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Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

A97-31331Mars 2001 Aerobot/Balloon System Overview

Kerry T. Nock, J. Balaram & Matthew K HeunJet Propulsion Laboratory

California Institute of TechnologyPasadena, California 91109

I. Steve SmithNASA Wallops Flight FacilityWallops Island Virginia 23337

Terry GamberLockheed Martin Aeronautics

Denver, Colorado 80201

Abstract

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, and4) 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 2005were 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 for relaycommunications capabilities and a 90 day missionduration. Key assumptions of the study included 1)a gondola mass on the order of 10 kg includingscience instruments, (plus deployable sciencepackages), 2) "constant" density altitude,superpressure balloon design without landingcapability, and 3) cruise altitude of 5-8 km abovereference level.

The study effort concluded that the MABS missionis feasible based on conservative assumptions onenvironmental conditions and technical readiness,provided early and significant NASA investment inballoon system technology is initiated.

Background

In the mid-1980s, the French and the Soviets beganstudying a Martian Aerostat mission for the 1994Mars opportunity. The French were to supply theballoon system and the Soviets were to provide thegondola and deliver the system to Mars. Thesystem design concept eventually evolved into a6-micron 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"anding 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 rn/s) were to be expected. The implications ofthe high winds meant that a balloon whichdescended to the surface on a landing snake couldbe destroyed if the drag on the snake became toohigh. Eventually, due to a combination ofprogrammatic and related technical problems, theprogram was canceled.

In 1994 another Mars balloon concept called theMars Aerial Platform (MAP) mission was proposedunder the NASA Discovery Program (Ref. 2). The

Copyright©1997 by the American Institute of Aeronautics and Astronautics, Inc. The U. S. Government has a royalty-free license to exercise all rightsunder the copyright claimed herein for Government purposes. All other rights are reserved by the copyright owner.

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MAP design consisted of a 12-micron, biaxialnylon 6, supperpressure balloon envelope and asmall, 7 kg gondola. The balloon was designed tofloat at a "constant" density altitude in nominalMartian environmental conditions. A two-balloonmission was proposed which included smallsurface meteorology packages on the entry vehicle.The MAP mission included a number of innovativefeatures but was criticized for the lack oftechnology readiness primarily in the balloonenvelope design.

Both the Martian Aerostat and the MAP missionconcepts have had a significant influence on theMABS design, the primary points of departurebeing balloon envelope material and design and theassumption of conservative environmental factors.

A significant advance in material design conceptswas achieved during the MABS effort. Theseadvances were due primarily to 1) the realization ofthe difficulty in finding both fracture resistance andtoughness in a single-component envelope, 2) thefocus of the NASA long-duration stratosphericsuperpressure balloon program on similar balloonenvelope material concepts, and 3) compositematerial design prototyping and testing carried outin the study. The resulting material concept hastwice the strength-to-weight ratio of previous Marsballoon envelope material candidates.

Contributions of Aerobot/Balloons toMars Exploration

The science experiment potential of a MarsAerobot/Balloon includes 1) imaging the surface ata scale and resolution relevant to future lander androver operations, 2) high resolution visible imagesand IR spectroscopy, 3) collection of data athundreds of scientifically interesting sites, 4)detection and mapping of subsurface ice and water,5) in situ measurements of atmospheric propertiesand winds, and 6) deployment of surface packagesat interesting sites for exploration.

The need for ultra-high resolution remote sensingwhich cannot be easily obtained from orbit isdriven primarily by the desire for increasing theprobability of successful Mars landings, roveroperation and sample return. After the Vikinglandings, engineers considered themselves luckythat both landers survived without damage due tolarge boulders found near the landing sites. TheMSP landers are smaller and thus susceptible todamage or adverse effects of smaller, morenumerous rocks. At this time, the landing hazardof these landers has yet to be quantified, however,it is clear that imaging data of the order of 20 cmresolution would be important to the assessment of

landing success probability and site selection. Inaddition, this class of imaging would be importantto rover operations planning and rover siteselection. At least part of this issue has beenresolved by the inclusion of descent imagingcapability for future rover missions. Unfortunately,such imaging only provides after the factinformation which can be used for post-landing (ifsuccessful) traverse operation planning and onlythen within a kilometer of the landing site.Aerobot data, on the other hand, could provide bothmedium resolution (2 m/pixel) wide area and ultra-high resolution sampled coverage which can beused to optimize and select future sites for roveroperation.

Currently there are about 80 major geologic unitsidentified from Viking data. It is not practical toexpect to be able to collect samples from all ofthese sites and return them to Earth in the nearfuture. An aerobot, with high resolution imagingand spectral instruments, could assist in the processof selecting optimum landing sites to maximize thevariety of terrains sampled. In addition, such dataenables calibration of global maps and data setstaken from orbital platforms.

Design Philosophy

The MABS design philosophy consisted of severalkey elements. Mission and "worst case"environmental requirements were identified.Environmental conditions and parameters wereresearched and analyzed as they relate to balloonsystem design. Previous Mars balloon proposalsand programs made optimistic assumptions onrange of environmental parameters expected. Forshort-duration missions, which fly over a limitedamount of Martian terrain, some of this optimism iswell placed. However, for missions which areexpected to last several months, large parts of theplanet are traversed and thus a wide range ofenvironmental conditions are experienced. Hereworst case means the worst expectedenvironmental condition during a mission. Worstcase does not mean to sum all of the adverseconditions ever seen on a global scale becausethese conditions are never experienced at the sametime at the same place. In addition, in order todevelop a system which is robust to changinglaunch opportunities it is important to consider theseasonal variations in atmospheric pressure in thedesign so that the basic design can be used for allconceivable arrival conditions.

A very important characteristic of a Mars balloonsystem design process is the tight design interplaybetween various elements. To optimize the totalsystem design, all system elements must be

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considered and the requirements analyzed. As anexample, the driving design requirement on thestrength of the balloon envelope material may notbe the level of supperpressure but instead may bedeployment and inflation (D&I) forces dependingon D&I method chosen. Another example relatedto material 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 Entry Systems:Cruise Stage: MSP-Based, VCP-Based, Adv. Tech.Entry Vehicle: MSP '98, PathfinderParachutes: MSP '98 Adv. Tech

Balloon Systems: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

Gondola 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 selecting a reference system design whichmeet the mission and science requirements, theadvanced technology development needs wereestablished. An overall program was constructedwhich factored in both the pre-project activities,such as balloon technology development anddemonstrations, and the project implementationtasks such as detailed design, fabrication system

integration and testing. Two cost estimates weredeveloped. 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.

Environmental Models

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.

