AIAA 97=1 447Mars 2001 Aerobot / BalloonSystem OverviewKerry T. Neck, J. Balaram & Matthew K. HeunJet Propulsion LaboratoryCalifornia Institute of TechnologyPasadena, California 91109

1. Steve SmithNASA Goddard Space Flight CenterWallops Flight FacilityWallops Island, Virginia 23337

Terry GamberLockheed Martin AeronauticsDenver, Colorado 80201

AlAA International BalloonTechnology Conference

June 3-5, 1997 / San Francisco, CA

For ~ermission to COPY or republish, contact the American institute of Aeronautics and Astronautics1801” Alexander Bell ‘Drive, Suite 500, Reston, VA 20191

Mars 200 I Aerobot/Balloon System Overview

Kerry T. Neck, J. Bahmrm & Matthew K HcunJet Propulsion Li~boratory

California Institute of TechnologyPasadena, California 91109

1, Steve SmithNASA Wallops Flight FacilityWallops Island Virginia 23337

Terry GamberLockheed Martin Aeronautics

Denver, Colorado 80201

AMEt&l

In late 1995, a study was initiated at the JetPropulsion Laboratory (JPL) of a 2001 MarsAerobot/Balloon System (MABS) MissionParticipants included NASA Goddard Space FlightCenter, Wallops Flight Facility (WFF), LockheedMartin Aeronautics (LMA), the French SpaceAgency (CNES) Toulouse Space Center, NASAAmes Research Center (ARC), and SpaceDynamics Laboratory (SDL) plus numerousindustrial partners. The purposes of the study wereto 1) determine technical feasibility of a longduration 2001 aerobot mission in the Martianatmosphere, 2) formulate a baseline concept, 3)identify pre project technology requirements and,4) develop a preliminary cost, schedule and plan.The study scope included definition andidentification of mission concept technical issuesincluding science instruments, gondola, balloonsystem design, entry vehicle and cruise spacecraftdesign, and launch vehicle performanceconsiderations.

Key constraints on the mission study were a 2001Mars launch opportunity (although 2003 and 05were also examined), a Delta launch vehicle,maximum use of Mars Surveyor Program (MSP)cruise and entry systems, the use of Mars GlobalSurveyor and MSP orbiter to relay communicationscapabilities and a 90 day mission duration. Keyassumptions of the study included 1) a gondolamass of the order of 10 kg including scienceinstruments, (plus deployable science packages), 2)“constant” density altitude, superpressure balloondesign without landing capability and, 3) cruisealtitude of 5-8 km above reference lCVCI.

The study effort concluded that the MABS missionis feasible based on conservative ilssunlp[ions onenvironmental and tcchnicid readiness providccl~iirly and significant NASA investment in balloonsystcm technology is initiated.

Backmound

In the mid- 1980’s, the French and the Sovietsbegan studying a Martian Aerostat mission for the1994 Mars opportunity. The French were to supplythe balloon system and the Soviets were to providethe gondola and deliver the system to Mars. Thesystem design concept eventually evolved into a 6micron thick mylar balloon which would beoverpressure during the day, descending to thesurface during the night (Ref. 1). At night, theballoon would rest on the surface on a guiderope or“landing snake” suspended from the gondola. Themass of the landing snake on the surface relievednegative buoyancy so that the gondola did nottouch the ground. The mission was to last about 10days before gas leakage reduced lift and kept theballoon from ascending during the day.

As early as 1992 the joint program wasexperiencing the detrimental effects of the collapseof the Soviet Union. The immediate impact was aslip of the launch to at least 1996. Thisprogrammatic uncertainty also resulted in fundingand related technical difficulties within the Frenchprogram. By early 1995 it was clear the missionwould need to be delayed again, this time to 1998.The French continued a scaled-back developmentprogram which had significant successes but alsohad some nagging balloon deployment failures.The shift of the mission arrival date to 1998 meantthat the mission would be arriving in the Northernhemisphere in winter when high surface winds(>30 n]/s) were to be expected. The implications ofthe high winds meant that a balloon whichdescended to the surface on a landing snake couldbc destroyed if the drag on the snake became toohigh. Eventually, duc to a combination ofprogrammatic and related technical problems, theprogram was cancclcd.

In 1994 iimthcr Milrs biilloon” concept called t heMilrs Aerial Pliltforn] (MAP) mission Wiis proposedunder the NASA Discovery Program (Ref. 2). ‘t%c

CopyrightO1’Y)7 by IIIC Anwrimn Instimw nf Aermmuti~s md Astronmatim, Inc. ‘lk U. S. (kmmnul hm t roydty-lru Iim)se to exercise Jll rightsumkr lhe mpyrigh! dnimul tmmin Ii)r Ckwernnwn[ purposes. All ntlwr nghls we reserved by the a,pyrighl owner.

sys(cm design, all systcm clcmcnts must bcconsidcrccl and the rcquircrncnts analyzccl. As anexample, [hc driving design requirement on thestrcnglh of the bitlloon cnvclopc material may notbc the Icvcl of suppcrprcssure but instead may bcdeployment and inflation (D&I) forces dependingon D&I method chosen. Another example relatedto materii~l choice is how sterilization, packing orstorage methods can drive envelope design. Theballoon is a system with a capital “S” and all thefactors influencing its performance must beconsidered to insure a successful design.

