGS solar array energy generation, BGS energy load and the state ofcharge of the BGS battery with the maximum energy load of 200 W-hr. The mass of the batteryis estimated at 0.5 kg for the Near-term technology horizon. The BGS energy load does not varydaily as the gondola load does because it is not affected by the BGS reel-up. The battery is fullyrecharged every day of the mission, except at the very beginning (days 0 to 8). The discharge ofthe battery down to the 50% level at the beginning of the mission is due to decreased energyproduction by the BGS solar array at high latitudes.175

Figure 5-44. BGS battery charge (maximum energy 200 W-hr)Figure 5-43 and Figure 5-44 show that energy generation is closely correlated to the duration of a“DARE day”. The longer the “DARE day” (as measured in hours), the higher the energygeneration level. We can conclude that purely for energy generation purposes it would bebeneficial to move with the westerly flow (not easterly, as in the above analysis). The duration ofthe “DARE day” would be longer than 25 hours and the energy generation levels would behigher. Westerly flows in the lower atmosphere of Mars develop in mid-latitudes of bothhemispheres during summer seasons (Ls=90º in the Northern hemisphere and Ls=270º in theSouthern hemisphere). However these flows are much weaker than the easterly flow duringspring and winter seasons. Further analysis is needed to determine energy generation benefits fora platform moving with a westerly flow at Mars.Table 5-9 summarizes the results of the solar arrays and batteries sizing effort.Table 5-9. Solar arrays and batteries sizing results.Solar ArrayBatterySize Mass Capacity MassGondola 0.4 m 2 0.8 kg 700 W-hr 1.8 kgBGS 0.8 m 2 1.6 kg 200 W-hr 0.5 kg5.4.6 Analysis of Effects of the Vertical Wind Gusts on the DARE Platform PerformanceGusts of the vertical winds can potentially drive the balloon into deeper layers of the atmosphere,where it could lose the superpressure and collapse, or into higher atmosphere, where theballoon’s envelope would experience higher stress due to higher internal super-pressure and mayrupture. We have looked at the effects of vertical winds on DARE platform to quantify theprobability of these negative outcomes.176

To estimate the depth to which the balloon would descend under a gust of downward wind weequate the drag force on the balloon to the increased buoyancy:F drag =F buoyancy or20.5ρ w A = ( M − ρ Va b b a b)gwhere ρ a is the atmospheric density, w – is the vertical velocity, A b – cross-sectional area of theballoon, V b – balloons volume, g – gravity.For w=1 m/s and substituting the values for the nominal DARE balloon: R b =17.24 m, A b =933m 2 , V b =21462 m 3 , M b =140 kg we get for the atmospheric density at the new floating altitude:ρρaa _ eqm/s.= 1.005, where ρ a_eq is the atmospheric density in equilibrium, 0.0066 kg/m 3 , with w=0The corresponding change in balloon altitude (for scale height of 10 km) is about 60 m. Thecorresponding change in the ambient pressure is about 1 Pa (from 265 Pa). The lowestsuperpressure estimated for the current DARE design is 10 Pa (nightly superpressure). Thus,vertical winds of 1 m/s will have a very small effect on DARE balloon performance.It is interesting to look at the level of the winds that would drive the balloon to an underpressurizedstate, at least at night. To get a 10 Pa change in the superpressure an altitude changeof about 500 m is needed (at 10 km altitude), corresponding to vertical wind w=3 m/s.According to mesoscale model simulations of Martian winds, these are high values for verticalwinds that can be expected only near the ground (lowest 1 kilometer) and on the slopes of themountains. At the floating altitude of the balloon (10 km) vertical winds not exceeding 1 m/s areexpected.Hence, we conclude that vertical wind gusts will not affect the performance of the DAREplatforms at Mars at floating altitudes several kilometers above the surface.5.4.7 Estimated Rates of Buoyant Gas Loss for DARE PlatformsWe have estimated buoyant gas leak rates for DARE balloons based on the data for balloon gasleaks from the ELBBO/Holzworth Campaign (http://lheawww.gsfc.nasa.gov/docs/balloon/workshop96/superpressure_flights.html). During the campaign several superpressure balloonswere flown for many days (~100 days), which enabled determining the leak rates. The same datawere used previously in estimating leak rates for Mars balloons (e.g. MABS). The slow rates ofgas loss observed in the ELBBO campaign lead to the conclusion that the leakage took placethrough the envelope seams. The loss rate, hence, depends on the length of the seams of theballoons envelope.Figure 5-45 summarizes the results of the ELBBO campaign. The gas loss data – in units ofkg/day – are shown as a function of the superpressure dP. We suggest that to use this data for177

Mars balloons the data need to be scaled to appropriate values of the dP. Without a detail modelof the gas leakage through seams, we assume a linear relationship between the gas loss and dP(see Figure 5-45).Figure 5-45. ELBBO leak rates summary and dP dependence.Several cases were considered in the gas loss analysis. The results of the analysis of the gas lossrates for the DARE Mars balloon are given in Table 5-10.In Case 1 we took the maximum loss rate from the ELBBO data and rescaled it to the size of theDARE Mars balloon (R=17.3 m), assuming a gore width of 2.5 m, consistent with the gore widthof the existing NASA balloons. The resultant gas loss was 25 g/day. At this rate in 12 days theballoon will not be able to maintain its inflation during night (the superpressure at night would be0 Pa). The required amount of gas to replenish an annual loss would be 17.2 kg and a 5-yearmission would require replenishment of 86 kg of gas. These numbers are given here just forreference, as Case 1 is not considered to be realistic, since it does not take into account thedependence of the loss rate on the balloon’s superpressure and other factors.Case 2 does consider the dependence of the gas loss rate on the superpressure. For the typicalsuperpressure of 250 Pa for DARE Mars balloon, the gas loss rate is about 9 g/day. At this lossrate the nightly superpressure would be lost in 33 days, and 1 and 5-year missions would require6 and 32 kg of buoyant gas to replenish, respectively. The nominal design of the DARE Marsballoon has 8.3 kg of hydrogen as buoyant gas. Carrying a gas replenishment system wouldsignificantly affect the mass budget in Case 2.178

Case#Table 5-10. Gas loss estimates for DARE Mars balloon for different assumptionsAssumptionsGaslossrate,g/dayTime toloose nightlydP=10 Pa,dayTotal gas mass toreplenish for 1-year mission, kg1 Max ELBBO 25 12 17.2 862 ELBBO scaled to lower 9.2 33 6.3 31.5Mars dP=250 Pa3 2 + assuming diurnalvariability of gas loss as afunction of dP4.6 66 3.2 164 1 + assuming factor of 10reduction in gas leak ratesdue to manufacturingimprovements5 2 + assuming factor of 10reduction in gas leak ratesdue to manufacturingimprovements6 3 + assuming factor of 10reduction in gas leak ratesdue to manufacturingimprovements2.5 120 1.7 90.9 330 0.6 3.20.5 660 0.3 1.6Total gas massto replenish for5-year mission,kgHowever, the maximum superpressure is achieved only once a day, before the buyant gas startsto cool as the sun sets. The lowest superpressure is seen at night (dP=10 Pa). To take thesefactors in to account, Case 3 assumes that the daily averaged gas loss is a factor of 2 smaller thanin Case 2. Case 3 represents the most realistic assessment of the gas leakage rates for Martianballoons at the current technology. Case 3 indicates that significant amounts of replenishmentgas would be need for a 1 and 5–year missions developed with current technology.Case 4 assumes that improvements in manufacturing will bring a factor of 10 reduction in the gasleakage rates. This would help to bring the amount of the replenishments gas to moremanageable levels. Cases 5 and 6 assume the 10-fold reduction in the leak rates due tomanufacturing advances together with dependence on the superpressure, and superpressure dailyvariability, respectively. This analysis highlights the necessity for the manufacturingimprovements in the development of space balloons. As the result of these improvements, a Marsballoon can float in the Martian atmosphere for a full Martian year without replenishing lost gas,and carry just small additional amounts of gas (1.6 kg) to enable longer, 5-year long missions.In the design being developed in this study we are going to assume conditions of Case 6 for theNEAR term technology. This means that our design does not require gas replenishment for aone-Martian-year mission.With low gas loss through seam leaks, gas loss through diffusion may become important. TheNear-Term envelope material, as defined in the enabling technology matrix, consists of a layer ofpolyethylene that is 3 µm thick and a layer of Mylar 1 µm thick. The diffusive properties of the179

composite material are not known, so the diffusive properties of the components were studiedindependently. The polyethylene layer has a maximum diffusion rate of 2 m 3 /day or 0.8 g/day ofthe hydrogen gas for the nominal balloon with the radius of 17.24 m. The Mylar layer has amaximum diffusion rate of 0.08 m 3 /day or 0.03 g/day. It is reasonable to assume that theminimum daily diffusion rate of the components will define the maximum daily diffusion rate ofthe composite envelope material. Hence, when the materials are layered the resulting diffusionrate will likely be less than 0.03 g/day. This loss rate is much lower than any of the gas loss ratesshown in Table 5-10 and thus diffusion does not present a problem for gas loss with the chosencomposite material.180

6 Pathways To Architecture Realization6.1 IntroductionIn this section we report on our efforts to seek pathways to realization of the DARE architecture.Possible pathways for DARE architecture realization include: inclusion of a DARE-enabledmission as one of the core missions into the NASA Mars exploration roadmap, a DARE-enabledMars Scout mission, a DARE-enabled Mars mission funded by the European Space Agency(ESA). To facilitate the development of the DARE architecture along these pathways, duringPhase II we had contacts with NASA mission and program managers, provided inputs to Marsand Venus exploration advisory groups (MEPAG and VEXAG), and to NRC studies; establishedconnections with European colleagues; participated in conferences and briefings, and submittedtwo papers on the DARE concept. These efforts discussed in more detail below.6.2 Discussions with NASA Mission and Program PlannersWe have discussed the DARE concept with JPL Managers in the Mars Pre-Projects andAdvanced Studies Office. The Mars Program funding has been significantly reduced ascompared to earlier plans due to the rcent emphasis on the development of the STS replacementand lunar exploration. This funding reduction may present an opportunity for DARE missions asthey can below-cost.We have also learned about the process through which new technology can be inserted into theMars Program. The may come either from articulation of new science that requires newtechnology and/or from new advanced studies that identify new technologies. Our intention is tofind opportunities for insertion of the DARE concept into that process.After discussions with JPL Mars technology managers, we explored opportunities for partneringon Mars Scout proposals that would employ elements of the DARE architecture. Unfortunately,it is widely believed that NASA would not select an aerial vehicle for the current Mars Scoutproposal, hence even past supporters of the aerial vehicles at Mars are not planning to proposethem for Scout 2001/13 (the exception is the Aerial Regional-scale Environmental Survey ofMars or ARES, that was selected and then rejected at the last round of Scout proposals).6.3 Conferences, Papers, BriefingsDr. Pankine presented the DARE concept at the 6 th Annual NIAC Meeting in Seattle, WA onOctober 19-20, 2004.Dr. Pankine presented the DARE concept at the Fall Meeting of the American GeophysicalUnion (AGU) in San Francisco, CA on December 13-17, 2004. Figure 6-1 below shows theposter presented at the AGU.181