Dust

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 less diurnal variation therebyreducing the balloon material strengthrequirements. The highest balloon skin strengthrequirements come from an optically clearatmosphere with no dust. Thus, the worst case forballoon design is an optical depth of zero. The duststorms at Mars are unpredictable but follow a

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general, seasonal pattern, and near-zero opticaldepths are possible.

Thermal inertia

The surface of Mars undergoes large diurnaltemperature swings which are governed, in part, bythe thermal inertia of the surface at a particularlocation. In the absence of atmosphericparticipation in the thermal radiation (optical depthof zero), the surface temperature determines themagnitude of the upwelling IR radiation in theatmosphere. Locations with high thermal inertiaundergo less-severe diurnal surface temperatureswings and provide less diurnal variation in the

Given the typical ranges of albedo and thermalinertia that exist at Mars, the thermal inertia of thesurface is usually the most significant of the twoparameters for driving the balloon design.Unfortunately, high albedo and low thermal inertiatend to correlate at Mars.

Ratio of albedo to thermal inertia. K

A parameter K was defined to be the ratio ofalbedo to thermal inertia at a given location.Because low albedo and high thermal inertia areboth favorable for balloon design, a low value of Kis desirable. Although K is a simple figure of merit,

UJDC(/}CO111QCO.QCUJD.3CO

LUQ

ENTRYI

INCREASED INCREASED BALLASTATMOSPHERIC ATMOSPHERIC DROP

PRESSURE(ISOTHERMAL

ATMOSPHERE)

TEMPERATURE(CONSTANTPRESSURE)

BALLOON DENSITY IS CONSTANT

- MASS (m) ALTERED ONLY BY BALLAST DROP OR GAS LEAK

- VOLUME (V) FIXED DUE TO NON-EXPANDING ENVELOPEBALLOON AL WA YS SEEKS CONSTANT DENSITY ALTITUDE

WHEN SUPERPRESSURED, BALLOON ALTITUDE CHANGES DUE TOATMOSPHERIC VARIATION OR BALLAST DROP ONLY

END OFMISSION

SURFACEFigure 1. Superpressure Balloon Behavior

upwelling thermal radiation. High thermal inertia isfavorable for balloon design because it decreasesthe diurnal temperature variation of the balloongas.

The SI units for thermal inertia are J m'2 s""2 K"'.

Albedo

The albedo of the Martian surface affects theamount of daytime solar radiation that is reflectedonto the balloon from below. Locations with highalbedo provide additional daytime solar loading onthe balloon and magnify the diurnal temperaturevariation for the balloon gas. Thus, low albedo isfavorable for the balloon design.

locations with high values of K should be closelyevaluated for their potential detrimental impact onballoon performance.

Topography

Traditionally, the 6.1 mbar pressure surface derivedfrom radio occultation data is defined to be the 0km point on the planet. The deepest plains are 5 kmbelow the reference altitude (Hellas Planitia), andthe highest mountains are as much as 25 km abovethe 0 km plane (Olympus Mons). Thus, a 30 kmtopographic variation exists at Mars.

Topography is important for balloon design forthree reasons. First, very large balloons are

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required to fly over the highest mountains. Second,the thin Martian atmosphere makes entry difficult.Atmospheric entry at a low location is desired.

Third, balloons flying at reasonable altitudes (say,6 km) may fly above high plateaus (at, say, 5 km)and be very close to the surface of the planet. Theballoon will see 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° < Ls< 360°), Mars is closer to the sun than during thenorthern hemisphere summer (90° < Ls < 180°).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 toatmospheric pressure variation. On diurnal timescales, Mars' pressure profile varies only slightlybut significantly, less than ±10% per day. The

diurnal pressure variations result from the passageof weather systems.

Atmospheric 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 Boundary Layer Model (BLM). The MarsBLM estimates environmental conditions at onelatitude, for one Martian day on a variable, closely-spaced altitude grid (5 m/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 OutputsInputsOptical Depth, xSolar Longitude, LsSurface AlbedoSurface TISurface Pressure, P0Solar Flux

Outputs (24/day)T,PSolar ,Direct RadiationSolar Diffuse RadiationIR Diffuse RadiationSolar Zenith Angle

As stated above, the MABS effort was a feasibilityassessment for a Mars 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 Inputs. 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.

We 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 be morespecific in longitude, topography, and time of yearthan the GDC. Table 3 summarizes some of theparameters for each design case. The latitude range

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corresponds to the region of the planet, the solarlongitude Ls corresponds to the expected arrivaland flight windows, and the topography indicatesthe range of topography to be evaluated.

For each design case, the lowest thermal inertia,highest albedo, and highest K locations wereidentified. From these parameters, BLM inputswere selected that reflect the worst-case conditionsfor the region. Table 4 shows the details of thethermal inertia and albedo evaluations for the NDCenvironment. The thermal inertia and albedochosen for the BLM inputs are somewhat lowerand higher, respectively, than the worst cases forthe region to account for measurementuncertainties.

Tables 5, 6, and 7 show similar data for the EDC,SDC, and GDC respectively.

For each BLM design case, the selected altitudeswere converted into surface pressures. Previouswork done for the MAP proposal provided chartsof balloon diurnal superpressure variation as afunction of Ls at given latitude using latitude-averaged albedo and thermal inertia (Ref. 2). Forthis study, we selected the worst-case Ls for eachdesign case based on the MAP charts. Thecorresponding lowest surface pressure PO w a sselected by taking the surface pressurecorresponding to the worst-case L, and reducing itby 10 %, the maximum diurnal variation expectedfor surface pressure. Table 8 shows the results ofthis analysis for each of the design cases. Note thatthe suffix "h" on the design cases refers to aplateau that would be unfavorable for balloondesign as discussed above.

Table 9 shows the BLM inputs selected for theballoon design activity. These eight atmosphereswere deemed the worst case environments forballoon flight.

Mars General Circulation Model (GCM). TheMars General Circulation Model (Mars GCM)grew from an Earth weather model in the late1960s. It simulates the dynamics of the Martianatmosphere on a global scale, much like weathermodels for the Earth. Atmospheric parametersavailable from the GCM include winds, solarradiation (direct, upward diffuse, and downwarddiffuse), infrared radiation (upward anddownward), temperature, pressure, and topography.

Initially, the Mars GCM modeled only two verticallayers in the atmosphere. As the cost ofcomputational power decreased, improved verticaland horizontal resolution was added to the GCM.Topographic effects have been included. The MarsGCM now models the effect of dust on radiation,dust-wind interactions, and the seasonalsublimation patterns of the polar ice caps (Ref. 3).