System and subsystem alternatives were identifiedand the more attractive options were developed andcompared in order to select a baseline design.Table 1 illustrates a few of the subsystem designalternatives considered in the study.

Table 1: Subsystem Design Options

Cruise and Entrv Svstems:Cruise Stage: MSP-Based, VCP-Based, Adv. Tech.Entry Vehicle: MSP ’98, PathfinderParachutes: MSP ’98 Adv. Tech

Balloon Svste~Envelope Material : Mylar C, Nylon 6, CompositesScrim Material: Kevlar, Dynema, PBOThermo Optical Surface: Transparent, White, & AlBalloon Geometry: Cylindrical (AR = 3-5), SphereBuoyant Gas: Helium, HydrogenDeployment & Inflation : Top and Bottom Tanks

w/Top Bubble, Bottom Tanks with BottomBubble

Reefing : Collar Straps, SleeveEnvelope Storage: Folded, Wound, RolledPayload: 10-20 Kg

GOndo la Systems:System Arch.: Dedicated Prcsr, Shared ComputerComputer : COTS, MCMThermal Stability: RHU, Resistive Heaters, Cold

ElectronicsStructure/Thermal : Mars Rover - Based, NewCommunications: New Millennium - Based UHF,

MSP ’98, Martian Aerostat, New UHFSun Sensing: APS Camera, Digital Sun SensorAltitude Sensing: Radar, LaserVelocity Sensing: Imaging (Day), VLBI, Doppler

Radar, LIDAR

After sclccling a reference systcm design whichmeet the mission and science requirements, theildviin~cd tcchrrology development needs werecsti~bl ishcd. An overall progriml was constructedwhich factor-cd in both the prc-project activities,such its balloon technology ctcvclopmcnt anddemonstrations, and the prc)jcct implementation

tasks such as detailed design, fabrication systcmintegration and testing. Two cost estimates weredcvclopcd. The first estimated was performed bythe study team members which was called a “grassroots” estimate. The second was done by the JPLIndependent Cost Models Estimation (ICME)based on historical cost performance of similarsystems with new ways of doing business factoredinto the result.

En vironmental Model~

The environment has a first-order impact onballoon design and flight dynamics. Obviousexamples of the effect of environment are (a) theglobal atmospheric circulation which dictates theballoon ground track and (b) the atmosphericdensity which determines the required balloon size.Less obvious is atmospheric radiation (both solarand infrared) which helps determine balloonenvelope strength requirements. Because there isconsiderable temporal and spatial variability in theMartian environment, an essential aspect of thefeasibility study was to determine the range ofenvironmental parameters and the worst-caseconditions for balloon design. Figure 1 describesthe basic behavior of a superpressure balloon atMars given different atmospheric temperatures andpressures. Later a similar chart will be shownwhich relates specific environmental parametersand conditions to balloon behavior.

Several environmental factors were evaluated fortheir impact on balloon design: dust, surfacethermal inertia, surface albedo, topography,atmospheric surface pressure, and the time of yearof the flight. Each of these factors and theirinteractions are discussed below.

12!M

Mars is famous for its dust storms which vary inextent from localized events to planet-encirclingstorms. High levels of atmospheric dust increasethe opacity of the atmosphere in both solar andthermal radiation wavelengths. Dust moderates theimpact of environmental radiation by increasing theoptical depth of the atmosphere. From the point ofview of balloon design, high dust optical depth isfavorable because it reduces the magnitude of thediurnal variation in atmospheric radiation. Withmoderated radiation, the temperature of the balloongas undergoes ICSS diurnal variation therebyreducing the balloon material strengthrcquircmcnts. The highest balloon skin strengthrcquircmcnts come from an optically clearatmosphere with no dust. Thus, the worst case forballoon design is an opti~id depth of zero. The (iuststorms iit Mi~rs arc unpredictable but follow n

3

[hc thin Martiim atrnosphcrc makes entry difficult.Ahnosphcric entry at a low locution is desired,

Third, balloons flying at reasonable ahitudcs (say,6 km) may Ily above high plateaus (at, say, 5 km)and bc very CIOSC to the surface of the planet. Theballoon will scc the full diurnal variation of thesurface temperature because the thin layer ofatmosphere between the surface and the balloondoes little to attenuate the thermal radiativeenvironment. A problem may occur when flyingover areas of high topography and low thermalinertia, such as Alba Patera (5 km, 40 deg N, 110deg. W). The close proximity to the surfaceaccentuates the diurnal variations in balloon gastemperature.

Solar Flux

The eccentricity of the Martian orbit means that thesolar energy received by the planet is more intenseduring the southern hemisphere spring andsummer. Thus, the southern summer provides themost severe diurnal radiation variations, a first-order effect for balloon design. A secondary effectof the variation of solar flux is atmospheric surfacepressure.