The poster generated substantial interest amongst the participants of the conference. The posterpresented a chance to raise awareness of the concept among planetary scientists and to discussvarious aspects of the concept with a wide spectrum of experts:- Luther Beedgle, JPL – imaging instruments;- Martin Knapmeyer, DLR German Aerospace Center – seismometer networks on Mars;- F.T. Flynn, NASA Ames Research Center – astrobiology instrumentation;- Alison Skelley, University of California Berkeley – astrobiology instrumentation - MarsOrganic Analyzer (MOA);- Bruce Murray, California Institute of Technology – impact craters in the Polar Regions astargets for high-resolution imaging;- Lori Fenton, Arizona State University – use of mesoscale models of Martian winds forDARE analysis and trajectory simulation;- J. M. Bryant, Wind Erosion Laboratory Department of Geography University of Guelph,Canada – observations of dust storm genesis from an airborne platform;- Andrew Ingersoll, California Institute of Technology – observations of the processes in thePolar Regions;- Shane Byrne, MIT, observations of the polar caps and layered terrain phenomena;and others.Figure 6-1. AGU DARE concept poster.182

We have submitted a white paper on the DARE concept to the Robotic Access to PlanetarySurfaces subcommittee of the NASA Robotic & Human Exploration of Mars StrategicRoadmapping Committee. The goal of submitting the white paper is to get NASA interested inthe DARE concept and to include the enabling technologies into the strategic explorationroadmap.Dr. Pankine gave a presentation on the DARE concept to the Mars Advanced Studies Group atthe JPL on May 12, 2005. The goal of the presentation was to introduce scientists, engineers andmanagers at JPL to the DARE concept and to receive feedback. The overreaching goal was to getNASA JPL interested in the DARE concept and in developing the enabling technologies. As theresult of this presentation a discussion of some of the engineering aspects of the concept wasinitiated with the JPL engineers (B. Mitcheltree).Dr. Pankine presented the DARE concept at the 17 th ESA Symposium on European Rocket andBalloon Programmes and Related Research that was held on May 30-June 2 2005 in Sandefjord,Norway. The paper submitted to the conference is attached to the report – see Appendix 1. Thepurpose of the talk was to introduce a broad international ballooning community to the DAREconcept and to receive feedback. The ultimate goal is to get the European Space Agency (ESA)interested in the DARE concept and in developing the enabling technologies. As the result of thepresentation Dr. Pankine received valuable feedback from A. Vargas who was previouslyinvolved with the Mars 96 planetary balloon mission and is a planetary balloon expert.Dr. Pankine presented a paper on the DARE concept at the AIAA 5th Aviation, Technology,Integration, and Operations Conference (ATIO), 26-28 September 2005, Arlington, VA. Thepaper is attached to the report in Appendix 2.Dr. Pankine gave a presentation on the DARE concept at the 7 th Annual NIAC meeting inDenver, CO, October 10-11, 2005. The presentation can be found in Appendix 3.Dr. Pankine participated in the meeting of the Mars Exploration Program Analysis Group(MEPAG) in Monrovia, CA on November 2-3, 2005 and of the Venus Exploration AnalysisGroup (VEXAG) in Pasadena, CA on November 4, 2005. The goal of participating in themeetings was to make the Mars and Venus communities aware of the DARE capabilities andinfluence technology development programs for Mars and Venus.The roadmap for Mars exploration is developed in more detail than the Venus explorationroadmap. Mars exploration roadmap (in its current state) does not include missions with anyaerial vehicle. However, DARE can become part of the Mars Exploration program as a Scoutmission in 2011/2013, or as the architecture for the Surface Network mission planned for 2020(see Figure 3-1).The Venus exploration community turned out to be more responsive to aerial vehicles, andballoons in particular. Dr. Pankine participated in the break out session of the AtmosphericScience group at the VEXAG meeting. The goal of the group was to provide technologydevelopment recommendations to the NASA managers. The group recommended (among otherthings) development of the vertical and horizontal atmospheric mobility technologies. Thisrecommendation may lead to a future follow up work on the DARE BGS system for Venus.183

As a follow up to the VEXAG meeting we have submitted an abstract for the DARE presentationfor the Chapman Conference that would focus on Venus exploration.At the MEPAG meeting Dr. Pankine had an opportunity to discuss a potential DARE MarsAstrobiology mission with Dr. Mike Mumma, Director of the NASA GSFC Center forAstrobiology. The contact with Dr. Mumma can potentially lead to a follow up Marsastrobiology mission study involving guided balloons.In November of 2005 we have initiated contact with the Caltech group that advocated a tierscalableapproach to planetary exploration (Dr. Wolfgang Fink). The approach has manysimilarities to the DARE architecture approach 1 .We have briefed Dr. Samad Hayati, the manager of the Mars Technology Program at JPL on theDARE concept on December 21, 2005. We have asked Dr. Hayati to suggest pathways forDARE further development.Dr. Hayati suggested Mars Scout opportunity of 2011/13. He clarified that Scout proposals canidentify technologies that require further development (for example, a technology at TechnologyReadiness Level (TRL) 4 needs to be advanced to TRL 6). In this case the Scout proposal wouldneed to contain the budget and the schedule for proposed technology development. NASA thendetermines what technologies are needed. Often, development of technologies identified inrejected proposals is being later sought through RFI.Some of the technologies identified in DARE are already being developed at JPL: lightweightsun sensor (~20 g); navigation; surface-relative navigation.Dr. Hayati suggested several “selling points” for DARE:- imaging sides of the cliffs;- pinpoint landing. While this technology is being developed by NASA, some sites on theplanet cannot be reached due to arrival geometry. DARE also offers additional benefit ofimmediate response to discovery.Dr. Hayati noted that in general technologies are being developed at NASA are tied to “baselinemissions” - Mars Science Lander (MSL), Astrobiology Science Lander (ASL), more capableMER-like rovers. These are so-called “focused technologies”. Other technologies are solicitedthrough NASA Research Announcements (NRA). An NRA for 2007 is expected in 2006.In December of 2005 we have hosted a briefing on the DARE concept for Dr. Dave Pieri of JPL.We have contacted Dr. Wolfgang Fink of Caltech, one of the authors of the paper on multi-tieredapproached to planetary exploration 1 . We have sent Dr. Fink materials on DARE concept andinvited him for a briefing at GAC Headquarters.184

We have contacted Dr. Marcel van den Berg of ESA regarding the Venus Aerobot study that heis leading 43 , and provided him with the summary of our DARE effort.We have submitted a technology summary on Venus BGS to the Venus ESA Cosmic VisionMission working group (E. Chassefière: eric.chassefiere@aero.jussieu.fr, M. Roos-Serote:roos@oal.ul.pt, D. Titov: titov@linmpi.mpg.de, C. Wilson: wilson@atm.ox.ac.uk, O. Witasse:owitasse@rssd.esa.int). The summary was prepared based on our modeling of the BGSperformance at Venus in Phase I and II of the effort.6.4 Media Interviews, Articles and Press ReleasesGAC has issued a press release on the development of the DARE concept on September 27,2005.“Sailing the planets: Exploring Mars with guided balloons.ALTADENA, CA – Mars rovers, Spirit and Opportunity, have, by now, spent almost two yearson the surface of Mars. They traveled several miles each, frequently stopping and analyzingscientific targets with their cameras, spectrometers and other instruments to uncover evidence ofliquid water on Mars in the past. Their mission is a smashing success for NASA.But what if NASA had a platform on Mars that was able to cover these distances in a matter ofhours instead and study the rocks on the surface in the same detail as rovers do? Scientific returnfrom such a vehicle would be immense – scientists would be able to study the whole planet ingreater detail in a time span of a single year.While orbiters can look at virtually any point on the surface of a planet, they lack the resolutionprovided by instruments on rovers or landers. Rovers, on the other hand, have limited mobilityand cannot travel very far from their landing site. As the atmosphere of Mars is very thin, anairplane at Mars would last for just an hour until it runs out of fuel.Global Aerospace Corporation of Altadena, CA proposes that the Mars exploration vehiclecombining the global reach similar to that of orbiters and high resolution observations enabled byrovers could be a balloon that can be steered in the right direction and that would drop smallscience packages over the target sites. The concept being developed by the Global AerospaceCorporation is funded by the NASA Institute for Advanced Concepts (NIAC).Balloons have been long recognized as unique, scientific platforms due to their relatively lowcost and low power consumption. Two balloons flew in the atmosphere of Venus in 1984. In thepast the inability to control the path of Mars balloons has limited their usefulness, and thereforescientific interest in their use.Global Aerospace Corporation has designed an innovative device, called Balloon GuidanceSystem (BGS) that enables steering a balloon through the atmosphere. The BGS is an43 A. Phipps et al., “Mission and System Design of a Venus Entry Probe and Aerobot”, IAC-05-A3.P.03, 56 th IAC, 17-21 Oct.2005, Fukuoka, Japan.185

aerodynamic surface – a wing – that hangs on a several kilometer-long tether below the balloon.The difference in winds at different altitudes creates a relative wind at the altitude of the BGSwing, which in turn creates a lifting force. This lifting force is directed sideways and can be usedto pull the balloon left or right relative to the prevailing winds.Floating just several kilometers above the surface of Mars, the guided Mars balloons can observerock formations, layerings in canyon walls and polar caps, and other features – at very highresolution using relatively small cameras. They can be directed to fly over specific targetsidentified from orbital images and to deliver small surface laboratories, that will analyze the siteat the level of detail rovers would do. Instruments at the balloon's gondola can also measuretraces of methane in the atmospheric and follow its increasing concentrations to the source on theground. This way the search for existing or extinct life on Mars can be accelerated.The figure illustrates a guided balloon platform (with exaggerated dimensions) operating at Marsoverlaying a Mars Express image [Copyright ESA/DLR/FU Berlin (G. Neukum)] of canyonwalls. The top of the balloon is aluminized, hence it reflects the Martian scene around it.Dr. Alexey Pankine of Global Aerospace Corporation presented the results of the study at theAIAA 5th Aviation, Technology, Integration, and Operations Conference (ATIO) in Arlington,VA on September 26, 2005.”186