The GCM was used to develop balloon flighttrajectories based on winds that are representativeof the conditions expected upon arrival (Ref. 4)Forecasting the weather at Earth is a difficultendeavor, and predicting the environmentalconditions at Mars is a daunting task. The GCMprovides representative weather that could beexpected at Mars at the arrival times underconsideration. Thus, the GCM is appropriate fordevelopment of mission strategies and profiles.

The Mars GCM requires several inputs including(a) the initial solar longitude Ls, (b) the initial massof the atmosphere, (c) the total atmospheric dustload, (d) the global distribution of surface thermalinertia, (e) the global distribution of planetaryalbedo, and (f) the assumed ratio of theatmospheric optical depth in the IR wavelengths tothe atmospheric optical depth in the solarwavelengths, tm/tso]. Viking data guided theselection of inputs for the GCM runs employed inthe MGA simulations developed for this study. Forexample, the total mass of the atmosphere wasselected such that the predicted seasonal surfacepressure variations match the Viking measurements(Ref. 2).

GCM Output. The GCM output used for predictingMGA trajectories is arranged in a planetary gridupon which atmospheric parameters are reported.The following parameters are given at each gridpoint:

• atmospheric temperature,• surface pressure,• the solar zenith angle,• direct solar flux,• downward diffuse solar flux,• upward diffuse solar flux,• downward IR flux,• upward IR flux,• the east-west wind (zonal) component, and• the north-south (meridonal) wind component

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Table 3 BLM Design case parameters.

NDCEDCSDCGDC

Latitude30°


Table 8. BLM input surface pressures.

NDC (0 km)NDCh (5 km)EDC (0 km)EDChJS km)SDC (0 km)SDCh (5 km)GDC (0 km)GDCh (8 km)

Worst-case Lsfrom MAP chart

225°

240°

275°

240°

Corresponding P0[mbar]

6.75

7.25

7.40

7.25

BLM P0[mbar]

6.03.86.54.26.74.36.53.2

CorrespondingSolar Flux [W/m2]

718

728

718

728

Table 9. BLM Inputs.

NDC (0 km)NDCh (5 km)EDC (0 km)EDCh (5 km)SDC (0 km)SDCh (5 km)GDC (0 km)GDCh (8 km)

tr-i0.0

0.0

0.0 .

0.0

N Latitude[°]

30°

-20°

-30°

-20°

TI[SI units]

80

70

160

70

albedo[-]

0.32

0.36

0.28

0.36

L,n225°

240°

275°

240°

Po[mbar]

6.03.86.54.26.74.36.53.2

Solar Constant[W/m2]

718

728

718

728

The pressure at any altitude z can be calculated by

where H is the atmospheric scale height, taken tobe 11.18 km, and zsurf is the surface altitude.

The GCM spatial grid contains 7 vertical divisions(up to about 10 km above the planet's surface), 40longitude divisions, and 25 latitude divisions. Theparameter grid is updated sixteen times per Martianday.

The Mars GCM run that was used for the trajectorypredictions for MABS begins at Ls = 220° and has'tm = ^sol-

Figure 2 illustrates the relationship between thesevarious environmental factors and conditions andthe behavior of a superpressure balloon at Mars.

Balloon Systems

The Balloon System is defined as the BalloonFlight System (BFS) and the Deployment andInflation System (D&IS). The BFS consists of theBalloon and the Gondola (which will be addressedseparately in this paper). The balloon included theenvelope, seams, end fittings, load core, inflationtube, diffusers, payload tether/shock attenuator, and

separation hardware. The D&IS included theballoon container, deployment hardware/sequencer,tanks, gas and control hardware.

As stated previously, the approach was to heavilyleverage ongoing aerobot/balloon activities, takeadvantage of previous development studies andtesting, and to design for the worst caseenvironmental conditions. It was identified veryearly in the study that two technologies—theenvelope material and the D&I of the balloonvehicle, would be critical to the success of a MarsAerobot/Balloon mission. Furthermore, both ofthese systems had to be integrated as a system anda systems design approach was imperative. Thedesign of the balloon and the design of the D&IS,heavily influenced the other. Neither could bedesigned without assessing the impact on the other.

A balloon design tool (Ref. 6) was developedwhich resulted in iterative equilibrium solutions. Itallowed for solutions for cylindrical and sphericalballoon shapes based on meeting steady stateday/night extremes. The principle unknown wasrequired material strength. Inputs includedmaterial areal density, radiative properties,atmospheric and thermal environment, performancerequirements, and design constraints. Outputsconsisted of system masses, physical properties,and balloon temperatures.

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Balloon Flight System

The performance of superpressure (SP) balloonshave been well documented. Balloons of the sizerequired for a Mars mission have been successfullyflown in excess of 90 days. However, the materialused in these balloons, primarily mylar, had veryhigh areal densities, problems with pin-holing andfracture, and were not designed for atmosphericentry deployment and inflation. Experience withscaling up of these balloons to larger designshighlighted many inherent material and fabricationproblems. Application of other materials such asNylon-6, while alleviating some of the problemsbut not all, had very limited successfuldemonstration. In addition, neither of the twomaterials could meet the design strength

gas integrity, and geometry, 4) accessories such asfittings, reinforcements, inflation, etc., 5)fabrication such as tolerances, 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 mass/unit area and macro stressstate, it presents increased difficulties during thedeployment and inflation process. The cylindricalballoon, offers advantages for ease of reefing, loadintroduction during the D&I, and improvedfabricationsuch as ease of maintaining tolerances.However, it results in a larger and heavier balloonand stability problems during inflation of theballoon during descent from altitude. A sphericaldesign was chosen because of mass limitations andstability concerns during the deployment and

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or homopolymers such as current balloon films ofpolyethylene (PE), mylar or nylon-6, identified themodification of one parameter to improve itsperformance comes at the expense of otherparameters. Often the gain in a desirable parameterwith a material selection is offset by a loss inperformance in another parameter.

Many different types of materials were consideredfor the Mars application. Leveraging workcurrently being pursued by the NASA BalloonProgram, it was decided to pursue the developmentof a "composite" "film" or material. The approachwas to take the best properties of differentmaterials and combine them in a manner whichwould offer the desired overall set of mechanicalproperties as well as alleviating the individualcomponent deficiencies. Several materialcombinations were investigated and a prototypematerial was made and tested. The material (Ref.6) consisted of a 3.5 micron Mylar base film,adhesively bonded to 55 denier Kevlar scrim in anorthogonal pattern, which was then laminated to 6micron SF-372® PE. The Kevlar scrimreinforcement provided the requisite strength withvery little weight penalty. The adhesive layersprovided the bonding between the differentcomponents as well as adding pinholing resistance.The mylar provided the substrate stiffness while theSF-372® provided pinholing resistance andfracture toughness.

