98-0614 from gto to the planets.pdf

8/14/2019 98-0614 From GTO to the Planets.pdf

1/9

FROM GTO TO THE PLANETSRobert Gershman, Paul Penzo, and Alexandria Wiercigroch

Jet Propulsion Laboratory - California Institute of TechnologyAbstract

Most Ariane 5 launches will provide slots forauxiliary payloads p to eight payloadsat 100 kg apiece. Recent analyses has established that these slots can be used to send smallscience payloads to Mars, Venus, or a range of near-Earth and main-belt asteroids. This paperdiscusses the Ariane 5 configuration, mission analyses, and launch scheduling issues associatedwith these opportunities and describes some applicable candidate mission concepts.

1. IntroductionThe Ariane 5 launchvehicle is designed tocarry large commercial satellites to orbit manyof which are delivered to a geosynchronoustransfer orbit (GTO). Like the Ariane 4 beforeit, the Ariane 5 will, in most launches, provideopportunities to orbit smalluxiliarypayloads. Ariane 4 has carried 22 suchpayloads each weighing up to 80 kg to Earthorbit. For Ariane 5 using the Ariane Structurefor Auxiliary Payload (ASAP) the maximummass is increased to 100 kg. Customers payone million dollars (US) for each slot to coverArianespace integration costs. Recent analysishas established that the Ariane 5 ASAP can besynergistic with state-of-the-art smallspacecraft concepts to providexcitingopportunities for planetary missions.Potential targets include Mars, Venus, theMoon, near-Earth asteroids, and main-beltasteroids. These planetary opportunities arebased on a launch to GTO followed byapplication of the mission design strategydiscussed below.

2. ConfigurationThe ASAP platform, which canbe used inmost Ariane 5 launch configurations, has eightmounting ingsplaced uniformly around thecircumference of the aft part of the payloadarea. Each can support a 100-kg payload in anarea 80 cm (height) x 60 (width) cm x 60 cm(depth). An important proposed variation,currently under study by Arianespace, is theso called banana or twin configuration,would provide for mounting a single payloadon two rings,doubling the upper limit onpayload mass and substantially increasing theavailable volume.3. Launch Opportunities andPerformance3.1 MarsThe opportune launch periods from Earth toMars appear about every 26 months and last2-4 weeks. This is the time intervalwhenescape energy sminimumand the deliverymass to Mars is maximized for a given launchvehicle. There may e two launch periods

-1 -


2/9

about a month apart, with one having a longerflightime to Mars than the other. Theshorter flight time trajectories are called Type1 and the longer Type 2.To adapt these short launch periods into theAriane 5 launch schedule, it is necessary to beaboard and launched into GTO many weeksbefore the Mars injection opportunities. AfterGTO launch, a wait time in space (in GTO oranother orbit) is acceptable, but a majortrajectory problem arises. This is how totransfer from the highlyccentricndspecifically oriented GTOo the correctescape direction and energy on a date whichfalls within the two week Mars launch period.In a very limited way, this can be done with asingle maneuver, performed with a solid motorfor example. A more general and very flexiblemethod, called the three-bum strategy, existsusing lunar and Earth flybys.For single or multiple burn options, the finalmaneuver must occur within the two weeklaunch period. If there is a single burn it mustoccurnear the equatorial plane to avoid acostly dogleg. Also, to avoid a severepropulsion penalty, the bum must be withinabout 15 degrees of GTO perigee. These tworestrictions limit the availability of Mars tocertain years and limits the length of the GTOlaunch period. Mars 2003 is available for thesingle burn option since certain Mars arrivaldates require a near quatorialnjection atEarth. In fact, bychoosing he Mars arrivaldate judiciously, it is possible to develop aGTO launch period of about 2 months for2003. In contrast, Mars 2001smarginallyavailable with Mars 2005 being not available.A brief ook at following years indicates aproblem for the single burn option until Mars2013.

To alleviate this problem, a three-burnstrategy has been developed (see Figure 1)which can deliver an auxiliary payload releasedin GTO off to Earth escape and on a pathwhich will deliver it to Mars. In addition, thisstrategy is general enough to allow transfer toalmost any solar system body, while acceptinga GTO date whichmay span wo or moremonths.

not to scale

To Mars

Figure 1: The Three-Burn StrategyFor Mars, the strategy requires the payload tohave a multi-burn capability with a totalvelocity increment in excess of 1200 m/s.Also, it muste capable, with groundassistance, of precise navigation, and a thrustcapability of about 0.5 g s.Normally, the injection from low Earth orbit(LEO) to Mars requires 3600 m/s or more.,however, the highly elliptic GTO provides2400 m/s, reducing the auxiliary payloadenergy equirement to about 1200 m/s. Thethree-bum strategy requires up to anadditional 300 m/s for trajectory shaping,depending on the GTO launch period desired.