The press-release generated substantial interested. National Geographic published a news storybased on the interview with Dr. Pankine. A partial list of the url’s to the sites that reprintedDARE story is below:WorldChanging.com: http://www.worldchanging.com/archives/003576.htmlNational Geographic:http://news.nationalgeographic.com/news/2005/10/1004_051004_mars_balloon.htmlSpaceDaily: http://www.spacedaily.com/news/balloon-05d.htmlEuricAlert: http://www.eurekalert.org/pub_releases/2005-09/gac-stp092605.phpSpace Flight Now: http://www.spaceflightnow.com/news/n0509/27marsballoon/Science Daily: http://www.sciencedaily.com/releases/2005/09/050928234356.htmPhysOrg.com: http://www.physorg.com/news6816.htmlRedNova News Space:http://www.rednova.com/news/space/253010/sailing_the_planets_exploring_mars_with_guided_balloons/BrightSurf.com: http://www.brightsurf.com/news/headlines/view.article.php?ArticleID=21123KeralaNext.com: http://www.keralanext.com/news/index.asp?id=386500LegendaryTimes: http://www.legendarytimes.com/index.php?op=news&func=news&id=4441MarsToday.com : http://www.marstoday.com/news/viewpr.html?pid=17898PostHumanBlues: http://posthumanblues.blogspot.com/2005/09/sailing-planets-floating-justseveral.htmlSpaceRef.com: http://www.spaceref.com/news/viewpr.html?pid=17898Universe Today: http://www.universetoday.com/am/publish/ballooning_mars.htmlAstroBioNet:http://www.astrobio.net/news/modules.php?op=modload&name=News&file=article&sid=1728&mode=thread&order=0&thold=0Mars Daily: http://www.marsdaily.com/news/balloon-05d.htmlFebruary 14, 2006 issue of New Scientist.com published an article on planetary ballooning basedin part on the interview with Dr. Pankine (“See Mars and Venus by balloon”, 14 February 2006,Kurt Kleiner, Magazine issue 2538).187

Dr. Pankine was interviewed by J. Papalardo, Associate Editor of the Smithsonian Air and SpaceMagazine. The article is scheduled to appear in the June/July 2006 edition of the magazine.Dr. Pankine was interviewed by Leonard David, writer for the popular science web siteSpace.com.188

7 SummaryThis report describes the work accomplished and results obtained during Phase II of thedevelopment of the concept for new planetary exploration architecture. During Phase II, we havemade significant progress in developing the concept as evidenced by the following summary ofkey accomplishments:• Developed advanced mathematical models of Balloon Guidance Systems (BGS)• Studied BGS performance in the atmospheres of Mars and Venus• Defined revolutionary science applications enabled by the DARE architecture• Researched potential instruments in support of the revolutionary science• Performed DARE platform design and mission trade studies• Developed conceptual design of the DARE platform• Developed conceptual mission scenario• Simulated and studied atmospheric trajectories of guided balloons at Mars• Researched enabling and enhancing technologies, developed a roadmap of technologydevelopment• Presented the DARE concept at several professional meetings, briefed NASA engineers,scientists and managers.• Completed this Final ReportDuring Phase II of the study we estimate that we have met nearly 100% of our objectives.189

8 Appendices190

8.1 Appendix 1. “Sailing the Planets: Planetary Science from GuidedBalloons”, paper presented at the 17 th ESA Symposium on EuropeanRocket and Balloon Programmes and Related Research, May 30-June 2,2005, Sandefjord, Norway.191

SAILING THE PLANETS: PLANETARY SCIENCE FROM GUIDED BALLOONSDr. Alexey Pankine (1) , Dr. Kim Aaron (1) , Nathan Barnes (1) , Kerry Nock (1)(1) Global Aerospace Corporation, 711 W. Woodbury Rd., Suite H, Altadena, CA 91001, USA,alexey.a.pankine@gaerospace.comABSTRACTWe present a new concept for a future (10-40 yearsfrom now) planetary exploration architecture. At thecore of the architecture are the Directed Aerial RobotExplorer (DARE) platforms, which are autonomousballoons with path guidance capabilities that can carryheavy scientific payloads and deploy swarms ofminiature robotic probes over multiple target areas. Thisarchitecture enables a multitude of observations that areimpossible or too expensive to make in any other way.The architecture enables surface imaging, in situsampling of the atmosphere and surfaces, radarsoundings, magnetic and gravity surveys and otherobservations at an unprecedented resolution. We presentan overview of the concept in the context of Marsexploration.innovative, lightweight Balloon Guidance System(BGS) and multiple, lightweight, deployablemicroprobes into a revolutionary architecture forplanetary exploration.1 CONCEPT OVERVIEWGlobal Aerospace Corporation is developing arevolutionary architecture that opens new and excitingpathways for planetary exploration. The work issupported by the NASA Institute for AdvancedConcepts (NIAC). The innovative system architecturerelies upon the use of Directed Aerial Robot Explorers,which are long-duration-flight autonomous balloonswith path guidance capabilities that can deploy swarmsof miniature probes over multiple target areas. TheDirected Aerial Robot Explorer (DARE) platforms willexplore the planets in concert with orbiter(s) and surfaceplatforms (landers, rovers, microprobes). The keyelements of the overall DARE architecture are: 1) longdurationplanetary Mars balloon; 2) light-weight balloonguidance system; 3) lightweight and efficient powergeneration and energy storage; 4) multiple lightweightdeployable microprobes; and 5) communication orbiter.They are shown schematically in Fig. 1.Balloons have been long recognized as unique, low-costscientific platforms due to their relatively low cost andlow power consumption. Indeed, the successful Venera-Vega Project [1] demonstrated technical feasibility ofdeploying a balloon on another planet and performingscientific observations from it. Concepts andtechnologies enabling planetary balloon exploration ofMars, Venus, Titan and the Outer Planets have beendeveloped [2,3,4,5,6,7]. The DARE architectureadvances these concepts to the next level of utility anduniversality by integrating the balloon platform with theFig. 1. Key elements of the DARE conceptThe DARE concept represents a highly adaptiveobservational platform capable of observing planetaryatmospheres and surfaces over long periods of timewithout consuming much power. A DARE platformwould orbit the planet using winds to guide theirtrajectory according to observational objectives. Studiesof the atmospheric dynamics, atmospheric chemical,and radiative processes on other planets would becomepossible at an advanced level. Small microprobes wouldbe deployed over the target areas and perform amultitude of tasks at the surface or while descending,such as chemical, biological, meteorological, or thermalanalyses, high-resolution imaging, measuring seismicactivity, etc. The data would be transmitted in real timeto the overflying DARE platform, processed ortemporarily stored onboard, and then relayed to theorbiter, or transmitted to the orbiter directly.At the heart of the DARE concept are long-durationplanetary balloons with trajectory control capabilitiescalled DARE platforms. A conceptual drawing of theDARE platform is shown in Fig. 2. The figure shows aMars balloon with a gondola and the deployed SinglewingBalloon Guidance System (BGS) on a long tetherbelow it. The Single-wing BGS is a wing and a rudderattached to horizontal boom. The drawing is forillustration purposes and is not to scale: the tether will

e several km long and the BGS will be much smallerthan the balloon.The difference in winds at different altitudes in theatmosphere creates a relative wind at the altitude of thewing (stronger winds are usually found at higheraltitudes on Mars). The relative wind creates the liftingforce that is directed sideways. The lifting forcedepends on the size of the wing, its aerodynamicproperties, the density of the atmosphere and on thestrength of the relative wind (which ultimately dependson the length of the tether).The BGS enables the DARE platform to be maneuveredrelative to the prevailing atmospheric winds and permitstargeted observations according to the missionobjectives. The BGS requires very little power (about 1W on average) to operate and can be made very light. Itis suspended below the gondola on a long (5-11 km)tether, depending on the height of the balloon above thesurface. The DARE platform employs a superpressureballoon and will remain aloft from 1 to 5 years and isable to visit different regions of the planet. SeveralDARE platforms can form a constellation and performsimultaneous observations over the planet. The gondolaon the DARE platform carries several small lightweightdeployable microprobes that can be released over targetsites to perform in situ analysis of the atmosphere orsurface. The gondola also houses scientific instruments,computers, the BGS deployment system (a lightweightwinch), solar panels, batteries and antennas forcommunication.Preliminary numerical analysis indicates that arelatively small (1 m 2 ) BGS wing is capable of movinga balloon platform with the velocity of the order of 1m/s in the direction perpendicular to the direction of thewinds at the balloon altitude (crosswind). The small butconstantly applied offset can result in quite significantchanges in a balloon trajectory.Technologies conceptualized in this study - such as verylong-duration-flight balloons, flight path guidance,microprobes, and planetary platform navigation andcommunications - have direct relevance to future in situexploration missions to Mars and other planets. Theresults of this study can be applied directly to thedevelopment of future missions to Mars, Venus andTitan.Below we describe the exploration applications of theconcept in the context of Mars exploration. We presentan outline of enabling technologies and describeconceptual Mars DARE platform. We show guidedballoon trajectory simulations at Mars illustrating thefeasibility of the proposed balloon guidance system.Fig. 2. Conceptual drawing of the DARE platform (notto scale)2 EXPLORATION APPLICATIONSDARE platforms can provide spatial coveragecomparable to that of satellites, but they additionallyenable opportunities for in situ atmospheric and surfaceanalysis with deployable microprobes and highresolutionsurface imaging. At Mars, DARE platformswould float close to the surface (6-12 km, depending onlocation and season) and could provide a wealth of newand unique observations. Some observations, such asobservations of magnetic anomalies on Mars, are quitepossibly only feasible from a suborbital platform. In thepast the inability to control the path of Mars balloonshas limited its usefulness, and therefore scientificinterest in its use. Without flight path guidancetechnology, a Mars balloon has a high probability ofimpacting high topography and it cannot be commandedto float over a particular study region. The BGS canvastly expand the capabilities of balloons for MarsExploration by providing the means to control theirpaths in the Martian atmosphere. In addition, a BGSreduces the risk of mission failure by avoiding regionswith high topography.The extended range of DARE platforms can provideopportunities for highly adaptive observations duringscience missions. Just like rovers, if an interesting targetis found, a DARE platform can be commanded toreposition itself to observe it. However, the range ofguided balloons is the entire planet, not the immediatevicinity of a rover landing site. A DARE platform can