The resulting material was successfully tested andhad the following properties:

Areal Density : 19.66 g/m2Tear Toughness: 12 N @ +23°C to 43 N @ -140°CStrength : > 2500 N/m from +23°C to -140°CSeam Strength : > 2500 N/m from +23°C to -140°C

The material was subjected to a range of radiationdose levels to determine if the material wouldsurvive the requisite sterilization procedures beforegoing to Mars. It was determined through testingthat no degradation of strength occurred forrequired radiation levels of 2 Mrad. Limited testingalso revealed that there were no observedcomponent or adhesive delaminations orembrittlement at -198°C.

Although the prototype material met therequirements established for a Mars material, muchwork and testing would still be required to optimizeand develop it for full scale production. However,it did successfully demonstrate the feasibility thatan adequate Mars material could be made withsufficient advanced development.

14 GAUGE MYLAR

ADHESIVELAYERS

ORTHOGONAL55 DENIER

KEVLAR SCRIM 25 GAUGEPOLYETHYLENE

Figure 3. Schematic representation of candidatecomposite material

Flight Performance

Several factors must be considered if a balloon is tosurvive for an extended mission duration. Thesefactors include the desired payload float altitude,altitude stability, thermal optical properties, gasselection, CO2 condensation, ultra-violetdegradation, dust collection, diffusion and effusion,and of course the overall flight environment.Various performance and design studies wereperformed using the balloon design tool. Inaddition, there was a review of previous SP flightdata. It was determined from the SP flight data thatwe could achieve a 90 day mission based onprojected diffusion and effusion gas losses. It wasalso determined that the balloon temperaturesbased on the extreme environment, camedangerously close to the atmospheric CC>2condensation point thereby necessitating analuminized top hemisphere of the balloon toprevent condensation from taking place during thecold Martian nights.

Deployment/Inflation System

Equally challenging as the material design and infact driving it was the method of deploying andinflating the balloon during descent through theMartian atmosphere. The deployment and inflationsystems are extremely tied to each other as well asthe balloon structure/shape, the balloon envelopematerial, and spacecraft volume and mass. Theyare a system and one cannot be modified without itaffecting all of the other systems. As a result, thespreadsheet design tool was used extensively todetermine the interaction between the varioussystems and to determine mass allocations anddistributions.

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Deployment. Deployment considerations included:1) packaging of the balloon such as density, foldingmethod, etc., 2) method of deployment such asgravity or mechanical 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 CNES/Toulouse 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 as part of this study effort for use indetermining balloon geometry and payload andentry vehicle mass distributions during thedeployment process. As a result of stabilityconcerns and mass distributions, it was decided thatthe balloon should be suspended from theparachute and between the inflation 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 inflation wasdetermined by the time it takes for the system toslow to an acceptable dynamic pressure to insure astable system plus the time it takes to inject the gasinto the balloon. Although a top down inflationsimilar to the VEGA and Mars '96 designs washighly desired, stability analyses coupled withother system masses and physical geometry led tothe placement 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 gasproceeded up through the center of the balloonthrough a central load carrying inflation tube untilwhere the gas exited into the top of the balloon.

Baseline Design

The baseline Mars balloon system is described infigure 4. below.

VOLUME: 10,500 m3DIAMETER: 27mBALLOON MASS: 55 kgGAS MASS: 12 kgPAYLOAD: 15 kgFLOAT ALTITUDE: 6.5 kmDAYTIME AP: 240 PaNIGHT AP: 20 Pa

MTRL: 3.5 |im MYLAR / 55 DENIER KEVLAR / 6 ,im SF-372AREAL DENSITY : 20.0 g/m2MATRL. STRENGTH: 2500 N/mSTRENGTH FACTOR OF SAFETY: 1.5COATINGS: TOP IS ALUMINIZED (TO PREVENT C02

CONDENSATION), BOTTOM WHITE

Figure 4. Baseline balloon design

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

Gondola and Science Instruments

Gondola Design Objectives

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, the goalis to incorporate the technology on-board theballoon for high-precision knowledge of theballoon position and orientation. These data areused for science sequencing and balloon tasksequencing on-board which promise todramatically improve the overall science qualityand data return from a Mars aerobot mission.

The specific study objective was to develop aballoon gondola concept capable of maximizing thescience return from a 90-day balloon aerobotmission to Mars. Concepts and options relating toMars Science Instruments, Mission SequencingConcepts, Planet-Wide Position Determination,Earth Communication, Power Generation, Storageand Management, Integrated Structure and

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Thermal Control, and Ballast/Micro-packageDeployment, were all investigated as part of theeffort. The emphasis was on identifying a veryconservative point design that would serve as astarting point for detailed study and designiterations in the next phase of study. While anactual mission would be driven by a combinationof cost and mass constraints, the initial effort wasto be towards a minimum mass design. In theprocess, key technology development items wereidentified and the major trade issues requiringfurther investigation in later studies were specified.Also, a preliminary costing of an implementationphase was performed.

Gondola Design Approach

Given the time available for the study, it wasdecided to extensively leverage earlier gondolastudy and design and development efforts from theCNES Martian Aerostat program and the MarsAerial Platform (MAP) Discovery proposal. TheMFEX/Sojourner rover design experience onmechanical/thermal systems as well as costing datawere also used in the study.

In order to minimize the mass of the system, it wasdecided to take an integrated approach fordeveloping the science instruments and thegondola. The implication of this was to consolidateinstrument data processing within the centraldata/computer system of the gondola with noseparate CPU for each science instrument.Wherever possible the science instruments wouldbe used to support on-board engineering functionssuch as navigation by means of ground imagesobtained by a science camera. Thermal control andmechanical structure would be integrated andpower regulation and management would beconsolidated.

It was also decided to adopt a reliability approachto gondola design involving the selection of high-reliability components for mission criticalfunctions, augmented with higher-risk/high-payoffcomponents for non-critical functions. Extensivetesting would accommodate the incorporation ofthese additional elements.

Gondola Design Assumptions

With regard to gondola communications, it wasassumed that relay communications would bepossible by using either the MGS 96 or MSP 98Orbiter communication links, with full availabilityof both orbiters during all communicationopportunities to permit maximum data return.Image data would be compressed 10-to-l to allow

the return to accommodate the availablecommunications bandwidth.