Surface Pressure

During the southern hemisphere summer (270° S L,c 3600), Mars is closer to the sun than during thenorthern hemisphere summer (90° < L. < 1800).The higher solar heat flux associated with theshorter Sun-Mars range during southernhemisphere summer results in additionalsublimation of the southern polar ice cap and moregas in the atmosphere. Thus, the planet-widesurface pressure is higher during southern summerthan northern summer.

The seasonal atmospheric pressure variation affectsballoon and mission design. Higher atmosphericpressure leads to higher float altitudes, all otherfactors being equal. The result can be up to 2 km ofseasonal float altitude variation for a givensuperpressure balloon due to seasonal atmosphericpressure variations alone. A long duration missionmust consider the effect of variable surfacepressure on float altitude. In terms of balloondesign for a long-duration mission, low surfacepressure seasons require the largest balloons andrepresent the worst case for balloon design.

There are both seasonal and diurnal time scales toi~tmosphcric pressure variation. On diurnal timescales, Mars’ pressure profile varies only slightlybut significantly, Icss than *IO% pcr day. The

diurnal pressure variations result from the passageof weather systems.

Atmosr)hcric Models

There are two atmospheric models that were usedto assist the balloon and mission design effortsduring the MABS study, a Boundary Layer Model(BLM), and a General Circulation Model (GCM).Both models and their uses are described in thesections that follow.

Mars Boundarv Laver Model ( BLM) . The MarsBLM estimates environmental conditions at onelatitude, for one Martian day on a variable, closely-spaced altitude grid (5 tn/division near the surface,250 m/division at 10 km). Inputs and outputs aresummarized in Table 2. The BLM outputs aregiven on time grid that has 24 points per Martianday.

Table 2 BLM Inputs and OutputsInputs I Outputs (24/day)ODtical DeDth, 7 ]T, P

=

Surface Pressure, P(, Solar Zenith Angle

As stated above, the MABS effort was a feasibilityassessment for a Mrtrs balloon mission. Worst-caseenvironmental conditions were selected for thedesign atmospheres. The BLM affords theopportunity to evaluate several sets of inputparameters, and, in the process, to determine the setof atmospheric parameters that provides the mostdifficult environment for balloon flight.

Selection of BLM @u& Selection of the inputparameters to the BLM was a major aspect of theenvironmental work during this study. We wereguided by experience with balloon design, previousstudies, and engineering judgment. Because theballoon was expected to have planetwide coverage,we examined all areas of the planet. Because ourdesign goal was a 90-day mission, a large portionof the Martian year was considered, starting withthe earliest possible arrival date and ending withthe 90 days beyond the latest possible arrival date.

Wc generated design environments for each regionof the planet: Northern Design Case (NDC),Equatorial Design Case (EDC), Southern DesignCase (SDC), and Global Design Case (GDC). TheNDC, EDC, and SDC were taken to bc morespecific in Iongiludc, topography, and time of yearthan the GDC. Table 3 summarizes some of theparanwtws for ~i~~h design cils~. Tbc li][ittld~ ri}ngc

5

Table 3 BLM Design rasc pitramctcrs.Latitude L. Topography

NDC 30°< N Lnt <60° 225° <L, <360° z<5kmEDC -30°< N Lat e 30° 225° <L. <360° z<5kmSDC -60°< N Lat <-30° 225° <L, <360° z<5kmGDC -60°< N Lat <60° 0°< L, <360° z<8km

Table 4. Northern Design Case (NDC)

Location TI albedo K = wTI T o p oW Longitude N Latitude [S1 units] [-] [ 1000/S1] [km]

Lowest TI 113 41 92 0.298 3.23 4.5Highest albedo 109 43 117 0.315 2.69 4.5Highest K 115 39 92 0.303 3.28 5.0BLM Inputs 80 0.32 0.0,5.0

Thermal Inertia and Albedo BLM inputs: 30° c N Lat c 60°, Alba Patera.

Table 5. Equatorial Design Case (EDC)

[ Location TI albedo K=a/TI Topo 1W Longitude N Latitude [S1 units] [-] [1 OOOISI] [ k m ]

Lowest TI 99 -3 84 0.288 3.44 7.0Highest albedo 99 3 109 0.355 3.26 6.5Highest K 97 1 92 0.345 3.75 6.5BLM Inputs 70 0.36 0.0,5.0

Thermal Inertia and Albedo BLM inputs: -30° c N Lat c 30°, Topography e 8 km, Just east of Olympus,Arsia, Pavonis, and Ascratus Mons

Table 6. Southern Design Case (SDC)

ELowest TI

B~M In utsThermal Inertia anc

Highest albedoHighest K

Location TI albedo K=a/TI Topo IW Longitude N Latitude [S1 units] [-] [ 1000/S1] [M]93 -39 176 0.213 1,21 5.0221 -45 176 0.268 1.52 4.6295 -59 234 0.333 1.42 -0.5221 -45 176 0.267 1.53 4.6

160 0.28 0.0,5.0Mbedo BLM inputs., - 60° c N Lat <-30°, Cimmeria

Table 7. Global Design Case (GDC)