-2-


3/9

The primary purpose of thehree-burnstrategy is to re-align the GTO major axis fromits near equatorial orientation to that whichwould be required inLEO if a single burn wereused to inject the payload directly to Mars.This is accomplished byusing a single ordouble lunar flyby which returns the payloaddirectly back to Earth so that it can performthis escape maneuver. This lastmaneuver isusually about 800 d s and is timed to occuron a predetermined date favorable toheEarth-to-Mars transfer. The last acceptableGTO launch date necessary to accomplish thethree-burn trajectory profile is about 30 daysbefore this final Earth flyby maneuver.An example of the earliest and latest date forthe GTO launch is presented in Figure 2 forthe Mars 2001 opportunity, as viewedfromthe ecliptic north pole. The total AV requiredfor thispportunity using a three-burnstrategy is shown in Figure.imilarly,Figures 4 and 5 present the three-burn velocityrequirement for GTO launch opportunities for2003 and 2005 respectively.Propellant requirements for navigation of thethree-burn strategy are relatively small because

x 1 0 5 km Moon Orbit

-6 -4 -2 0 2 x 10 5 km

of the long ime nterval between injectionfrom GTO to he inalmaneuver to escapeEarth, considerable tracking times available toreduceheavigation requirementsconsiderably. The first burn, injection romGTO, is the most critical. The payload, oncereleased from the Ariane may beracked for aslittle as 1-2 orbits, and a 1 0 - d s burn error atthis time and only the burn magnitude isimportant) wouldequire about a 50-m/scorrection one day out. The burn itself will beabout 730 d s . The correction may be reducedby breaking the maneuver into two parts, firstgetting into an intermediate orbit, say with a 5-10 dayperiod. Then, further tracking, ndbetter understanding of engineperformance,could educe the final burn error. Once thespacecraft is on the high ellipse toward itsapogeebeyond the Moon, which will takefrom 15 to 35 days, the remaining burns can beminimized for theexisting trajectory. Trackingwill be good, and correction requirements areestimated to be about 25 d s . A finalcorrectionwillbe equired after the escapemaneuver of about 460 m/s is performed. AnRMS estimate of all maneuvers required comesto about 50 m/s, and this is expected to beprimarily dependent on engine performance,not tracking or attitude control.

X I O 5 km Moon Orbit

- 8 I . . . / . : . . . . . . . . . . . . . . . . . . . .: . . . .-2 0 2 J 6 x 1 0 5 k m

Figure 2: Earliest and Latest Ariane Launch or GTO to Mars 2001)-3-


4/9

1 4 4 0 I I I I I I I Ianh Escape Flyby 2TMarch 2001l M M a r s Arrlval Date 3 September001

C7 = 9 km2/secZP

1380

2 1m3

lJUI1320

Bt 7 3 0 01280

1 2 0 10 ZO 30 40 50 10 10 800 100GTO Launch Dates from November 14,2000)

Figure 3: Mars 2001 Required AV

13601340

vti 1320E9 i 3 wE 12805 1260 240312201200

1 2 09 06 0 3 0 0GTO Launch Dates from May 28,2003


. 1 5 01 2 0 P O - 10 30GTO Launch Dates from September I 2005

3.2 VenusSimilar mission design considerations apply toVenusxcept that the Earth departureasymptote tends to be opposite rather thannear the sun, leading to the double lunar flybystrategy shown inigure 6 . Thecorresponding performance chart for the 2004Venus opportunitys shown inFigure 7.Note that if a AV of 1360 d s e c is available,the Ariane launch period can be more than 4-months long.3.3 Main Belt AsteroidsMain belt asteroids caneeached bylaunching to Venus as described above andthen using a Venus gravity assist. The AVrequired for several potential targets is shownin Table 1.3.4 Other TargetsOther targets of interest to sciencerereachable from GTO by more straight-forwardmeans, .e., a single burn at perigee. Theseinclude the Moon and near-Earth asteroids.The ASAP is also considered to be an excellentresource for space technology demonstrations,particularly those involving Earth re-entry.3.5 Performance SummaryTable 2 summarizes the ASAP performancefor many of the cases discussed above. Thefigure of merit is the useful payload deliveredto an Earth escape trajectory that intersectsthe orbit of the target body. Results are give nfor both of the baseline ASAP 100-kg slot andfor the banana assuming a 180-kg capability.Carrier spacecraft mass estimates arefrom acompanionpaper by Ken Leschly of JPL2.


-4-


5/9

A V 1 = 735 m sm an 27 MayAV2- 198mb3C an 21 JuneAV3- 471 nu=on 24 July

10

AV2 = 251dsecon 21 July . .