With regard to energy and power assumptions, thedesign was targeted for a reasonable worst casesolar power/energy condition corresponding to anorthern winter mission. Night-time scienceoperations were minimized to conserve energy, andall energy calculations were made using an 8 hourday, followed by a 16 hour night. A moderatelyhigh optical depth of 1.0 was chosen for the floataltitude corresponding to an optical depth of 1.7 atground level. If the actual optical depthencountered was extremely high (i.e. > 2) thenmission activities would be descoped byperforming less imaging and thermal emissionspectroscopy to stay within reduced power budgets.

It was also assumed that Radioisotope HeatingUnits (RHU's) would be acceptable for meeting thethermal design needs. This assumption allowed theuse of lumped thermal models in analyzing thegondola. A more detailed thermal analysis withcareful configuration of thermally sensitive itemsand appropriate thermal inertias and sources couldlead to a non-RHU thermal solution. However, thiswould have to be undertaken in the next phase ofstudy.

Mission Requirements on Gondola

The gondola was required to carry a sciencepayload with a mass of 3 kg, peak power of 3 W,and a volume of 3 liters. The peak data rate fromthe science instruments would be 500 kbits/s, witha maximum data storage requirement of 500 Mbits.

Communication capability was required fortransmitting balloon engineering and science datato MGS '96, MSP '98 and MSP '01 spacecraft. Itwas a requirement that the gondola acceptcommands from both the MSP '98 and the MSP '01spacecraft. The full range of balloon-centerdorbiter azimuth and elevation angles was to be usedduring any given pass. Storage of all engineeringmeasurements start at entry to insure recording ofdeployment and inflation data which are laterreturned to earth.

Navigation requirements were distinguishedbetween the accuracies that were to be achievedon-board the vehicle and those that needed to beachieved on the Earth after data return and analysis.On-board knowledge requirements for Altitude,Latitude and Longitude were respectively 1 m, 1km and 5 km. Post-flight accuracies were forAltitude, Latitude and Longitude were respectively1 m, 100m and 1 km.

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In addition to the sensor derived estimates of theaerobot position it was also desired to be able topredict the position of the aerobot into the future.The accuracy of these predictions would bedependent on the vertical balloon performancemodel and the planetary Global CirculationModels. A 1-day prediction accuracy requirementfor Altitude, Latitude and Longitude wasrespectively 100 m, 10 km and 100 km. Accuracyfor a week prediction was desired to be 200 m, 200km and 500 km and a 30-day accuracy of 500 m,500 km and 2000 km respectively in Altitude,Latitude and Longitude.

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 of

arc.• Altitude measurements better than +/- 1%.• Sun Elevation/Azimuth measurements better than

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

less than 3 s

> On-board clock accuracy better than 20 s.1 Short term stability of oscillator for one-way

Dopplerof I"10.1 Ballast/Probe drop activation time less than 60 s.

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

Gondola Science Imaging

As science imaging of the Martian terrain is one ofthe most important objectives of a Mars aerobotsystem, some of the issues relating to databandwidth and image quality are discussed.

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 1000 x 1000 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

UHF ANTENNA

NEUTRON SPECTROMETER

STAR SENSOR

BATTERIES

PROBES (6!

/—CPU / MEMORY

PROBE MAGAZINE

BUN SENSOR

POWER REGULATION

COMMUNICATIONELECTRONICS

CAMERA. HIGH RESOLUTION

1NERTIAL SENSORTHERMAL EMISSION SENSOR

SELECTIVE GAS SENSOR

AEROSOL SENSOR

ATMOSPHERIC SENSOR

CAMERA. OBLIQUE

Figure 5. Gondola Concept

— OPTICAL ALTIMETERCAMERA, NADIR

CAMERA POWER CONDITIONING

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MASS GONDOLA FUNCTIONAL ARCHITECTURE

FromBalloonSystem

Analog SignalsDigital Lines <Poww '

' ScienceCommand &Data

Controller-CPU-DIO.A2D- Data Storage-Clock

PowerConditioning

i

1

*

•* ——

Battery

PowerMgmt

Figure 6. Gondola Functional Block Diagram

wind-speed of 20 m/s the ground-track motion is576 km. Thus depending on the wind-speed theimage return on average could range fromapproximately 1 image per kilometer of travel atlow wind speed to a value of 1 image for every 3km of travel at higher speeds. While these imagescould be uniformly spaced along the ground track,it is more meaningful to concentrate these imagesat specific science targets. It is also important thatthe images that are taken at high-resolution be alsoaccompanied by a context image that allows thehigh-resolution image to be correctly interpreted.

There are several issues that impact the definitionof an appropriate imaging strategy . The balloon ata float altitude of 6 km above the terrain will beable to obtain nadir pointed contextual (or framing)images with a footprint of 10 x 10 km at aresolution of 10 m/pixel, and nadir pointed high-resolution images with a footprint of 200 x 200 mfootprint at a resolution of 0.2 m/pixel. Even asingle high-resolution image requires a contextualframing image to be useful for subsequentlocalization of the high resolution data.

The rotational dynamics and ground track motionof the balloon platform requires that high resolutionimaging employ one or more of the following

motion blur compensation methods: fast exposures,motion stabilizing optical path elements, or carefultiming of camera exposure/read-out. Even then, theyield of blur-free images may drop because ofgusts, dim ambient lighting, or constraints onexposure due to filter wheel use.

Let d be the imaging distance, r be the resolution atthe imaging distance, v be the balloon translationvelocity, Trou be the period of the balloon's rollmotion about its vertical axis, Tpend be thependulum period of the balloons swing motion,QTO/; be the maximum roll angular velocity, £lpendbe the maximum pendulum angular velocity, and 0be the maximum pendulum angle. Typical valuesderived from Mars Balloon/Atmospheric modelsand French/Wallops experiences with high-altitudeballoon experiences indicate values of d = 6000m; r = 0.2 m; v = 60 m/s; Tmu = 1200 s; Tpenj = 30s; 9 = 0.25 deg; Qr,,u= 1.5 deg/sec; £lpend= 0.05deg/sec. This data indicates that the effect of thependulum motions is to contribute to an apparentground-track motion of 6 m/s. If the pendulumswing is aligned with the ground-track motionthen it is possible to time the camera exposuressuch that the pendulum induced motion effectscompensate for the actual ground-track velocity.

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Another technique for motion blur compensation isto synchronize the image data readout with theapparent ground-track velocity. This is possibleonly if the detector axes aligns with the ground-track motion, a condition that will occur once every72 km for the data set discussed here. Analysisalso indicates that the exposure time for a 1-pixelblur is 3.3 ms for a balloon ground velocity of 60m/s.

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 200 m 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.