deploy small rovers, miniature geo-chemicallaboratories or it can distribute small surface andatmospheric sampling probes over Mars at the site ofinterest. In addition, they can be deployed with greattargeting precision, an important goal for future smallprobes.These small probes, landers or rovers do not requireindividual heat shielding for atmospheric entry and thuscould be miniaturized so that many could be carried. Aguided balloon can deliver multiple seismological,surface heat flux or meteorological stations to preselectedlocations to form a network of surface stations.A single DARE platform carrying a 30 kg payloadwould be able to deliver fifteen 2 kg probes to differentlocations. DARE platforms could deploy networks ofseismological and meteorological surface stations.Small probes would be deployed over potential landingsites to provide close up images of the terrain and toemplace navigation beacons. DARE platforms can carryarrays of magnetometers to study crustal magneticanomalies.DARE platforms could provide a new approach to roversite selection. DARE platforms could provide missionplanners with detailed information for planning futurerover missions, scout potential sites for sample returnmissions, and provide detailed high-resolution imagingon the distribution of rocks, slopes, and other hazards atpotential landing sites. Deployed surface microprobescould give preliminary readings on the surfacecomposition or presence of telltale signs of life.Multiple landing site options can be visited over thecourse of a single DARE mission and landing siteselection could be made with enhanced confidence.Furthermore, this approach could be extended toexploring routes for rover sample collecting excursions,choosing landing sites for human exploration, andchoosing areas to search for life.DARE platforms can provide data to help select sites forlanding subsurface exploration probes and they can alsodeploy these probes. DARE platforms can carrysounding radars and neutron spectrometers to determinelocations where underground ice or water are closest tothe surface. By reducing the amount of drilling ordigging involved in obtaining a sample of theunderground ice or water, the mission’s chances forsuccess will be greatly improved.The proposed architecture will have applications notjust at Mars, but also at Venus, Titan and Jupiter [9].3 ENABLING TECHNOLOGIESA number of technologies were identified that enablethe proposed DARE architecture. These include:advanced materials for the design of Mars balloons;balloon guidance system; entry, descent and inflation(EDI) technology; navigation and guidance algorithmsin Mars winds; and balloon performance modelsdevelopment. Below we discuss with more details thefirst three of these technologies.There are other technologies, not listed above, that arenecessary or enhancing, but we do not discuss them hereas they will come about one way or the other fromcommercial, defense or space systems developments.3.1 Advanced Balloon Envelope MaterialsMaximizing the useful mass of the balloon’s scientificpayload requires reducing the masses of all the othercomponents. Advanced lightweight and strong materialsneed to be developed to reduce masses of parachutes,heat shields, balloon envelope and other componentsthat take up a great proportion of the mass delivered toMars.Of special importance are the materials for balloonenvelopes. Reducing the surface density of the balloonfilm by using composite materials (such as Mylar-Kevlar-Polyethylene scrim [5]) can lead to substantialincreases (~10 kg) in useful payload mass. The highatmospheric altitudes (10-20 km) and heavy payloads(~100 kg) envisioned in our concept require largeballoons (20-70 m in diameter). The greater stressexperienced by the envelopes of these large balloonswill also be supported by the advanced materials.3.2 Balloon Guidance SystemThe ability to guide DARE platforms to the sites ofinterest or to avoid high topography is crucial to theDARE architecture. The proposed BGS consists of anaerodynamic surface (wing) hanging below the balloonon a very long (several km) tether. Several designs arebeing investigated: a Single-wing BGS (see Fig. 2) anda Dual-wing BGS that can be seen on Fig. 1. Themagnitude and the direction of the lifting force can becontrolled by changing the roll angle and the angle-ofattackof the wing. The horizontal component of thetotal force produced by the wing can be used to changethe path of a balloon in the winds. Preliminarynumerical analysis indicates that a guidance system witha relatively small (1 m 2 ) BGS wing is capable ofmoving a conceptual DARE platform with the velocityof the order of 1 m/s in the crosswind direction.The aerodynamic surface of the BGS will be operatingat low Reynolds numbers (~1000). For these lowReynolds numbers, drag coefficients are significantlyhigher and maximum lift coefficients are a little lowerthan for higher Reynolds number operation. The

educed Lift-to-Drag ratio is quite a challenge forsystems which must generate their lift aerodynamicallywhile providing power to overcome the drag. However,for our system, the weight is supported by buoyancy.The "lift" from the wing is directed close to horizontaland predominantly across the flight path of the balloon.The drag acts mostly to slow down the balloon, and isrelatively unimportant to the operation of our system.One of the advantages of the proposed BGS is that itdoes not require power for propulsion. A small amountof power (~1 W on the average) is needed only forcommunications and to adjust the control surfaces of theguidance system once a day, or even less frequently,depending on the path control objectives. Alternativeapproaches to balloon path guidance employing, forexample, an engine driven propeller would requiresignificantly higher levels of continuous power input(~1000 W).3.3 Entry, Descent and InflationEntry, Descent and Inflation (EDI) are the sequence ofevents that occur over a short time-span following thearrival of the spacecraft carrying the DARE platform toa planet and before the platform achieves its floatingaltitude in the atmosphere of a planet. Fig. 3 belowillustrates the sequence of events that take place duringEDI.terminal velocity on the parachute, the inflation phasecommences. The gondola and attached equipment arelowered, and the weight is used to stretch out theballoon envelope vertically. The shock generated duringdeployment of the envelope as it goes taut is absorbedby a load tether inside the balloon. The inflationequipment is mounted below the balloon, and will bejettisoned when the inflation of the balloon is complete.An inflation tube is used to carry the gas up to the initialinflation bubble towards the upper end of the balloonand also to support the deployment shock. Once theballoon has reached a size comparable to the parachute,the parachute will be cut away. Upon completion ofinflation the inflation hardware can be released, and thesystem will ascend towards an equilibrium float altitude,at which point the trajectory control system can bedeployed.An alternative approach that was considered andrejected is to land prior to inflating the balloon. Thedisadvantages of this approach are the increased massdue to the necessary landing system and the danger thatthe balloon can be damaged during inflation because ofthe interference by significant horizontal winds.The Mars balloon EDI technology is currently beingstudied at NASA.4 MARS DARE PLATFORM DESCRIPTIONThis section gives an overview of our currentunderstanding of the conceptual design of the MarsDARE platform.4.1 Mars BalloonsFig. 3. EDI sequenceThe entry and descent phases would be very similar toprevious Mars missions. A blunt conical heat shieldreduces the enormous interplanetary velocity andabsorbs the heat generated by passing through theatmosphere. The aft cover is released and a supersonicparachute is deployed at a Mach number of ~2. Thisslows the descent rate significantly into the lowsubsonic range to the point the internal components nolonger need protection from the ram air and the heatshield can be dropped. As the system approachesThe Mars DARE platforms will employ superpressureballoons that maintain a constant atmospheric densitylevel in the atmosphere. The height of a constant densitylevel above a reference level will vary by about akilometer depending on the season, diurnal phase,terrain altitude and location on Mars. We envisionballoons with an aluminized top and white paint bottomto prevent condensation of CO 2 on the surface of theballoon during periods of extreme cold (temperaturesbelow 140K), possible at night in the Polar Regions.Films with controllable thermo-optical properties maybecome available in the future, and these will help inreducing the stress on the envelope due to heating of theenvelope.4.2 Balloon Guidance SystemThe BGS will be made from lightweight materials(Carbon-Carbon composite) with a total mass of a fewkg. The BGS will be deployed from its stowage positionon the gondola a short time after completion of the

alloon inflation during atmospheric descent. The BGSwill be folded and stowed below the gondola in theentry probe on the way to Mars. The deployment systemmay include a lightweight winch. The BGS willcommunicate with the control computer on the gondolavia a radio link.4.3 Lightweight MicroprobesGuided DARE platforms are well suited to distributelightweight surface and atmospheric sampling packagesover the planet. These probes won’t need the heatshielding for atmospheric entry and thus could beminiaturized. In addition, they can be deployed withgreater targeting precision, an important goal for futuresmall probes.5 MARS DARE TRAJECTORIESTrajectories of guided Mars balloons were simulatedusing “real” winds from the Mars Global CirculationModel (MGCM) [8]. As the cross-track controlvelocities depend on the wind difference between thealtitude of the balloon and the BGS, they change as theplatform moves through the atmosphere and encountersdifferent wind conditions. In the analysis shown belowthe cross-track control velocities were constant duringsimulations for simplicity. However, these controlvelocities do correspond to the most probable winddifferences that are expected for the analyzed seasonand locales.A variety of innovative approaches to vehicle shapes,delivery methods and tracking are possible. Forexample, the miniature surface geo-chemical laboratorycan resemble the NASA’s Deep Space 2 probe in itsdesign – a hard shell staying at the surface with thepenetrator or drill sampling the surface. Small droppablerovers can be deployed to investigate surface sites inmore detail over short periods of time.4.4 Sensor TechnologiesDevelopment of small sensors with new capabilities willenable new applications for DARE platforms.Miniaturized pressure-temperature sensors can be usedon dropsondes to provide multiple soundings of theatmosphere on a global scale. In another example,embedding a radar antenna into a balloon envelope canprovide unprecedented resolution for subsurfacesoundings. Small seismometer-penetrators can be usedto create a network of seismological stations acrossMars.Fig. 4. Trajectory of the free-floating balloon over NorthPole at Mars, L s =180°4.5 Power Generation and Energy StoragePower generation and energy storage are importanttechnologies that would enable the DARE platform tostudy Polar Regions at night or employ powerful radars.Amongst these technologies we consider improved solarcells (30% efficiency CSi arrays), fuel cells (400-800W/h-kg), wind turbines and other technologies.4.6 Communications OrbiterWe assume that a telecommunications orbiter, similar incapabilities to the Mars Telecommunication Orbiter(MTO) will be available at the time of the architectureimplementation. The MTO-like orbiter will enableusage of small antennas and transmitters – similar insize and capabilities to the communication package ofthe Mars Exploration Rovers (MERs).Fig. 5. Trajectory of the DARE platform over NorthPole at Mars, L s =180°Fig. 4 and Fig. 5 show trajectories of the free-floatingballoon and of the DARE platform, respectively. Thefigures show the Martian North pole in polar projection.