Figure 6: Earliest and Latest Ariane Launch or GTO to Venus 2002)

tbs by Paul

Figure 7: Venus 2004 Required AV

4. Launch Scheduling for PlanetaryOpportunitiesSending an auxiliary payload to Mars or Venusnecessitates changes in the Ariane manifestingprocess. Mars opportunities involve 3-monthlaunch periods with abrupt termination times,after which the velocity increment from G T Oto Mars rises prohibitively. While a launch toGTO is very likely in the 3-month period, itwill not be known until about one year fromlaunch who the main passenger will be. Thiswill necessitate doing any integration analysis

withmore han one mainpassenger.Also,provisions mustbemade for delays to helaunch vehicle or he main passenger occurringnear the timeof aunch that would slip theMars payload out of its launch period.Provisions houldbemadeoransfer theMars payload to another GTO launch (if any)within the period. Another fallback would beto revert to a Type 3 or 4 trajectory with laterEartheparturesandmuchongerlighttimes).

-5-


6/9

Table 1: Examplesof Main Belt Asteroid Opportunities

Asteroid

teresPallasEgeriaAmphitriteHestia

Vesta

No. Radius I DeDarturearth2 263 Jul-2002104 A ~ ~ 2 0 0 2

I I Ilight ApproachEarthC Carrier AV Time Velocity Rang8(km2/sec2) 3.4.6.2.61.9 (AU)Wsec)years)km2/sec2)

10.33.60.6 3.2.4.01.54.5.6.5

7.81.7.2.8.5.9 2.1.8

.5 1.4

334ar-2004

1.8.1.0.4.79 100 Mar-2004 2.3.6.9.4.7626ar-20043.6 6.7.3.4.317ar-20043.0.4.7.4.155ar-20043.6.8.5.4.7

Table 2: ASAP Performance SummaryV e n u s

/ 0 4

91 4 0 0

LARGE. 6

-6-


7/9

5. Mission ConceptsMostof the work on potential planetaryutilization of ASAP has been targeted at Marsand some of the more promising concepts arediscussed hereand summarized in Table .5.1 Aerobot ConceptsThe exploration of Mars with aerobots is onetype of a mission concept under considerationas an Ariane 5 auxiliary payload. The use ofaerobots in Mars exploration has severaladvantages over landers, rovers, and orbiters.Whereas landers and rovers have limited rangeand imaging perspective, he use of an aerobotmovingabove the Martian surface at severalkilometers in altitude willhave substantiallygreater mobility and area coverage capability.Furthermore, the collecteddata taken byanaerobot will bridge he resolution gap betweendata obtained by an orbiting satellite and thatfromround-limitedover. Thedevelopment, testing, and use of aerobots toexplore the Martian surfaceelies first ondemonstratingeyechnologiesnd thenprogressing to missions withreatercapability. Two mission concepts underconsiderationre the AerobotTechnologyExperimentnd the SolareatedMontgolfiereNind Sail Rover.

Figure 8: Aerobot ExperimentThe AerobotTechnologyExperimentseeFigure 8 will be a low-cost flight experimentdesigned to demonstrate the key technologiesfor the exploration of Mars with aerobots andto demonstrate the feasibility of an airborneplatform at Mars for scientific exploration.The payload, weighmg between 25 to 32 kg,is inserted into the Martian atmosphere via anaeroshelldelivery. Next, adrogue chute isdeployed which extracts the main parachutethat is connected to the mainballoon. Theballoon nflates n 60 seconds or less. Theballoon will carry an instrument payload thatincludes a miniature wide angle camera, spotspectrometer, and magnetometer. The missionlifetime is approximately 7 days.

Table 3: Mission ConceptsMicro-satellite Micro-banana

Entry/Descent/Landing Payload Quantity 8 Massayload Quantity & MassLcaLa m 3 mm r r Unit TotalEDL lnst r P/L/L EDL l ns t r P/L/L =

f 2 d '3 d (k g )kg)k g ) Quan. (k g ) (kg)kg)kg ) Quan. (k g) 2Seismic Net 5 2 7 4 2 8 s2 7 2 1 4xlimate Net 3 1 4 1 6 s1 4 8CScience Lander x x C 1 2 8 2 0 2 4 0 SScience Lander x x x c 3 5 1 05 1 4 5 ABalloonExperiment x x x c 3 5 1 05 1 4 5 AWindRover x x X C 6 4 1 0 1 1 0 6 4 1 0 2 2 0 sNano-rover x x x c 4 1 5 1 0Com.Orbiter X 1 0 0 1 1 0 0 - 1 2 0 1 1 2 0 s1 0 0 1 1 0 0cience Orbiter - ~ 1 2 0 1 1 2 0 sCarrier: c = cruisecraft (spinner) S = sciencecraftBananamounting: S = symetricalabout hrust axis A = asymetrical