Science Sequencing: Concepts

On-board state determination involves processingsensor data to obtain knowledge of balloon positionin the form of latitude and longitude as well asdistance vectors to specific targets. It providesinformation on orientation such as the tilt of theballoon platform, and its pointing and line-of-sightsto targets. Furthermore, state determination allowscomputation of rate information such as theground-track velocity, the climb rate, balloongondola swing rate, and rotational speed of rollmotion about the balloons vertical axis.Knowledge of this state information allows theMars aerobot to optimally execute its on-boardscience and engineering sequences. For example,position information could be used to activate nadirpointed cameras as the aerobot over-flies a sciencetarget site. Accurate relative position informationwith respect to a target and knowledge of theballoon's orientation could be used to maximizeoblique camera coverage of the target. To performuseful science with the specificity required byplanetary scientists, on-board position accuracybetter than 10 km would be desired with higher,target-relative accuracies of 1-2 km as the ultimategoal. Rate information on the balloons angular

GONDOLA SCIENCE DATA ACQUISITION EXAMPLESAMPLING OF HUNDREDS OF SCIENCE SITES

DAY DAY

NIGHT

SCIENCETARGETSITE

NIGHT

GROUNDTRACK600-1700 km/8hr-day

NEAR-IRIMAGE4 m/pixel

VISUALIMAGE2 m/pixel

20 m/pixel20km

Figure 7. Science sequence concept

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motion could be used to control the exact momentof triggering a camera, such that motion blureffects are minimized. Altitude climb rate datawould be used by on-board sequence control tocontrol ballast drops to avoid terrain hazards.

Post-flight state determination would be made bycorrelating aerobot images to planetary imagedatabases, by more refined processing of variouscelestial sensor data, and by detailed analysis ofDoppler data from orbiter return of balloon data.

At or near an important science target site it isdesirable to take more images, possibly in differentspectral bands. It may be desirable to increase thequality of an image transect by increasing imageoverlap between successive high resolution images,taking oblique view images in preferred directions,or obtaining higher quality images by carefultiming of the exposures to minimize motion blur.Other actions that can be taken near science targetsare to activate dormant sensors, and optimize themix of sensor sampling activities and rates. Onemay also drop in-situ science micro-packages atsuch science targets.

Optimal science data return can be achieved in anumber of ways. One technique is to exploit thecapabilities of the on-board data systems by databuffering and increased ground-interaction throughthe command link. The aerobot would return"thumb-nail" images or data sets, and a subsequentcommand from the ground could be used to selectthe images that would be transmitted at highresolution. Another technique is to enhance theprecision of the on-board sensors and positionestimators. With this more careful targeting andstate-driven sequencing of science data acquisitioncan be achieved. Yet another technique is toimplement on-board recognizers for specificscience data patterns. The aerobot would then tailorits sensor data acquisition in response to triggeringsignals from the science pattern or templaterecognizers.

Among the alternatives considered for sequencingwere a Fixed Timer Sequence which wouldperform a timer reset at the night-to-day terminatorand activate science data acquisition at fixedintervals. A variant of this method is an Earth-based timer sequence where the timer activationtimes are set from earth based on predicts of thenext-day wind-blown path of aerobot. An on-boardstate-driven sequence would tie data acquisition tothe balloon's on-board knowledge of its "state" i.e.latitude and longitude and a list of desired sciencetargets. When the balloon knows it is near a target,it adapts its science gathering sequencingaccordingly. A science-driven sequence would

adapt data gathering in response to simple on-boardanalysis (data quality, pattern recognition) ofscience data. For example, if the on-board NeutronSpectrometer profile indicated water, then thiscould result in increased Thermal EmissionSpectroscopy measurements. Finally, a DataBuffer Based Sequence would require only asummary report to be sent back to earth with the hi-resolution stored on-board to await an up-loadselection command the next day. Of thesetechniques the methods based upon estimating theposition of the aerobot and driving the scienceactivity to the position state appears to be the mostfeasible, together with some limited data bufferingto selectively control the transmission of high-resolution data sets.

Science Instruments Summary

The following strawman science instruments wereselected for the MABS design:

High-Resolution Nadir Camera. Provides 200 X200 m image at 0.2 m/pix resolution image for nearground-truth data in different spectral bands using a6 position filter-wheel. Blur compensation is byusing fast shutters and/or motion-stabilizing opticalpath elements.

Wide-Angle Nadir Camera. Provides 10 X10 kmcontext for high-resolution images at 10 m/pixelresolution.

Medium-Angle Oblique Camera. Provides photo-geology side view images of mountainous/canyontargets for get-strata analysis.

Thermal Emission Spectrometer (TES). Providesmineralolgical information at 100 m resolution.

Neutron Spectrometer. Provides water, carbonates,nitrates composition two orders of magnitude betterthan orbital data.

Atmospheric Sensors. Supports atmosphericscience, and balloon performance modeling.

Aerosol Sensor. Provides dust and aerosolconcentrations.

Selective Gas Sensor. Provides information onwater, methane gas levels.

A near IR spectral imager was also identified as avery desireable instrument to address surfacemineralogy, but was included too late in the studyto incorporate into the detailed gondola design

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The total mass of all these science instruments wasestimated to be 4950 g including contingency andmargin ranging from 25% to 30%. Daytime andnighttime energy requirements were estimated tobe 32.2 and 7.3 W-hr, respectively.

On-board State Estimation

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

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 position estimation. The tiltestimates are also used to distinguish betweentranslation and rotation effects when using frame-to-frame image registration based methods ofground-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 nighttime 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 laser rangersas well as radar altimeters.

The ground-track velocity is needed to understandthe wind patterns and the long-distance balloontrajectory. The velocity can be integrated for high-accuracy knowledge of ground-track, as well as forcompensating for measurements of orbiter-to-gondola Doppler measurements. Sensors to

provide this information include inertial sensors,frame-to-frame image based comparison methods,Doppler radar, and celestial position differencing -especially at night when high accuracy positionestimation using star/moon fixes is possible.

Determination of the gondola latitude andlongitude 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 signalacquisition/loss 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.

Nighttime on-board position and velocity would beobtained by a sensor strategy that consisted of thefollowing:

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•Successive Phobos elevation measurements whenPhobos is visible (approx 3 hrs), illuminated,not in shadow, and when balloon rotation alignfield-of-view.

•Successive bright-star(s) elevation measurementsevery 15 minutes.

•2-4 orbiter (MGS, MSP'98) fixes using signalacquisition/loss or Doppler

• Full 6-DOF inertial estimator if power budgetpermits.

• Doppler radar sensor if mass and power budgetpermits.

• Height measurements concurrent withradiometric measurements.

• Ground-track velocity estimation by differencingcelestial position measurements.

An earth-based position estimate would beobtained by correlating nadir context framesreturned by the aerobot with Mars maps. No VLBItechniques are anticipated.