Location TI albcdo K=a/TI TopoW Longitude N Latitude [S1 units] [-] [ 1000/s1] [km]

Lowest TI 99 -3 84 0.288 3.44 7.0Highest albedo 99 3 109 0.355 3.26 6.5Highest K 97 I 92 0.345 3.75 6.5BLM Inputs 70 0.36 0.0,8.0

Thermal Inertia and Albcdo BLM inputs:, -60°< N Lat c 60°, Topography< 8 km, Cimmcria - Just east ofOlympus, Arsia, Pavonis, and Ascratus Mons

Ed loon FM S“ ~1 Vslcnl

The pert’ormarwc of supcrprcssurc (SP) ball(wnshave been well documented. Balloons of dw sizerequired for a Mars mission have been successfullyflown in excess of 100 days. However, thematerial used in these balloons, primarily mylar,had very high areal densities, problems with pin-holing and fracture, and were not designed foratmospheric entry deployment and inflation.Experience with scaling up of these balloons tolarger designs highlighted many inherent materialand fabrication problems. Application of othermaterials such as Nylon-6, while alleviating someof the problems but not all, had very limitedsuccessful demonstration. In addition, neither ofthe two materials could meet the design strength

gas inlcgrity, and geometry, 4) accessories such asfillings, rcinforccmcnts, inflation, etc., 5)fabrication such as tolcranccs, reliability,packaging, sterilization and 6) load introductionsuch as payload attachment and shock attenuationduring the D&I process. While the spherical shapeis the best from a massh.tnit area and macro stressstate, it presents increased difficulties during thedeployment and inflation process. The cylindricalballoon, which offers advantages for ease ofreefing, load introduction during the D&I,improved fabrication advantages such as ease ofmaintaining tolerances, it results in a larger andheavier balloon and stability problems duringinflation of the balloon during descent fromaltitude. A spherical design was chosen because ofmass limitations and stability concerns during the

wK

3003Inu- 200coCcga 100wn3W o

L

W“’--’”s”l—C’E’R—

If 1

DAY

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Figure 2. Mars superprcssure balloon response to its environment

requirements, limitations on areal density and still deployment and inflation process whichbe successfully deployed and inflated. It was dcscribcd later.decided to take advantage of some of the rcccntcomposite film work that was currently beingpursued by NASA.

~. There are many structural designconsiderations which must be accounted for in lhcdesign of a balloon structure, the shape, wbclhcrspherical, cylindrical or “natural shape” cacb haveadvantages and disadvantages. Factors included: I )macro and Iocalizcd stress distributions such asload introduction, 2) Ihc toti~l n~itss of COIIIpOIWIIIS

and their distribution, 3) scam design for slrcngth,

will be

Envelope Material. The identification of anadequate material is probably the most criticalfactor in the design of a balloon. This has beendemonstrated many times over the years inballooning. One cannot consider only a material’sstrength or optical properties, although both arevery important piwamctcrs. Otlcn these can bcncga[ed by poor performance in other parameterssuch as lack of elongation or very poor or noIrac[urc tougbncss, [hereby very intolerant ofdCt’CCIS which itll balloons hilvc. Pursuit of

9

Pc? JymcnLI( Dcployrncnt consithmtions inclttdcd:I ) packaging of the balloon such as density, foldingIncthod, Ctc., 2) mc[hod of deployment such asgravity or mcchttnical feed, and location within theoverall flight train, 3) container design such astoroidal, cylinder, etc., 4) reefing method ifrequired, 5) sequencing of the events, and 6) theoverall loading on the balloon envelope andgondola from the deployment process andparachute opening shock.

A survey was performed of air launched balloonsystems to establish baselines for a possible design.Information was compiled on the various systemswhich included the successful Soviet/FrenchVEGA and the U.S. Off-Board Jammer System(OBJS) as well as others. Of particular help wasthe detailed amount of work , data and assistanceprovided by CNESII’otrlouse concerning their Mars’96 development efforts. Although the Mars ’96Martian Aerostat system encountered difficultiesduring the two high altitude drop-and-deploy tests,the information and lessons learned identified keystability problems. A stability model wasdeveloped for use in determining balloon geometryand payload/spacecraft mass distributions duringthe deployment process. As a result of stabilityconcerns and mass distributions, it was decided thatthe balloon should be suspended between theinflation hardware, gondola and heat shield.

Inflation. The inflation system design is influencedby: 1) the location of the system whether at the topor base of the balloon, 2) gas selection, 3) inflationrates, 4) mass of total system and distributionthereof, 5) stability and inflation/ballooninteraction, 6) temperatures, 7) tankage and 8)sequencing. Of key concern was whether inflationcould be completed before impacting the surface.The time allowed for deployment and inflationbeing determined by the time it takes for the systemto slow to an acceptable dynamic pressure underthe main parachute which results in a stable systemand the time for gas filling which is dependent onthe amount of gas and the rate it must be injectedinto the balloon. Although a top down inflationsimilar to the VEGA and Mars ’96 designs washighly desired, stability analyses coupled withother systcm masses and physical geometry led tothe placcmcnt of the inflation hardware at the baseof the balloon, similar to the OBJS, The heliuminflation gas, was injected into the base of theballoon through a sonic diffuser. The gasprocccded up through the center of the balloonthrough a central load carrying intlation tube untilwhere the gas exited into the top of lhc balloon.