c 2 :u l v g g : : s . E Unit Total 2

scout 6 4 1 0 4 4 0 s4 1 0 2 2 0x

-7-


8/9

The SolarHeatedMontgolfiere/Wind SailRovermission concept (seeigure 9) isdesigned to take advantage of the fact that aontgolfiere, or hotairballoon,canmakemultiple soft landings and deliver a payload,such as a wind sail rover, to the surface. Thepayload weighs approximately 10 kgndincludes a 4-kg instrument payload. Afterinsertion into the Martian atmosphere, thesolar heated balloon delivers a wind rover tothe surface. The rover will carry out soil androck sample analysis andwillacquire highresolution images over many kilometers. Theballoonwill then rise to a low altitude andacquireigh-resolution stereoscopic surfaceimagesover a 12-hour period, the expectedlifetime of the balloon. The mission ifetimeof the wind rover is approximately 30 days.5.2 Network ConceptsThere areeveral network-type missionconcepts that benefit rom a very ow-mass(less than 10 kg),0-cmeroshellesignsimilar to that being developed for Deep Space2. In these concepts, the objectives are carriedout with two or more probe deliveries to theMartian surface. Targeting of the probes tospecific interesting sites or tocover a givenarea depends on the missionciencerequirements. For example, the site selectionmay be based on geology,mineralogy, orimaging requirements.A long-term (approximately 20 years)investigation of thenatureandvariability ofthe Martian climate system is considered tohave high priority in the Solar SystemExploration Road Map. A network of probeson the surface of Mars will define the climatenaturend variability. In addition, these

Figure 9: Solar HeatedMontgolfiereMind Sail Roverprobes willollect data on the thermalstructure ofheed planets atmosphereduring probe entry. A current probe designincludes a 4.7-kg entry, descent, and andingpackage and a 2.1-kg instrument within a 40-cm diameter aeroshell. Ideally, the probes aredistributed across the Martian surface forclimatemonitoring purposes. Each probeoperates autonomouslyndommunicatesthrough a relay satellite. Options existorlong-term data storage overeveral years in theevent that no orbiters available.Deployable Camera or Spectrometer

Figure 10: Probe with Various PayloadsAvailable at http://eis.jpl.nasa.gov/

roadmap/site/slide-destiny.htm1

-8-
http://eis.jpl.nasa.gov/http://eis.jpl.nasa.gov/


9/9

Another mission concept will nvestigate theMars surface with various payloads. In thislow-mass payload design,esshan10 kg,there exists the opportunity tosend two ormore probes to the planets surface, each witha different instrument payload such as adeployable camera or spectrometer (see FigureIO). The current mission design will require arelay orbiter for communications.The seismicity and internal structure of Marscan also be investigated with a network-typemission. In this concept, multiple probes aredelivered to Mars, impact its surface, ndlodgebeneath the surface.These probes arevery low-mass, approximately 3.3 kg and aredesigned to self-right themselves during entry.The ideal mission design includes network ofmicroprobes and may utilize a modified NewMillenniumrogram Deep Space 2microprobe.Each probe willommunicatethrough a relay orbiter and will lastapproximately one year.5.3 OrbitersMany of the mission concepts underconsideration require the use of a relay orbiterfor communications. Ideally, such a satellite inorbit around Mars willenefitll Marsexploration missions. One such design s theMultiple Relay Satellite (see Figure 1l), a 100-kg Mars orbiter that uses the ASAP launchcapability. The current design ncludes a 5 -year in-orbit lifetime.

6. ConclusionAriane 5 auxiliary payload capability to G T Ocouldbeavailable as early as mid-1999andwille routine by 2003. This capabilitycoupled with the trend toward smaller, smarterscience spacecraft provides excitingnd

Figure 11: Mars Relay Satellitechallenging new opportunities for planetaryscientists and missiondesigners.7. AcknowledgmentSpecialcknowledgment is due to JacquesBlamont of CNES who first suggested the useof ASAP for planetary missions and createdthebanana concept. We also acknowledgethe contributions and support many of ourcolleagues at JPL, CNES, and Arianespace,particularlyJohnBeckman who started JPLdown this trackandKimLeschly who hasturned these ideas into credible spacecraftconcepts.The work described in this paper was carriedout at the Jet Propulsion Laboratory,California Institute of Technology, undercontractwith he National Aeronautics andSpace Administration.8. References1. Ariane V ASAP Users Manual publishedby Arianespace, 1997.2. K. Leschly & G. Sprague, CarrierSpacecraft Using the Ariane5 GTO SecondaryLaunch Opportunities, this conference.

-9-

98-0614 from gto to the planets.pdf

Documents