The season is the Northern Hemisphere summer(L s =180°).The goal of this simulation was to demonstrate guidanceof the DARE platform in relatively weak summer windsin Polar Regions. The cross-track control velocity wasconstant in this simulation at 0.1 m/s corresponding toweak vertical gradients of the winds in the polarsummer atmosphere. The path guidance objective wasto keep the balloon at the latitude of 60° N. As can beseen from Fig. 5 the guidance is quite efficient and theDARE platform trajectory is confined to a narrowlatitudinal band in between 60° and 70° N. In contrast,the free-floating balloon in Fig. 4 quickly driftsnorthward after the start of the simulation and remainsin the vicinity of 80° N latitude afterwards.6 SUMMARYA concept for a new architecture for planetaryexploration is described in the context of Marsexploration. The key elements of the architecture are:long-duration-flight autonomous balloons, balloontrajectory control, lightweight power generation andstorage, and multiple, deployable microprobes foratmosphere and surface exploration. A relatively smalland light balloon trajectory control device would enablerepositioning the platform on a global scale for in situanalysis and targeted deployment of atmospheric andsurface probes. Deployment of probes from balloonseliminates atmospheric entry and deceleration hardwarethus reducing probe mass and permitting more sciencepayload or more probes. The DARE architecture willenable low-cost, low-energy, long-term globalexploration of the atmosphere and surface of Mars andother planets. Additional information can be found atwww.gaerospace.com/projects/DARE/DARE.html7 REFERENCES1. Sagdeev, R. Z. et al. Overview of the VEGA Venusballoon in situ meteorological measurements, Science,231, 1411-1422, 1986.2. Cutts, J. A. et al. Venus Aerobot MultisondeMission, AIAA Balloon Tech. Conference, 1-10, 1999.3. Greeley, R. et al. The Mars Aerial Platform MissionConcept, AIAA paper 96-0335, 1996.Fig. 6. Trajectory of the DARE platform guided fromthe Southern to the Northern hemisphere at Mars.Fig. 6 shows the simulated trajectory of the DAREplatform crossing from the Southern Hemisphere intothe Northern Hemisphere. The background shows thelow-resolution map of Martian topography, with thehighest elevations shown in lighter shades (Tharsisregion centered at about -90° W longitude, 0°longitude), and the lowest elevations shown in darkershades (Hellas basin centered at about 60° E longitude,–45° S longitude). The season is late spring in theSouthern Hemisphere (L s =255°). The cross-trackcontrol velocity of the DARE platform in this case was1 m/s. The DARE platform starts at –60° S, 180° E(lower right corner of the figure). The control objectivefor this simulation was to transport the platform fromthe Southern Hemisphere to the 80º N latitude. Thestrong equatorial flow at this season threatens to run theDARE platform into the high topography of the Tharsis.The guided DARE platform is able to avoid crashinginto the Tharsis by crossing the equatorial region beforeit gets pulled into the zonal equatorial flow. Thisanalysis illustrates that the proposed BGS is useful inguiding balloons at Mars.4. Jones, J. A. and Heun, M. K. Montgolfier BalloonAerobots for Planetary Atmospheres, AIAA Paper 97-1445, 1997.5. Nock, K.T. et al. Overview of a Mars 2001Aerobot/Balloon System, 12th AIAA Lighter-Than-AirTechnology Conference, San Francisco, 19976. SAIC, Titan exploration with advanced systems,NASA CR-173499, 1983.7. Tarrieu, C. Status of the Mars 96 AerostatDevelopment, 44th Congress of the InternationalAstronautical Federation, paper IAF-93-Q.3.399, 19938. Haberle, R. M. et al. Mars atmospheric dynamics assimulated by the NASA Ames General CirculationModel, J. of Geophys. Res., 98, 3093-3123, 1993.9. Pankine, A. et al. Directed Aerial Robot Explorers(DARE) for Planetary Exploration, 34th ScientificAssembly of the Committee on Space Research, PaperPSB1-0074-02, 2002.

8.2 Appendix 2. “Sailing the Planets: Exploring Mars from GuidedBalloons”, paper presented at the AIAA 5th Aviation, Technology,Integration, and Operations Conference (ATIO), 26-28 September 2005,Arlington, VA.192

AIAA 5th Aviation, Technology, Integration, and Operations Conference (ATIO)26 - 28 September 2005, Arlington, VirginiaAIAA 2005-7320Sailing the Planets: Exploring Mars from Guided BalloonsAlexey A. Pankine * , Kim M. Aaron † , Nathan C. Barnes ‡ , Kerry T. Nock §Global Aerospace Corporation, Altadena, CA, 91001We present a new concept for a future Mars exploration architecture. At the core of thearchitecture are the Directed Aerial Robot Explorer (DARE) platforms, which areautonomous balloons with path guidance capabilities that can carry heavy scientificpayloads and deploy swarms of miniature robotic probes over multiple target areas. Thisarchitecture enables a multitude of observations that are impossible or too expensive tomake in any other way. The architecture enables surface imaging, in situ sampling of theatmosphere and surface, radar soundings of the subsurface, magnetic and gravity surveysand other observations at an unprecedented resolution. We present an overview of theconcept in the context of Mars exploration.I. IntroductionGLOBAL Aerospace Corporation isdeveloping a revolutionary architecture thatopens new and exciting pathways for planetaryexploration. The work is supported by the NASAInstitute for Advanced Concepts (NIAC). Theinnovative system architecture relies upon the useof Directed Aerial Robot Explorers, which arelong-duration-flight autonomous balloons withpath guidance capabilities that can deploy swarmsof miniature probes over multiple target areas.The Directed Aerial Robot Explorer (DARE)platforms will explore the planets in concert withorbiter(s) and surface platforms (landers, rovers,microprobes). The key elements of the overallDARE architecture are: 1) long-duration planetaryMars balloon platform(s); 2) balloon flight pathguidance; 3) autonomous navigation and control;4) lightweight and efficient power generation andenergy storage; 4) lightweight deployable sciencepackages; 5) communication relay orbiter; and 6) Figure 1. Key elements of the DARE conceptcooperation between surface, airborne and orbitalelements of the architecture. The elements of the architecture are shown schematically in Fig. 1.Balloons have been long recognized as unique, low-cost scientific platforms due to their relatively low cost andlow power consumption. Indeed, the successful Venera-Vega Project 1 demonstrated technical feasibility ofdeploying a balloon at another planet and performing scientific observations from it. Concepts and technologiesenabling planetary balloon exploration of Mars, Venus, Titan and the Outer Planets have been developed 2-7 . TheDARE architecture advances these concepts to the next level of utility and universality by integrating the balloonplatform with the innovative, lightweight Balloon Guidance System (BGS) and multiple, lightweight, deployablemicroprobes into a revolutionary architecture for planetary exploration.* Project Scientist, 711 W. Woodbury Rd., Altadena, CA, 91001.† Chief Engineer, 711 W. Woodbury Rd., Altadena, CA, 91001‡ Engineer, 711 W. Woodbury Rd., Altadena, CA, 91001, AIAA Member.§ Senior Engineer, 711 W. Woodbury Rd., Altadena, CA, 91001, AIAA Member1American Institute of Aeronautics and AstronauticsCopyright © 2005 by Global Aerospace Corporation. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

II. Concept OverviewAt the heart of the DARE architecture concept are long-duration planetary balloons with trajectory controlcapabilities called DARE platforms. A conceptual drawing of the DARE platform is shown in Fig. 2. The figureshows a Mars balloon with a gondola and the deployed Single-wing Balloon Guidance System (BGS) on a longtether below it. The Single-wing BGS is a wing and a rudder attached to horizontal boom. The drawing is forillustration purposes and is not to scale: the tether will be several km long and the BGS will be much smaller thanthe balloon.The DARE platform is highly adaptive and capable of observingplanetary atmospheres and surfaces over long periods of timewithout consuming much power. A DARE balloon would orbit theplanet using winds to guide their trajectory according toobservational objectives. Studies of atmospheric dynamics,atmospheric chemical, and radiative processes on other planetswould become possible at an advanced level. Microprobes would bedeployed over the target areas and perform a multitude of tasks atthe surface or while descending, such as chemical, biological,meteorological, or thermal analyses, high-resolution imaging,measuring seismic activity, etc.The BGS enables the DARE platform to be maneuvered relativeto the prevailing atmospheric winds that makes possible targetedobservations according to the mission objectives. The BGS requiresvery little power (about 1 W on average) to operate and can bemade very light. It is suspended below the gondola on a long (5-11km) tether, depending on the height of the balloon above thesurface. The DARE platform employs a superpressure balloon thatwill remain aloft from 1 to 5 years and is able to visit differentregions of the planet. Several DARE platforms can form aconstellation and perform simultaneous observations over theplanet. The gondola on the DARE platform carries several smalllightweight deployable microprobes that can be released over targetsites to perform in situ analysis of the atmosphere or surface. Thegondola also houses scientific instruments, computers, the BGSdeployment system (a lightweight winch), solar panels, batteriesand antennas for communication.Preliminary numerical analysis indicates that a relatively small(1 m 2 ) BGS wing is capable of moving a balloon platform with thevelocity of the order of 1 m/s in a direction perpendicular to thedirection of the winds at the balloon altitude (crosswind). Thissmall, but constantly applied, offset can result in quite significantchanges in a balloon trajectory.Technologies needed for DARE platforms - such as very longduration-flightballoons, flight path guidance, microprobes, andplanetary platform navigation and communications - have directrelevance to future in situ exploration missions to Mars and otherplanets. The results of this study can be applied directly to thedevelopment of future missions to Mars, Venus and Titan.Below we discuss the development of the DARE architecture.Figure 2. Schematics of the DAREplatform.III. DARE Architecture DevelopmentIn developing the DARE architecture we start by defining the exploration capabilities afforded by the newarchitecture. We then consider the observations and measurements together with potential instruments that enablethe new exploration capabilities. The next step is to define potential mission scenarios – this illuminates the enablingtechnologies, demonstrates feasibility of the concept and helps to focus the design. Based on the analysis of the2American Institute of Aeronautics and Astronautics