The sensor set selected for the gondola consists ofinertial sensors (3-gyro & 3-accelerometer), a Sunsensor, a Phobos and star camera, a laser or radaraltimeter. The total mass of the sensor hardware is1325 g including contingency/margin ranging from5% to 30%. Day and nighttime energyrequirements were estimated to be 3.0 and 0.9 W-hr, respectively.

Earth Communication Approach

The communication to Earth would be through theUHF Balloon Relay link of the MGS Orbiter (ifstill healthy) and the similar link through MSP '98Orbiter. Redundancy would be achieved bygondola hardware to support both orbiter links. Inaddition, the command link capability of the MSP'98 orbiter would be used tailor the gondolamission for maximum science return in the form ofnew target lists, as well as providing orbiterephemeris information to the aerobot for use withone-way Doppler measurements. The hardwareimplementation leverages the New MillenniumDeep Space #2 mission and MGS UHF-linkdevelopment, with some additional development(e.g. conventional transceiver) if on-board one-wayDoppler measurement was found to be necessary.

The total mass of the communication systems wasestimated to be 428 g including contingency. Dayand nighttime energy requirements were estimatedto be 2.2 and 4.4 W-hr, respectively.

Data and Computer System

The on-board computer and data system supportsthe compute intensive functions during day-time

operations such as image compression; TESaveraging and Fourier transforms; image-basedground-track motion determination and imageanalysis; and navigation sensing computations.Minimal operations are supported at nightincluding Moon/Star based navigation and ambientsensor data acquisition.

The baseline system consisted of a Power PC-603EMulti-Chip Module Controller/Data-Storage unitthat would support low-power features such ashigh/low processor clock rates and sleep modes.Power usage would be 5 W at 50 MHz clock, 3 Wat 25 MHz clock, and 20 mW during sleep mode,with just enough power to keep-alive SRAM andbus-controller. In addition, low-dissipation energysources in the form of linear regulators and/ordirect power source operation with battery orcapacitor would be used to support sleep modeoperations at night thereby avoiding switchingpower supply pre-load power dissipation.Aggressive power management of active/inactivememory and I/O modules would result insubstantial power saving. In addition, a fieldprogrammable gate array (FPGA) or micro-controller would be used to buffer data of low-ratesensors at night to avoid frequent wake-ups of themain processor. A 1 Gbit solid-state recorderwould be used for on-board data storage.

The total mass of the computer and data systemswas estimated to be 740 g includingcontingency/margin ranging from 5% to 30%.Daytime energy requirements were estimated to be8.6 W-hr and night-time needs were estimated to be7.9 W-hr.

Power System

The solar array size is designed for a surfaceoptical depth of 1.7 at 40 deg N during the northernwinter. High-efficiency GaS (21.5%) cells are used.An energy margin of 25%, a battery depth-of-discharge of 75%, a converter efficiency of 75%;and a battery efficiency of 80% are assumed for thedesign.

For a total estimate energy need of 124 W-hr, the"?system consists of a solar cell of area 0.76 m ,

with a mass of 1520 g (not including mechanicalsupport structure), and a battery mass of 540 g.

Mechanical/Thermal System

The design here is for a light-weight multi-functional structure, with a "Warm ElectronicsBox" design inherited from the Sojourner roversystem. The ambient thermal environment has lessday/night variations than that of a rover, and the

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initial thermal design using a lumped thermalmodel, generic temperature requirements, andRHU's should be adequate. Detailed thermaloptions to be investigated in the next study includeheat-pipes, and phase change materials. Themechanical system accommodates a ballastturret/controller 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 totalmass of the ballast packages was assumed to be3000 g.

Mass. Energy Budgets Summaries

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

Table 10 Mass and energy budget

Item

Science InstrumentsCommunicationPosition & AttitudeSensorsComputer and DataSystemPowerStructure, Thermal,Cabling

Payload Total

Ballast

Mass

GO

49504301325

745

26003155

13205

3000

Day-TimeEnergy(W-hr)

32.22.23.0

8.6

0.80

46.8

0

Night-TimeEnergy(W-hr)

7.34.40.9

7.9

1.60

22.08

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

Gondola and Mission Design Considerations

The following considerations were addressed tosome level as part of MABS gondola systemdesign:

• Commercial-Off-The-Shelf (COTS) parts vsCustom components.

• Instrument computers vs General PurposeComputer

• ASIC/DSP vs General Purpose Computer• Position/Attitude sensitivity as function of sensor

selection and design parameters• On-board 1-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 Shuttering 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 exo-skeletal structure vs temporary

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

Shadowing

Balloon Delivery System

Design 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 system that is scheduled to enter Marson July 4, 1997. The balloon/aeroshell is launchedin an inverted configuration similar to Pathfinderand Mars '98. The cruise stage design and

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equipment layout is derived from an LMASpaceprobe/Discovery design to reduce the non-recurring cost. A modest level of technologyadvance reduces mass and allow the use of a lowcost Delta 7325 launch vehicle. The total balloonsystem mass of 212 kg can be accommodated withthis lower mass cruise stage and the smallerdiameter aeroshell of Mars '98.

Launch Configuration

The current best estimate for the MABS launchmass is 534 kg. which provides a 15% launchmargin for the Delta 7325 capability of 616 kg.Table 11 summarizes subsystem masses. Theaeroshell is supported on the launch vehicle in aninverted configuration. The 96 cm diameter cruisestage structure mates to the Delta upper stageadapter with a standard clampband. Figure 8illustrates a possible launch configuration. Thecruise stage attaches to the backshell at the otherend of the 96 cm cylinder with an interface ringand six bolts. The entire stack is carefully balancedto meet the Delta center-of-gravity requirementsfor spin-up to 70 RPM for launch. An umbilicalline runs from the balloon system through thebackshell to the cruise stage. The cruise stage solararrays is mounted to the back of the equipmentdeck.

Table 11. Reference Cruise and Entry SystemMass Summary

SubsystemPowerTelecomAd&CC&DHThermal ControlStructure MechanismsPropulsion - DryPropellantEntry Vehicle*

Total Mass, kg4721

6115

1817

175385

* Includes 212 kg Balloon System

Flight Element Total 526Launch Adapter 8Total Launch Mass 534Delta 7325 Capability 616Launch Margin 15 %

Mechanical Design

The flight delivery system consists of the cruisestage, backshell, and aeroshell. A singlemonopropellant propulsion system mounted on thecruise stage provides all the propulsiverequirements for cruise. Entry is spin stabilized so

the propulsion system can be jettisoned at entry.The propulsion system consists of two propellanttanks and two Rocket Engine Modules (REM).Each tank carries 7.5 kg of hydrazine. Each REMconsists of one aft pointing 22N (5#) thruster andtwo opposing 22N roll thrusters. The DeliverySystem design can fully support the missionrequirements including meeting the dynamicrequirements for balloon deployment.