Biisclinc DC sign

The baseline Mars balloon systcm is dcscribcd infigure 4. below.

VOLUME: 10,500 m3OIAMETER: 2? mBALLOON MASS S5 kgOAS MASS; 12 kgPAYLOAD 15 kgFLOAT ALTITUOE: 6.S kmOAYTIME AR 240 PaNIGHT AP 20 PC

MTRL 3.5pm MYLAR/ SS OENIER KEVLAR / 6 pm SF-272AREAL OENSIN: 20,0 @m2MATRL. STRENGTH: 2S00 NfmSTRENGTH FACTOR OF SAFETW 1.5COATINGS: TOP IS ALUMINI’2E0 (70 PREVENT C02

CONDENSATION), B0770M WWTE

Figure 4. Baseline balloon design

Further details of the design studies and results arecontained in (Ref. 6)

Gondola and Science Instruments

Go dola Design Ob.n iective~

Earlier Mars balloon concepts did not incorporateany autonomy for on-board decision making (Ref.1,2). Even in the more advanced French concept,all science data was timer driven with localizationof the balloon only possible after data acquisitionusing earth-based analysis of images, crude sun-sensor data, and Doppler analysis of orbiter relay ofballoon communications. Global position accuracywas no better than about 30 km and available atthis resolution only when Sun and Dopplermeasurements were simultaneously available. Thisresulted in a low “yield” of science data since dataacquisition was not tied to specific science targets.When combined with the inherent limitations onthe bandwidth of the communication link from theMars balloon to Earth, the overall science returnwas severely compromised. In this study, weaimed to incorporate the technology for on-board,high-precision knowledge of the balloon positionand orientation which when used on-board forscience sequencing, and balloon task sequencing,promised to dramatically improve the overallscience quality and data return from a Mars aerobotmission.

The specific study objective was to develop aballoon gondola concept capable of maximizing thescicncc return from a 90-day balloon acrobotmission to Mars. Concepts and options relating toMars Scicncc Instrurncnts, Mission SequencingConcepts, Planet-Wide Position Determination,Earth Communication, Power Generation, Storageand Management, [ntcgratcd Structure and

II

In addition m the sensor derived cslimaws of tbcacrobot position it was also desired to bc able mpredict the position of the acrobot into the future.The accuracy of these predictions would bcdependent on the vcr(ical balloon performancemodel and the planetary Global CirculationModels. A l-day prediction accuracy requirementfor Altitude, Latitude and Longitude wasrespectively 100 m, 10 km and 100 km. Accuracyover a week was desired to be 200 m, 200 km and500 km and a 30-day accuracy of 500 m, 500 kmand 2000 km respectively in Altitude, Latitude andLongitude.

In addition a number of additional internal/derivedrequirements were also developed:

. Internal temperature ranges from -40”C to +40”C

. Tilt measurement better than +/- 5 minutes ofarc.

● Altitude measurements better than +/- 1 Yo.● Sun Elevation/Azimuth measurements better than

+/- 30 minutes of arc.. Data/Computer system reset times less than 60s.. Data/Computer system sleep-mode wake-up time

less than 3 sec

POWE

Cj

SUN

R REGI

(OkE LC

iMERA.

● On-board clock accuracy better than 20s.● Short term stab~l~fy of oscillator for one-way

Doppler of I .. BalhW/Probe drop activation time Icss than 60s.

Figure 5 is a perspective conceptual view of theMars Aerobot gondola. Figure 6 is a functionalblock diagram for the Mars Aerobot gondola.

Gondola Science Imarzing

As science imaging of the Martian terrain is one ofthe most important objectives of a Mars aerobotsystem, we now discuss some of the issues relatingto data bandwidth and image quality.

The typical maximum data return is 480 Mbit for amission at 40 deg latitude. Assuming that 90% ofthis data is given over to science images, indicatesa total of 430 Mbits of science image data return.Compression of each 1000X1OOO 8 bit image by afactor of 10 gives 0.8 Mbits/image for a totalreturn of 540 images/sol. The balloon’s ground-track motion during an 8 hour period of day-lightat an average speed of 60 m/s is 1728 km, and for a

h / - - U H F /d/TEN!4A

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CAMERA. OBLIOUE –-’”- “\, ,‘\ “. . .(.1}~1 l~~L j+LT[hlETcr{

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Figure 5. Gondola Concept

13

Another tcchniquc for motion blur compcnxrtion isto time the image data readout in synchrony withthe apparent ground-track velocity. This is possibleonly if the cictcctor axes aligns with the ground-track motion, a condition that will occur once every72 km for the data set discussed here. The dataalso indicates that the exposure time for amaximum l-pixel blur would be 3.3 ms or faster.