mission scenarios we can identify enabling technologies and develop conceptual design. The sections belowdescribe these steps in the DARE architecture development in more detail.A. Exploration CapabilitiesDARE platforms can provide spatial coveragecomparable to that of satellites, but theyadditionally enable opportunities for in situatmospheric and surface analysis with deployablescience packages and high-resolution surfaceimaging. At Mars, DARE platforms would floatclose to the surface (6-12 km, depending onlocation and season) and could provide a wealthof new and unique observations. Someobservations, such as observations of magneticanomalies on Mars, are, quite possibly, onlyfeasible from a suborbital platform. In the past theinability to control the path of Mars balloons haslimited their usefulness, and therefore scientificinterest in their use. Without flight path guidancetechnology, a Mars balloon has a high probabilityof impacting high topography and it cannot becommanded to observe a particular region ofinterest. The BGS vastly expands the capabilitiesof balloons for Mars Exploration by providing themeans to control their paths in the Martianatmosphere. In addition, a BGS reduces the riskof mission failure by avoiding regions with high topography.Figure 3. Olivine outcrop and DS-2 landing ellipse(NASA/JPL/ASU)DARE platforms will be capable to float heavy and power intensive payloads – 90 kg and 200 W in the 3 to 10years time horizon and 170 kg and 400 W beyond 10 years. These payloads could include large telescopes for ultrahighresolution observations of the surface or radars for subsurface soundings.The extended range of DARE platforms can provide opportunities for highly adaptive observations duringscience missions. Just like rovers, if an interesting target is found, a DARE platform can be commanded toreposition itself to observe it. In this case, the range is the entire planet, not the immediate vicinity of a rover landingsite. A DARE platform can deploy small rovers, miniature geo-chemical laboratories or it can distribute smallsurface and atmospheric sampling probes over Mars at particular sites of interest. In addition, these small probes canbe deployed with great targeting precision, an important goal for future small probes. Fig. 3 shows a thermal map ofa portion of the Martian canyon with a narrow band of the olivine outcrop on the bottom (shown in purple). Theyellow ellipse shows the landing error ellipse for a direct-entry probe landing from approach to the planet (theellipse’s dimensions are approximately 120 by 20 km, typical of the failed Deep Space 2 probe landing accuracy).This example shows that interesting geology on Mars, with spatial dimensions on the order of a kilometer, would bevery hard to target with conventional, direct-entry probes. DARE platforms, on the other hand, can target geologicfeatures with deployable science packages with greater precision.DARE platforms could provide a new approach to rover site selection, by providing mission planners withdetailed information for future rover missions, scouting potential sites for sample return missions, and providingdetailed high-resolution imaging on the distribution of rocks, slopes, and other hazards at potential landing sites.Microprobes could give preliminary readings on the surface composition or presence of telltale signs of life.Multiple possible landing sites could be visited over the course of a single DARE mission with final landing siteselection being made with enhanced confidence. Furthermore, this approach could be extended to exploring routesfor rover sample collecting excursions, choosing landing sites for human exploration, and choosing areas to searchfor life.DARE platforms can provide data to help select sites for landing subsurface exploration probes and they can alsodeploy these probes. DARE platforms can carry sounding radars and neutron spectrometers to determine locationswhere underground ice or water are closest to the surface. By reducing the amount of drilling or digging involved inobtaining a sample of the underground ice or water, the mission’s chances for success will be greatly improved.The proposed architecture will have applications not just at Mars, but also at Venus, Titan and Jupiter 9 .3American Institute of Aeronautics and Astronautics

Figure 4. Trajectory of the free-floating balloon Figure 5. Trajectory of the DARE platform overover North Pole at Mars, Ls=180°North Pole at Mars, Ls=180°Fig. 4 and Fig. 5 illustrate the balloon guidance capabilities of the DARE platform. Fig. 4 and Fig. 5 showtrajectories of the free-floating balloon and of the DARE platform, respectively. The figures show the Martian Northpole in polar projection. The season is the Northern Hemisphere summer (Ls=180°). The goal of this simulation wasto demonstrate guidance of the DARE platform in relatively weak summer winds in Polar Regions. The cross-trackcontrol velocity was constant in this simulation at 0.1 m/s corresponding to weak vertical gradients of the winds inthe polar summer atmosphere. The path guidance objective was to keep the balloon at the latitude of 60° N. As canbe seen from Fig. 5 the guidance is quite efficient and the DARE platform trajectory is confined to a narrowlatitudinal band in between 60° and 70° N. In contrast, the free-floating balloon in Fig. 4 quickly drifts northwardafter the start of the simulation and remains in the vicinity of 80° N latitude afterwards.B. Potential Observations and InstrumentsScience mission applications for DARE at Mars would include Mars sample return assist - (see below inMission Scenarios); search for present or past life habitats (hydrothermal sources); characterization of the water,atmosphere and climate on Mars in the past and today; and study of the structure and evolution of Mars.In view of these objectives DARE enables a number of observations on the global scale:- in situ analysis of the atmosphere (altitude profiles, abundances, isotopic ratios);- high spatial resolution imaging (1-10 cm) of the surface and subsurface (visible and IR cameras, spectrometer,radar, radiometer, etc);- high-spatial resolution (1-10 km) magnetic andgravity surveys;- emplacement of networks of surface stations -seismological, geophysical (surface heat flux, soilproperties, subsurface structure - magnetometers),meteorological (pressure, temperature, near surfaceprofiles of atmospheric parameters, surface winds);- targeted delivery of surface probes – penetrators,biological, mineralogical, chemical in-situ analyzers.Consider, for example, one of the potentialmeasurements afforded by the DARE architecture - in situmeasurements of the atmospheric constituents, in particular– the measurement of the fractionation ratio of the carbonisotopes C 12 and C 13 . This measurement supports thesearch for microbial life on Mars. Methane-makingorganisms discriminate between isotopes as they feed on a4American Institute of Aeronautics and AstronauticsFigure 6. Tunable Laser Spectrometer forAtmospheric and Sub-surface gas measurements onMars (NASA JPL)

global reservoir of CO 2 . If the measured ratio ofthe carbon isotopes in methane were differentfrom the isotope ratio in the CO 2 , it would offerstrong evidence for a biological source. Aperfect candidate instrument for thesemeasurements is a Tunable Laser Spectrometer(TLS). A TLS that is being developed at NASAJPL for atmospheric and subsurface gasmeasurements on Mars is shown on Fig. 6 Theinstrument is very light and small. It can becarried on board the DARE platform or beincorporated into deployable probes. DAREplatforms carrying TLSs enable the planetarywidesearch for biological sources.In another example, DARE can survey thecrustal magnetic anomalies, discovered in theSouthern Hemisphere by the Mars GlobalSurveyor (MGS) 10 . Fig. 7 shows the map of theanomalies. Understanding these anomalies willprovide clues on crustal evolution on Mars andFigure 7. Map of crustal magnetic anomalies on Mars(GSFC NASA)advance understanding of the Martian extinct dynamo. Orbital measurements lack resolution to study the anomalies.A DARE platform with an array of magnetometers enables high-resolution observation of crustal magneticanomalies and detection of weak anomalies via gradient measurements.C. Potential Mission ScenariosWe have developed several mission scenarios, namely Crustal Magnetic Anomalies Survey, Surface NetworkEmplacement and Mars Sample Return assist for analysis in the study. We have used the Surface NetworkEmplacement Scenario to develop conceptual design of the DARE platform.An objective of the Surface Network Emplacement mission would be the establishment a network of 5meteorological and seismological stations on the surface of Mars. A DARE platform would carry 5 surface stationsweighing 5 kg each, and capable of operating for 1 year on the surface. The surface stations would rely on DeepSpace 2 11 (DS-2) and NetLander 12 heritage. They would be able to withstand high impact velocities (DS-2 was 160-220 m/s). The expected stations’ impact velocities would range from 40 to 60 m/s, depending on season. Advancedtechnologies would allow reduction in mass of the station (from about 20 kg NatLander stations), which wouldenable the deployment of many stations from a single platform. Advanced power generation and storagetechnologies would enable emplacement of the surface stations in the polar latitudes. The surface station payloadFigure 8. Tetrahedron network of surface stations at Mars (M. Knapmeyer).5American Institute of Aeronautics and Astronautics

would include a seismometer, meteorological package and soil analysis package. The mission profile would consistof the DARE platform flying to pre-selected sites and deploying the stations and then continue with a surfaceimaging mission afterwards. Preferred positions for the surface stations sites are shown on Fig. 8. This missionwould produce a map of the internal structure of Mars and validate atmospheric circulation models.By design, the DARE architecture can take advantage of the cooperative opportunities between the variousplatforms. As an example of a cooperative mission consider the Mars Sample Return assist mission scenario. In thismission scenario the DARE architecture creates a Mars-wide transportation system. The objective of the mission isto enable sample return from multiple sites on a single return vehicle. The mission profile is depicted schematicallyon Fig. 9. The mission profile is as follows: multiple rovers collect samples at different sites on the surface of Mars.The samples are placed into canisters for transfer. The DARE platform flies over the rover sites and collects thesamples in canisters. The drawing in Fig. 9 shows just one of the many possible mechanisms for the transfer fromthe surface to the overflying DARE platform. In this approach the sample canister is shot into the atmospheretrailing a micro-tether. The micro-tether is snatched by the BGS system of the DARE platform and the samplecanister is thus collected. There could be other approaches to the sample canister transfer. If the collection attemptfails, the DARE platform will attempt collection on a subsequent overflight. After the sample is transferred to theDARE platform, the platform flies to the site of the sample return vehicle, where the sample canister is dropped. Thecanister is then collected by a canister collection rover and is placed onto the sample return vehicle. The sequencerepeats until several samples from different sites on the planet are placed onto the sample return vehicle. Thesample return vehicle then takes off and returns the samples to Earth. This mission would enable return of samplesfrom several distinct sites on a single return vehicle, maximizing mission scientific return.Figure 9. Mars Sample Return Assist mission profileIV. Balloon Guidance System (BGS) TechnologyThe ability to guide DARE platforms to the sites of interest or to avoid high topography is crucial to the DAREarchitecture. The BGS technology consists of an aerodynamic surface (wing) hanging below the balloon on a verylong (several km) tether. Several designs are being investigated: a Single-wing BGS (see Fig. 10) and a Dual-wingBGS that can be seen on Fig. 11. The difference in winds at different altitudes in the atmosphere creates a relativewind at the altitude of the wing (stronger winds are usually found at higher altitudes on Mars). An example of awind profile at Mars is shown on Fig. 12. The relative wind creates the lifting force that is directed sideways. The6American Institute of Aeronautics and Astronautics