The cruise stage structure provides the interface tothe launch vehicle and provides mechanical supportfor the sensors, solar array, and antennas for cruise.The cruise stage thermal control is passive withspecific coatings to reduce heating near Earth.

The aeroshell consists of the composite heat shieldand the backshell which are coated with a 1.5 cmlayer of SLA-561 ablator for dissipating the entryheat load. The entry ballistic coefficient of 50kg/m2 is lower than Pathfinder which aids inreducing parachute deployment velocity andopening shock. (The entry overload is only 15 gwhich is lower than Pathfinder). After thehypersonic phase, a drogue parachute is deployedfrom the top of the backshell. This parachute is

Figure 5. Mars Balloon Delivery System

similar to the Mars '98 design and is deployed by amortar. The heatshield is released and droppeddown on a long tether shortly after drogueparachute deploy. The drogue parachute is used todeploy a much larger main parachute that is used toreduce the dynamic pressure for balloon inflation.The mass of the heatshield is used to stabilize theballoon during inflation. The backshell andparachute are jettisoned at >0.5 kilometers altitude(relative to reference level) after balloon inflation iscomplete. Entry and deployment are planned overlow lying terrain (0 to -2 km) to insure successfuldeployment before the system hits the ground.

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Electronics

There are eight unique command and data handling(C&DH) cards (13 total with redundant cards) andeight unique power distribution cards to control theflight system 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.

Entry 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 drogueparachute 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. Thedrogue parachute carries away the backshell but theheatshield is retained until balloon inflation iscomplete.

The flight system has strong heritage fromViking/Pathfinder and Discovery Stardust. LMAbuilt the aeroshells 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.

Summary

The Mars 2001 Aerobot/Balloon System Studyeffort, using conservative design assumptions, hasresulted in a feasible design for a long-durationmission to Mars. The key technology requirementsfor this mission have been identified and atechnology plan has been constructed. A baselinemission and system concept has been defined wellenough to generate rough estimates of a totalmission cost. Excluding launch, cruise spacecraft,entry vehicle and project reserves, a Mars Aerobot

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mission is estimated at less than $50M (FY97$).The cost estimate is quoted in this manner becausesuch a mission could piggyback onto a futureplanned Mars mission if the launch vehicle andentry vehicle are large enough to accommodate theballoon system. If a dedicated launch vehicle,cruise stage and entry vehicle are required, the costof an Aerobot mission is approximately three timeshigher. The lower cost makes such a mission abouttwice the cost of the 1997 Pathfinder Rover. Theupper cost level is roughly comparable to the costof future dedicated Mars lander and orbitermissions.

The MABS study focused on the 2001 opportunityand up until October 1996, this opportunity wasviable provided the advanced technology effortshad proceeded. These efforts did not go forward atthat time which has resulted in 2003 being the nextpractical Mars Aerobot/Balloon opportunity.

What are the next steps for flying an aerobot atMars? First and foremost an aerobot must beconsidered by the Mars program planners asanother feasible tool for exploration andpreparation for future missions including samplereturn. An aerobot can provide images at a scaleand resolution relevant for scientifically valuableand safe lander and rover operation. In addition toproviding scientists with a unique vantage point forexploration, an aerobot can contribute to the searchfor possible life. As an aerobot cruises above theMartian landscape it has the capability to exploreliterally hundreds of potentially interesting sites toexobiologists at a resolution relevant toidentification of promising places to visit withlanders and rovers.

Another key step in making a Mars balloon oraerobot mission a reality, is insuring that long-leadballoon envelop technology is pursued by NASA.The Ultra-Long Duration Balloon (ULDB)technology program for Earth scientific ballooningcontributes to technology for a Mars balloon,however, because the Mars requirements are sounique it is critical that focused Mars balloondesign technology goes forward if Martian aerobotsor balloons are ever to be considered viablemission candidates in either the Mars Explorationor the Discovery Programs.

Advancement of autonomous navigation andcontrol technologies allow the aerobot to takeadvantage of unique vantage points along its windblown path. Such technology enables the aerobotto point cameras to predetermined targets ofopportunities when the system is within closeproximity. Furthermore, these technologies could

trigger the release of small scientific packages atplaces of high interest.

The MABS design has demonstrated that anaerobot can fill the gap in data resolution betweenorbit and ground. In addition, aerobots are directlyrelevant to the Mars Exploration program goals andto future Mars surface sample return missions.

Finally, as an aerobot ranges over the Martiangeography for purposes of discovery, people willidentify with it and become engaged in theexploration as explorers themselves. Students willfollow its progress across the planet and anticipatenew sights, new discoveries and new mysteries ofMars.

Acknowledgments

The authors would like to acknowledge theprogram support of J. Cutts, C. Elachi, S. L. Miller,D. Shirley at JPL; H. Needleman atNASA/GSFC/Wallops; and J. Blamont, and J-M.GuilbertofCNES.

We would also like to thank all the members of theMABS study team for their efforts including G.Ball, J. A. Jones, L. Lee, R. Levin, D. McGee, J.Olivieri, B. Parvin, E. Slimko, T-L. Tran, A.Vaisnys, J. Wall, E. Washington, L-C. Wen, and A.Yavrouian at JPL; H. Cathey, S. Raque, M. Said,W. Schur, and J. Simpson at NASA's GSFCWallops Flight Facility; B. Clark, B. McCandless,M. McGee, S. Price, N. Smith, and B. Sutler atLockheed Martin; and J. Cantrell of Utah/SDL.We also appreciate assistance of Dr. R. Haberle atNASA's Ames Research Center in the developmentof Mars balloon environmental models and the useof his Mars General Circulation Model to simulateaerobot trajectories across Mars. Thanks also toJim Schaeffer NASA's Ames Research Center forrunning the GCM codes and providinginterpretations of the seasonal wind patterns.

Special thanks are given to Dr. J. Blamont for hisinstigation of the Mars balloon concept, A. Vargasof CNES/Toulouse for his help in understandingthe lessons learned from the French/Soviet Marsballoon program and for his insights into theMABS design, and to Dr. R. Greeley of ArizonaState University for his help in establishing thepotential role of an aerobot in the Mars ExplorationProgram.

A portion of the research described in this paperwas carried out by the Jet Propulsion Laboratory,California Institute of technology, under contractwith the National Aeronautics and SpaceAdministration.

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