To accommodate the imaging needs from a Marsballoon platform within the imaging context wehave discussed, it is useful to define the notion of aTransect Image Set consisting of 1 context framingimage, 10 successive overlapping or non-overlapping high-resolution images, and highlyencoded results of interest operators/filters appliedto up-to 40 additional high resolution images. Sucha data set provides redundancy in case ofunavoidable motion blur, a fully or partiallyconnected “ground-truth” high resolution imageswath 200m wide of length 2 km inside a single 10km framing image. and ultra-compressed usefulscience/mission data for remaining 8 km of thepath. This science data sequence is illustrated inFigure 7.

Scicm S’cauu Cone@

On-board state determination involves processingsensor data to obtain knowledge of planetaryposition in the form of latitude and longitude aswell as distance vectors to specific targets. Itprovides information on orientation such as the tiltof the balloon platform with respect to vertical, andits pointing and line-of-sights. Furthermore, statedetermination allows computation of rateinformation such as the ground-track velocity, theclimb rate, balloon gondola swing rate, androtational speed of roll motion about the balloonsvertical axis. Knowledge of this state informationallows the Mars aerobot to optimally execute itson-board science and engineering sequences. Forexample, position information could be used toactivate nadir pointed cameras as the aerobot over-flies a science target site. Accurate relative positioninformation with respect to a target and knowledgeof the balloon’s orientation could be used tomaximize oblique camera coverage of the target.To perform useful science with the specificityrequired by planetary scientists, on-board positionaccuracy better than 10 km would be desired withhigher, target-relative accuracies of 1-2 km as theultimate goal. Rate information on the balloonsangular motion could be used to control the exact

GONDOLA SCIENCE DATA ACQUISITION EXAMPLESAMPLING OF HUNDREDS OF SCIENCE SITES

o

NIGHT

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THE RMAL-IR HI-RES IMAGESPECTRA

200 m

SCIENCETARGETSITE

NEAR-IRIMAGE4 m/pixel

b

why

20 In/pixel20 km

Figure 7, !%icncc scqucncc concept

]5

rcquircmcnts were estimated to bc 32.2 W-hr andnight-time needs were estimated to bc 7.3 W-hr.

9n -board Stiltc Es(ima(iort

On-board state estimation consists of determiningthe attitude of the gondola, as well as the positionand height of the acrobot.

One quantity that needs to be determined is the tiltof the gondola with respect to the local vertical.This data is needed to support accuratemeasurement of the sun/star/moon elevation angleswhich are needed for planetary position estimation.The tilt estimates are also used to distinguishbetween translation and rotation effects when usingframe-to-frame image registration based methodsof ground-track motion determination. Sensors thatprovide this information are clinometers under nearsteady state conditions, and accelerometers andgyroscopes in cases of significant pendulum andtranslation dynamics.

The other attitude measurement is thedetermination of the gondola angle with respect totrue north. This data is useful in predicting theimage rotation between successive frames whenperforming frame-to-frame image based motiondetermination. The absolute rotation angle is alsoneeded to allow integration of translationaccelerations and velocities obtained from theground-track imaging or inertial sensors. Sensorsthat provided this information include a Roll rategyro for short term, moderate accuracy deltarotations, inertial estimator using the full 6 gyroand accelerometer measurements for short andmedium term, high accuracy delta rotation, Sunand Phobos azimuth angle measurements for day-time absolute rotation, and Phobos and Bright-Starazimuth angle for night-time absolute rotation.

Determination of the height of the gondola abovethe ground is needed to understand image scalewhen interpreting science images, and whenperforming frame-to-frame motion determinationor landmark position determination. The data isalso useful in supporting emergency ballast drops.Sensors that provide this data include Iascr rangersas well as radar altimeters.

The ground-track velocity is needed to undcrstttndthe wind patterns and the long-distance balloontrajectory. The velocity can be integrated for high-accuracy knowledge of ground-track, as WCII as forcompensating for measurements of orbitcr-to-gondola Doppler measurements. Sensors toprovide this information include inertial sensors,frame-to-frame image based comparison mc(hods.Doppler radar, and cclcstial position diffcrcncing -

especially at nigh[ when high accuracy positionestimation using stadmoon fixes is possible.

Determination of the gondola latitude andIongitudc is the primary means of supporting state-driven science sequencing. In addition this data isused to predict the location of the sun, moon andorbiter positions when performing celestial orradio-metric sensing. The sensing approaches thatcan achieve this include the low accuracy methodof detecting terminator crossing, moderate day-time accuracy from successive Sun elevationmeasurements, higher day-time accuracy fromsuccessive Phobos/Sun elevation measurements,moderate/High nighttime accuracy fromsuccessive Phobos and/or Bright Star elevationsensing, moderate short and medium accuracy byinertial estimation (6-DOF rate and gyro), moderateaccuracy information from radio metric data e.g.signal acquisition/loss or Doppler profiles of anorbiter signal, and very high target-relativeaccuracy by measuring deviation of unique, clearlydiscriminated landmark features in nadir andoblique low-resolution images (e.g. craters) fromon-board map-based predicted values.