Figure 10. Single-wing BGS Figure 11. Dual-wing BGSlifting force depends on the size of the wing, its aerodynamic properties, the density of the atmosphere and on thestrength of the relative wind (which ultimately depends on the length of the tether).The magnitude and the direction of the lifting force can be controlled by changing the roll angle and the angleof-attackof the wing. The horizontal component of the total force produced by the wing can be used to change thepath of a balloon in the winds. Preliminary numerical analysis indicates that a guidance system with a relativelysmall (1 m 2 ) BGS wing is capable of moving a conceptual DARE platform with the velocity of the order of 1 m/s inthe crosswind direction.The aerodynamic surface of the BGS will be operating at low Reynolds numbers (~1000). For these lowReynolds numbers, drag coefficients are significantly higher and maximum lift coefficients are a little lower than forhigher Reynolds number operation. The reduced Lift-to-Drag ratio is quite a challenge for systems, which mustgenerate their lift aerodynamically whileproviding power to overcome the drag.However, for our system, the weight issupported by buoyancy. The “lift” from thewing is directed close to horizontal andpredominantly across the flight path of theballoon. The drag acts mostly to slow downthe balloon, and is relatively unimportant tothe operation of our system.One of the advantages of the proposedBGS is that it does not require power forpropulsion. A small amount of power (~1 Won the average) is needed only for radiocommunications between the BGS and thegondola and to adjust the control surfaces ofthe guidance system, by means of anactuator, once a day, or even less frequently,depending on the path control objectives.Alternative approaches to balloon pathguidance employing, for example, an enginedriven propeller, would require significantlyFigure 12. Mars wind profilehigher levels of continuous power input(1000s of Watts).V. Other Key TechnologiesThe reduction in mass of the floating system is the main driver behind the identification of other keytechnologies. If incorporating an advanced technology results in a reduction of mass of the DARE platform, morescientific payload can be carried and more science can be accomplished. Key technologies that reduce the mass of7American Institute of Aeronautics and Astronautics

the system include advanced balloon materials, advanced structural materials, inflation hardware technologies,advanced solar power generation technologies, lightweight science sensors, etc.Apart from the mass reducing technologies there are other key technologies relevant to the planetary ballooningmission. These include Entry, Descent and Inflation (EDI) technology; night power generation, balloon guidancetechnology, autonomous navigation and control technologies - and such.There are other technologies, not listed above, that are necessary or enhancing, but we do not discuss them hereas they will come about one way or the other from commercial, defense or space systems developments.In our analysis of key technologies, we consider three technology time horizons: Current, which corresponds to aTechnology Readiness Level (TRL) of 8-9 and requires up to 3 years to get to successful mission operation; Nearterm,which corresponds to a TRL of 3-6 and requires 3 to 10 years; and Far-term, which corresponds to a TRL of 1to 3 and requires more than 10 years. In this section we focus on the Near-term technology horizon. Thesetechnologies have been assumed in preliminary system trades and have been used to develop an example designdiscussed in Section VI. Below we discuss advanced balloon envelope material and EDI technologies.A. Advanced Balloon Envelope MaterialsMaximizing the useful mass of theballoon’s scientific payload requiresreducing the masses of all the othercomponents. Advanced lightweight andstrong materials need to be developed toreduce masses of parachutes, heatshields, balloon envelope and othercomponents that take up a greatproportion of the mass delivered to Mars.Of special importance are thematerials for balloon envelopes.Reducing the surface density of theballoon film by using compositematerials (such as Mylar-Kevlar scrim -Polyethylene 5 ) can lead to substantialincreases in useful payload mass. Thehigh atmospheric altitudes (10-20 km)and heavy payloads (~100 kg)envisioned in our concept require largeballoons (20-70 m in diameter). The Figure 13. Composite material for a Mars Balloon envelopegreater stress experienced by theenvelopes of these large balloons will also be supported by the advanced materials.For the DARE platforms using Near-term technology we are considering a composite material consisting of 1-micron Mylar, a 38-Denier PBO thread and a 3-micron PE film. The surface density of this material is estimated at0.012 kg/m 2 . A similar, but thicker, material was proposed in previous Mars balloon study 5 . The need for acomposite material arises from the fact that it is impossible to find all the needed mechanical properties in a singlematerial. Hence, in this proposed composite material Mylar provides substrate stiffness, Polyethylene providesfracture toughness and pinhole resistance, while the scrim provides high-strength at low surface mass. Fig. 13 showsa sample of a composite envelope material.B. Entry, Descent and InflationEntry, Descent and Inflation (EDI) are the sequence of events that occur over a short time-span following thearrival of the spacecraft carrying the DARE platform to a planet and before the platform achieves its floating altitudein the atmosphere of a planet. Fig. 14 on the next page illustrates one concept for a sequence of events that takeplace during EDI.The entry and descent phases would be very similar to previous Mars lander missions. A blunt conical heatshield reduces the enormous interplanetary velocity and absorbs the heat generated by passing through theatmosphere. The aft cover is released and a supersonic parachute is deployed at a Mach number of ~2. This slowsthe descent rate significantly into the low subsonic range to the point the internal components no longer needprotection from the ram air and the heat shield can be dropped. As the system approaches terminal velocity on theparachute, the inflation phase commences. The gondola and attached equipment are lowered, and the weight is used8American Institute of Aeronautics and Astronautics

Figure 14.DARE EDI sequenceto stretch out the balloon envelope vertically. A load member inside the balloon could absorb the shock generatedduring deployment as the envelope goes taut. The inflation equipment is mounted below the balloon, and will bejettisoned when the inflation of the balloon is complete. An inflation tube is used to carry the gas up to the initialinflation bubble towards the upper end of the balloon and could also to support the deployment shock. Once theballoon has reached a size comparable to the parachute, the parachute will be cut away. Upon completion ofinflation, the inflation hardware will be released, and the system will ascend towards an equilibrium float altitude, atwhich point the trajectory control system can be deployed. The Mars balloon EDI technology is currently beingstudied by NASA.An alternative approach, that was considered and rejected, is to land prior to inflating the balloon. Thedisadvantages of this approach are the increased mass of a landing system and the danger that the balloon can touchthe ground and be damaged during inflation by significant winds.The inflation hardware (the tanks holding the buoyant gas under pressure, valves and pipes) can have asubstantial mass – for example, in a previous Mars balloon study 5 the inflation hardware for 11.5 kg of heliumweighed 75 kg. For the Near-term technology horizon hydrogen can be used instead of helium as a buoyant gas andstored in a cryogenically cooled container before the start of the inflation sequence. Warming up cold hydrogen forinflation requires additional equipment, however the estimated savings in mass are still substantial – we estimatethat the mass of the inflation hardware for 8.5 kg of hydrogen for Near-term DARE mission can be reduced to 14 kg.VI. Preliminary System Trades and Example DesignThis section describes preliminary system trades and an example design of the DARE platform. A SurfaceNetwork Emplacement mission scenario was chosen to focus the design effort and Near term technology horizonwas used in the example design.To limit the parameter space for the system trades analysis we fix the mass of the entry vehicle to 340 kg. Thisentry vehicle mass is based on the previous Mars balloon mission studiess 3,5 and assumes Delta 7326 launch rocketwith a 616 kg launch and trans-Mars injection capability. No advances in launcher and carrier spacecraft9American Institute of Aeronautics and Astronautics

Figure 15.DARE flight system definitions and massestechnologies are assumed. We further assume that the EDI hardware and the balloon flight system constitute 200 kgand 140 kg of the entry vehicle mass, respectively. Fig. 15 illustrates the mass budget. Using a more capable launchvehicle would increase the balloon flight system mass.A. Preliminary System TradesFig. 16 shows the payload mass and the balloon radius of the DARE platform as a function of the altitude aboveFigure 16. DARE platform payload mass and 10balloon radius as a function of height aboveAmerican Institute of Aeronautics and Astronautics

Mars for the three technology horizons and for a single southern summer season on Mars. The height is measuredabove the Mars areoid, which is a reference level similar to the sea level on Earth. Solid lines represent payloadmasses (left axis), while the dashed lines refer to the balloon radii (right axis). For the Far-term technology thepayload mass does not change with altitude, since it was assumed that the balloon material is practically weightless.Hence, as the balloon size increases with higher float altitude, this does not require any additional mass. Theassumption about zero mass balloon demonstrates the limits on the Mars balloon payload mass.This plot illustrates the dramatic increase in the payload mass with the decreasing balloon envelope mass. For aDARE platform at 10 km altitude the payload mass increases from about 10 kg for Current technology to about 90kg for the Near-term and to 190 kg for Far-term technology horizons.DARE platforms would employ superpressure balloons that would that remain at a constant atmospheric densityaltitude. As the Martian atmospheric temperature and pressure changes seasonally and geographically, DAREplatforms would float at different altitudes. The changes in atmospheric density structure can be quite dramatic onMars due to very thin CO 2 atmosphere that partially freezes in parts of the year. The height of a constantatmospheric density level can vary by as much as 4 km across the planet (being lower at the winter pole and higherover the equator) for the same season, and by 2 km for same location throughout a year. In addition, dust storms cansignificantly warm up the atmosphere, which would reduce the altitude of a constant atmospheric density level.Since DARE is envisioned as a platform with global reach and operating for over a year, the platform needs to bedesigned for the floating altitude that ensures that the platform is sufficiently high above the topography to allow forthe BGS operation and avoid collision with the surface.The topography of the southern hemisphere of Mars is generally 2 to 3 km above the reference level. The worstcasedesign conditions for the DARE platform are operating in the southern hemisphere during a dust storm. At least3 km altitude above the topography is required for operation of the BGS. This requirement translates into height of 6km above reference level during a dust storm, which in turn corresponds to 10 km height for normal conditions.With the Near-term technology these conditions define a DARE platform with a 34.4 m balloon and capable offloating an 87 kg payloadB. Example DARE Platform DesignIn this section we discuss one example DARE platform design based on Near-term technology and the SurfaceNetwork Emplacement mission scenario. Fig. 17 shows the packaging of the entry vehicle. The aft (top) part of theentry vehicle is occupied by the parachutes and a canister the houses the balloon envelope. The gondola is shownbelow the balloon envelope canister and above the folder BGS wing. The wing separates from the gondola andunfolds, while being lowered on a winch upon ascent to the floating altitude during the EDI. Below the folded BGSis the cryogenic inflation hardware.Figure 17. DARE platform entry vehicle packaging11American Institute of Aeronautics and Astronautics