For the Mars Aerobot, daytime on-board positionand velocity would be obtained by a sensor strategythat consisted of the following:

. Terminator crossing twice a day.● Successive sun elevation measurements every 15

minutes● Successive Phobos measurements during periods

when Phobos is visible (approx 3 hrs), awayfrom the sun, well illuminated, and when theballoon rotation aligns field-of-view.

. 2-4 orbiter (MGS, M98) fixes using signalacquisitionfloss or Doppler

. Image frame-to-frame motion determinationevery few minutes when high accuracyknowledge of ground-track position or velocityis needed.

. Full 6-DOF inertial estimator if power budgetpermits.

. Landmark deviation based position determinationin close vicinity of target after earth validationof approach based on received image analysis.

● Height measurements concurrent with imagingand radio-metric measurements.

. Doppler radar sensor if mass/power budgetpermits.

Nighttirnc on-board position and velocity would bcobtained by a sensor strategy that consis(cd of thefollowing:

● Succcssivc Phobos elevation mcasurcrncnts whenPhohos is visible (approx 3 hrs), illuminated,

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initial thermal design using a Iumpcd thcrmi~lrnodcl, generic tcmpcraturc rcquircmcnts, andRHU’S should bc adequate. Detailed thermaloplions to bc investigated in the next study includeheat-pipes, and phase change materials. Themechanical system accommodates a ballastturrctfcontrollcr to drop multiple 600 g science-probe ballast packages.

The total mass of the thermal and structural systemwas estimated to be 3155 g. In addition the ballastpackages were assumed to total 3000 g.

Table 10 summarizes the mass and energy budgetfor the gondola system.

Table 10 Mass and energy budget

Day- Night-Item Mass Time Time

Energy Energy(g) (W-hr) (W-hr)

I Science Instruments I 4950 32.2 7.3Communication 430 2.2 4.4Position & Attitude 1325 3.0 0.9SensorsComputer and Data 745 8.6 7.9SystemPower 2600 0.8 1.6Structure, Thermal, 3155 0 0Cabling 1

Payload Total 13205 46.8 22.08I I

Ballast \30001 o 0

gondola External Interfaces

The following interfaces were considered duringthe design of the gondola:. Aeroshell Mechanical Interfaces - deployment

latches, restraints etc.● Deployment/Inflation System Data Interface. Cruise Stage Power Interface● Cruise Stage Thermal Interface. Balloon Mechanical Tether Interface● Balloon Sensors Data Interface

~n Dc~“s ““ .,. .

The following list the considerations which wereaddressed to sornc level as part of MABS gondoladesign:. Coll~l~lcrci:ll -Off-The-Shelf (COTS) parts vs

Custom components.

● [nstrurncnt computers vs General PurpcxxComputer

● AS IC/DSP vs General Purpose Computer. Position/Attitude sensitivity as function of sensor

selection and design parameters● On-board I -way Doppler localization vs On-

board Signal acquisition/loss localization● Star/Moon Tracker performance enhancement vs

Star/Moon Tracker Mass/Cost● Doppler radar vs Image frame-to-frame velocity

determination● Science Payoff vs Accuracy of Position

Determination● Imaging and TES Quality vs Optical Depth● Fast Shattering vs Motion-Compensation Optics● Neutron Spectrometer Performance vs Mounting

Location● Flat vs Tilted Array vs Mission Latitude vs

Optical Depth. RHU vs Heater vs Heat Pipes. Power vs Latitude of operation. Night-time science vs Power/Energy system mass● Integral exe-skeletal structure vs temporary

Launch, Cruise, Entry and Deploy structures.. Tether Length Sensitivity to Mass, Rotation and

Shadowing

Balloon Deliverv Svstern

Overview

The balloon delivery system consists of the entryvehicle, which contains the balloon system, and thecruise stage which delivers it to Mars. The primaryrequirements of the balloon delivery system are toprovide a controlled thermal environment in cruise,target the entry vehicle to the required entrycorridor, and to decelerate the entry vehicle so thatballoon deployment and inflation can begin. Thedelivery system for the MABS is designed for aMars direct entry from the approach hyperbolictrajectory. The entry system design is derived fromthe Mars ’98 architecture and uses the same 2.4 mdiameter blunt cone aeroshell design with a cruisestage for external mounting of avionics, solararrays and sensors. The balloon, gondola and allthe support equipment is contained in the entryaeroshell. The cruise stage is spin stabilized duringcruise and entry. Most of the spacecraft hardwareis redundant. The blow-down propulsion systemmounted on the cruise stage provides all the TCMdelta V and attitude control in cruise. The directentry draws heritage from the Discovery/Pathfinderdirect entry systcm that is scheduled to enter Marson July 4, 1997. The balloon/acroshcIl is Iaunchcclin an inverted configuration similar to Pathfinderand Milrs ’98. The cruise stage design andequipment layout is dcrivccl from an LMASpaccprobc/Discovery design to rcducc the non-

.