Figure 18. DARE example gondola conceptFig. 18 shows the DARE platform gondola in more detail. In this example the gondola houses the circular solarpanel at the top, five deployable surface stations positioned on the circumference of the gondola, high-resolution IRspectrometer and visible camera. The winch and stored tether of the BGS are also shown. The camera and thespectrometer would be swung into position for observations after the completion of the EDI sequence. Table 1summarizes the mass and power budget for this example DARE design.VII. SummaryA concept for a new architecture for planetaryexploration is described in the context of Marsexploration. The key elements of the architecture are:long-duration-flight autonomous balloons, balloontrajectory control, lightweight power generation andstorage, and multiple, deployable microprobes foratmosphere and surface exploration. A relatively smalland light balloon trajectory control device would enablerepositioning the platform on a global scale for in situanalysis and targeted deployment of atmospheric andsurface microprobes. Deployment of microprobes fromballoons eliminates atmospheric entry and decelerationhardware thus reducing overall mass and permittingmore science payload or more probes. The DAREarchitecture will enable low-cost, low-energy, long-termglobal exploration of the atmosphere and surface ofMars and other planets. Additional information can befoundatTable 1. DARE design mass and power budgethttp://www.gaerospace.com/projects/DARE/DARE.htmlAcknowledgmentsThis study is funded by NASA Institute for Advanced Concepts (NIAC) under the Universities Space ResearchAssociation Research Grant No.: NAS5-03110.12American Institute of Aeronautics and Astronautics

References1 Sagdeev, R. Z. et al., “Overview of the VEGA Venus balloon in situ meteorological measurements”, Science, Vol. 231, pp.1411-1422, 1986.2 Cutts, J. A. et al., “Venus Aerobot Multisonde Mission”, AIAA Balloon Tech. Conference, pp. 1-10, 1999.3 Greeley, R. et al., “The Mars Aerial Platform Mission Concept”, AIAA paper 96-0335, 1996.4 Jones, J. A. and Heun, M. K., “Montgolfier Balloon Aerobots for Planetary Atmospheres”, AIAA Paper 97-1445, 1997.5 Nock, K.T. et al., “Overview of a Mars 2001 Aerobot/Balloon System”, 12th AIAA Lighter-Than-Air TechnologyConference, San Francisco, 19976 SAIC, “Titan exploration with advanced systems”, NASA CR-173499, 1983.7 Tarrieu, C., “Status of the Mars 96 Aerostat Development”, 44th Congress of the International Astronautical Federation,paper IAF-93-Q.3.399, 19938 Haberle, R. M. et al., “Mars atmospheric dynamics as simulated by the NASA Ames General Circulation Model”, Journal ofGeophysical Research, Vol. 98, pp. 3093-3123, 1993.9 Pankine, A. et al., “Directed Aerial Robot Explorers (DARE) for Planetary Exploration”, 34th Scientific Assembly of theCommittee on Space Research, Paper PSB1-0074-02, 2002.10 Connerney, J. et al., “Magnetic Lineations in the Ancient Crust of Mars”, Science, Vol 284, Issue 5415, pp. 794-798, 199911 Smrekar, S. et al., “Deep Space 2: The Mars Microprobe Mission”, Journal of Geophysical Research, v. 104, No. E11, p.27,013, 1999.12 Counil, J-L, et al., “NetLander mission to Mars - Payload and scientific objectives”, International Astronautical Congress,52nd, Toulouse, France; 1-5 Oct. 2001.13American Institute of Aeronautics and Astronautics

8.3 Appendix 3. “Sailing the Planets: Planetary Exploration from GuidedBalloons”, presentation at the 7 th Annual NIAC Meeting, Denver, CO,October 10-11, 2005.193

SAILING THE PLANETS:PLANETARY EXPLORATION FROM GUIDEDBALLOONS7 th Annual Meeting of the NASA Institute for Advanced ConceptsDR. ALEXEY PANKINEGLOBAL AEROSPACE CORPORATIONSAILING THE PLANETS 1

MARS ROVERS ARE A GREAT SUCCESS…Spirit view from Husband Hillsummit (NASA/JPL)… but their range is very limitedSAILING THE PLANETS Columbia Hills surroundings2(NASA/JPL/MSSS)

DARE – NEW PLATFROM FOR PLANETRAYEXPLORATION• An airplane will lastfor just a few hours• Airships propulsionsystems make themprohibitively heavy• Ordinary balloons areat the mercy of thewinds• Directed Aerial RobotExplorers (DARE) -guided long-durationballoon platformsMars Express/ESA-GACSAILING THE PLANETS 3

NEW ARCHITECTURE FORPLANETARY EXPLORATIONKEY ELEMENTS:• Long-DurationPlanetary BalloonPlatforms• Balloon Flight PathGuidance• AutonomousNavigation & Control• Lightweight PowerGeneration & EnergyStorage• Miniaturized ScienceSensors• Small DeployableScience Packages• Communication RelayOrbiter (MTO)• Synergy BetweenPlatforms ComprisingArchitectureSAILING THE PLANETS 4

MARS DARE PLATFORM SCHEMATICS• Superpressure balloon (Al top, white paintbottom, D=20-70 m)• Gondola:• Science payload (~100 kg)• Power generation & energy storage• Communications• Microprobes• BGS deployment system (a winch)• Tether (5-11 km)• Balloon Guidance System (BGS)

DARE ARCHITECTURE APPLICATIONSAND EXPLORATION CAPABILITIESSAILING THE PLANETS 6

EXPLORATION CAPABILITIES• Global planetary coverage• Heavy, power-intensive payloads (90 kgand 200 W in 3 to 10 years, 170 kg and400 W >10 years)• Long flight duration: 700 days (1 Marsyear)• Targeted overflight of surface sites andprecise delivery of science probes• Proximity to surface enables highresolutionimaging, elemental, magneticand gravity surveys not possible orchallenging from orbitOlivine outcrop and DS-2 landing ellipse(NASA/JPL/ASU)• In situ atmospheric chemistry andcirculation• Landing sites reconnaissance, navigationbeacon emplacementSAILING THE PLANETS Water ice lake inside a crater on Mars 7(ESA)

FRACTIONATION OF METHANEISOTOPES IN THE ATMOSPHERE• Methane-making organismsdiscriminate betweenisotopes as they feed on aglobal reservoir of CO2• Measure the C 12 /C 13 ratio inthe methane.• If it is different from theisotope ratio in the CO 2 , itwould offer strong evidencefor a biological source.Tunable Laser Spectrometer for Atmosphericand Sub-surface gas measurements on Mars(NASA JPL)• DARE enables planetary-widesearch for surface biologicalsourcesSAILING THE PLANETS 8

SURFACE TARGETS FOR HIGH-RESOLUTION IMAGINGVery small cratersBouldersDichotomy boundaryLayers in canyon/crater wallsOrigins of the outflow channels10 cm10 kmSAILING THE PLANETS 9

EMPLACEMENT OF SURFACENETWORKS ON MARS• Single DARE platform can carry tensof mini-labs• Meteorological & seismologicalnetworksSurface labs locationsNetLander Surface Module (ESA)Mars Microprobe(NASA) as an exampleof a mini-labSAILING THE PLANETS 10

MARS SAMPLE RETURN ASSIST• Multiple rovers collect samples at different sites• Samples and transferred to Sample Return Vehicle by DAREplatform• Science results: several samples from distinct sitesSAILING THE PLANETS 11

SAILING ACROSS MARTIAN EQUATORSimulated DARE trajectory over elevation contour mapDARE trajectory over MOLA topography(NASA)• 90-day late Southern spring, 1 m/s control velocity• Objective: navigate from Southern to Northern midlatitudesSAILING THE PLANETS 12

DARE AT VENUS, TITAN, JUPITER•VENUS- Targeted overflight of surface sites and precisedelivery of geophysical probes- Wind profiles and atmospheric composition atmultiple locations•TITAN- Global measurements of winds, gas abundances,surface chemistry with probes•JUPITER- Solar-Infrared Montgolfier balloons- Sample with probes distinct regions of theatmosphere (Great Red Spot, belt/zone)SAILING THE PLANETS 13

KEY TECHNOLOGIESSAILING THE PLANETS 14

KEY TECHNOLOGIES• Three technological time horizons:Current (0-3 years, TRL 8-9), Near (3-10years, TRL 3-6), Far (beyond 10, TRL 1-3)• Advanced Balloon Materials• Balloon Guidance System (BGS)• Entry, Descent and Inflation (EDI)• Navigation & Guidance in Mars winds• Mars Balloon performance modelingSAILING THE PLANETS 15

MARS DARE BALLOON• Low-mass high-strengthenvelope material• composite material• 1-µm Mylar/38-Denier PBOthread/3- µm PE film• areal density of 0.012kg/m 2• Nano-tubes fabric infuture?• Superpressure sphere• Al top, white bottom toprevent CO 2 condensationMars balloon conceptSAILING THE PLANETS Composite Mars balloon material 16

BALLOON GUIDANCE SYSTEM (BGS)• BGS is an aerodynamic surface suspendedon a tether several km below the balloon• Tether could be Zylon fiber, 5 to 20 timesstronger than steel, by weight. 10 km longtether weighs 0.5 kgDual-wing BGS• Variation in atmospheric wind and densitywith altitude result in a sideways liftingforce• 1 m 2 BGS creates sideways control velocityof 1-2 m/s in typical Martian winds and 8km tether• BGS wing operates at low Reynoldsnumbers at Mars (~1000), lift coefficientsof 0.6-1.4• Single-wing and Dual-wing BGS designs arebeing studiedSingle-wing BGSSAILING THE PLANETS 17

ENTRY, DESCENT & INFLATION (EDI)• Parachute deploys• Inflation commences• Parachute cut-off• Inflation equipment jettisoned•Platform ascends to floating altitude•The BGS is deployedAltitude profileSAILING THE PLANETS 18

SYSTEM TRADES AND EXAMPLEDESIGNSAILING THE PLANETS 19

PAYLOAD VS. ALTITUDE• Height of atmosphericdensity levels lowerby 4 km in dustyatmosphere• DARE to float 2-3 kmabove southernhighlands in duststorm• 6 km at τ=3• M=87 kg, R=17.2 m• Altitude of 10 km atnormal conditions

ALTITUDE CHANGE AFTER PROBE RELEASE• Releasing 30 kg of probes raises altitude by 3 km• Increase in super-pressure can be relieved byventing 1 kg of gas (out of 8 kg)SAILING THE PLANETS 21

ENTRY VEHICLE• Delta 7326 launch rocket, 616 kg Mars injection capability• 340 kg Pathfinder-type entry vehicle• EDI hardware 200 kg, balloon flight system 140 kgSAILING THE PLANETS 22

BALLOON FLIGHT SYSTEMSAILING THE PLANETS 23

GONDOLA DESIGNSAILING THE PLANETS 24

SUMMARY• DARE enables revolutionary planetary explorationcapabilities at Mars and other planets• DARE addresses NASA's Mars Exploration Program(MEP) goals by returning unique measurements incritical science themesSAILING THE PLANETS 25