‘rhcrc wc cigh[ unique command and data handling(C&DH) cards (13 total with redundant cards) andeight unique power distribution cards to control theflight systcm in cruise and entry. The C&DH andpower electronics are similar to Mars ’98. Thelander single board computer is an R6000 using thesame Vx works operating system and interfaces asMars ’98, The lander software is the derived fromMars ’98 and Stardust. The aeroshell also containsa barometric switch, timers, cable cutters, and abattery to implement deployment of the parachutesand aeroshell.

The lander telecom subsystem has been upgradedto use the Small Deep Space Transponder, SDST,to save mass and reduce cost. This also eliminatesthe separate boxes for the command decoder andthe telemetry modulation units The SDST isredundant. There are redundant solid state poweramplifiers located on the cruise stage. The cruisetracking passes occur every other day and arerequired to be at least four hours in length to satisfynavigation requirements. The cruise stage has aMedium Gain Antenna that is a horn design for theprimary cruise link. The cruise stage also has aLow Gain Antenna that is used for emergencycommands if necessary.

The attitude control system is derived fromDiscovery Stardust. The flight system is spinstabilized in cruise and uses sun sensors andredundant star cameras for attitude reference. Theattitude control system is used to reduce the spinrate from 70 rpm at injection to 5 rpm for cruise.

Entrv and Descent

The flight system is oriented to the desired entryattitude and the cruise stage is jettisoned at 5minutes prior to entry, A jettisonable cruisestructure allows for a clean aerodynamic shape forentry while reducing the entry mass and ballisticcoefficient. The ballistic coefficient is 50 kg/m2with a coefficient of drag of 1.6. The heat shieldshape is based on Viking and Pathfinder for whichthere is a large aerodynamics database. The drogucparachute is deployed at an altitude of about 7kilometers altitude based on the calculated entrytrajectory. The backshell is released at about 25 safter drogue parachute deploy. The main parachuteis deployed when the backshell releases. Thedroguc pamchutc carries away the backshcll but thelwotshiclcl is retained until balloon inflation isConlplctc.

HMiEtx

The flight systcm has strong heritage fromViking/Pathfinder and Discovery Stardust. LMAbuilt the acroshclls for both of these missions and isin the process of building the Mars ’98 aeroshell.The aerodynamics for the 70° blunt cone aeroshellare well understood from Viking and Pathfinder.Mars direct entry will be demonstrated byPathfinder and Mars ’98. The disk-gap-bandparachute design is the same design as Pathfinderand Mars ’98. The cruise stage structure andelectronics are the very similar to Discovery tominimize non-recurring cost.

New Technology

The cruise stage uses the Small Deep SpaceTransponder which will be demonstrated on NewMillennium Deep Space 1 mission in 1998. Noother new technology items were assumed in thestudy but several mass reductions are possible for a2001 launch based on current technologydevelopment efforts.

Trades/Risk Assessment

The major flight system trades conducted includedredundancy, the trajectory, and the delivery systemconfiguration. Redundant lander hardware hasbeen baselined to reduce mission risk and to takeadvantage of the similar Surveyor and Discoveryhardware and fault protection. Six options wereconsidered for the Delivery System design withvarying technology level, launch mass, andheritage. Minimal new technology was used toallow development of a well understood design.The Discovery design for the cruise stage wasselected since it was mass efficient and it wasdesigned for spin stabilization.

The entry heating is well within the parameters ofprevious missions such as Pathfinder. The lowerballistic coefficient of the entry vehicle allowsparachute deployment at lower velocities and lowerloads than Mars ’98.

SU!mMIY

The Mars 2001 Aerobot/Balloon System Studyeffort, using conservative design assumptions, hasresulted in a feasible design for a long-durationmission to Mars. The kcy technology requirementsfor this mission have been identified and atechnology plan has been constructed. A baselinemission and systcm concept has been defined WCIIenough to gcncratc rough estimates of a totalmission cost. Excluding launch, cruise spacecraft,entry vchiclc and project rcscrvcs, i~ Miirs Acrobot

21

●

Rcfcrcw. .

1. Tarricu, C., “Status of the Mars 96 AcrostatDcvclopnwnt” lAF paper, Graz, Austria, October]9930

2. Grcclcy, R., et. al., “The Mars Aerial PlatformMission Concept”, AIAA paper 96-0335, 34thAerospace Sciences Meeting, Reno, NV, January1996.

3. Pollack, J.B. et al. “Simulations of the GeneralCirculation of the Martian Atmosphere. PolarProcesses”. Journal of Geophysical Research,Vo]. 95, No. B2, pp. 1447-1473, 1990.

4. Schaeffer, J. Personal communication, March1996.

5. Heun, M. K., H. M. Cathey Jr., R. Haberle,“Mars Balloon Trajectory Model for MarsGeoscience Aerobot Development”, AIAA BalloonTechnology Conference, San Francisco, California,3-5 June 1997.

6. Smith, I. S,, H. Cathey, S. Raque, M. Said and J.Simpson, The Mars 2001 Balloon Design, AIAAInternational Balloon Technology Conference, SanFrancisco, California, 3-5 June 1997.

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