-
4th Annual AIAA/USU Conference on Small Sah~lIites Utah State UnlYlirait)l, Logan. Utah 28-30 August 1990
DESIGN OF AN ELECTRICALLY-PROPELLED ASTEROID RENDEZVOUS PROBE LAUNCHED - AS A SECONDARY PAYLOAD
- Ralph 0 Lorenz· Department of Aeronautics and Astronautics University of Southampton Southampton UK - Abstract
This paper summarises iii design study undertak~m as a final Yl:'ar proJect for the author's a.Eng, in Aero5pace SystQms Engineering. A spacecraft design IS outlined for a vehicle to perform iii. rendezvous 'With a Near Earth Asteroid after being launched as a - secondary payload into Geosynchronous Transfer Orbit. The 360 kg (dry) spacecraft wOlJld use iii biproP4'llant chemical propuh.ion system to manoeuvre In. and escape from, Earth orbit. Interplanetary manoeuvring ""Quid be accomplished ""Ith an ion - propulSion system.
V!'rious mission and system design aspects are described.
• The author IS no"" with the European Space Technol09Y and Resurch Centre (ESTEC1. Postbus 299. 2200 AG Noord""ijk. The Neth.dands -
Introductl0n - The obJectlve of thm study was to determlne what kind of mission could be accomplished with a modest spacecraft using electric propulsion laUnched as a secondary payload. In particular it e~amined ~ Near Earth Asteroid - Rendezvous (NEAR) mission. USlng a variety of advanced technologies anc techniques, a highly sophisticated and capable spacecraft can be constructed of a size SUltable for launch as a secondary payload. - This paper summarIses the re~ults of the study report (ref.1) WhICh IS 155 pages long. There is insufficient space here to repeat the details of "calculation methods and other background inform~tion, so if applying these results to other mISSions, take care! -
As an e~ercise. the mission was based around a number of components and technologies under development or manufactured in the U~. - Launch The Ariane launch vehicle was selected as a baseline for a number of - reasons:xt IS European, and therefore perhaps the least poiltically senSItIve of available boosters; also it is launChed frequently (about 10 tImes/year), and has a long hIstory of collaboration with secondary payloads (AMSAT OSCAR 10,13; UoSATs 3.4. Microsats; Viking etc.) -
The spacecraft IS designed as a 'satellite porteur' (ref.::) - built around the 1920mm/937mm launch adaptor (Just as the AMSAT Phase 4 spacecraft.) A cylindrical volume allocation (2.26m dia by 1m high) was o1.liIiiumed. No fi:led - launch mass limit was taken, as thIS WQuld depend on the prlmary passenger on the launcher. A dry mass in the region 300-350 kg was aImed for, makIng the wet ma5S around 700-800 kg. This mass can be accommodateo by uprating - the Ariane booster, WhICh is available 1n a number of verSlons wlth GTO payload masses from 1900kg to 4200kg. (see figures !!:,3) - The mission analysis 1S effectively dlvided into two main parts: the Earth orbIt pha~e and the interplanetary cruise. lnltially, the Earth orbIt phase was r&go1.rded as a formallty, the intentlon beIng to escape immediately using - a SOlld motor into a clrcular heliocentric orblt of radius 1 AU and use the electrlc propulsion system from there. However, when the effects of uSIng a slightly overslzsd motor were investigated, it was found that great - improvements in V-infinity (the hyperbolic departure velOCity) can be obtained by slightly increasing the magnItUde of the escape burn due to a non-lInearIty in the celestial mechanics (see figure 4.) - However, for the e~cess V-infinity to be useful, it must be in the correct directIon. This neceSSItates manoeuvring in Earth orbit - hen~e requirIng a restartable propulSIon system (i.e. not a solid.) Additionally, Oilnce the launch date is specihed by the pnmary passenger on the launCh vehicle, the spacicraft may have to wait in earth orbit for some months before departure. GTO is a most unfavourable orbit in which to spend such a period aerodynamlc drag, torques and heating, and most Importantly, the high radiation dose all degrade th. mission perform~ce. Therefore the craft
- -
manoeuvres into .in expanded Or-blt (say 89(00)(500 km) for thll!! phasIng penod as suggested in (,-ef.3.) This orbit reducs$ the radiation dose considerably: it also breaks the escape burn into two smaller burns so that a smaller motor can be used. The perigee is raised slightly for safety but kept faIrly low to maximise the V-infinity performance of the escape burn. An inclInation change manoeuvre would be performed just after entering the waiting orbit to control apsldal precession and nodal regression of the orbit. Inclination would be adjusted agaIn a few orbits before departure to orient the departure direction.
The Earth orbIt manoeuvres are ~ummarised in figure 5.
Mission - Deep Space Phase
A number cf potential targets were examined, selected by eye from the TRIAD (Tucson Revised Index of AsterOId Datal TIle (ref.4l and other asterOId mission studies (refs.5,6,7l. Delta-V requirements were evaluated assuming Hohmann transfer~ (not strictly correct given that some of the manoeuvres are conducted with a low-thrust propulSion system, but accurate enough for the purpose of this study.)
Thus two velOCIty changes, dslta-V-one and -two, are required for the misSion. These can be traded off against each other by varYIng the fraction of the Inclination change that is performed at rendel-yeus (parameter alpna.) The delta-V requirements fer some of the easIer targets are plotted 1n hgl.!re 6 with the effect aT varying alpha shown; actual figures for alpha=O.7 are given in table 1.
- The delta-V-one requirement is met by the hyperbolic departure velOCIty, plus a velocity increment from the ion propulSion system if necessary just after departure; delta-V-two is met by thii ion propulsion system In deep space.
The nomln.l targets selected are mildly inclined (less than 10 degrees to the ecllptic). and have apheli~ 1.BAU or closer. The three prlncipal targets - are Eros, Anteros and Bacchus. The capabllities of a combined chemical-electric and a chemIcal-only spacecraft of the seme launch mass are shown in figure 7. By comparing figures ~ and 7 ,it is seen that a chemical-only vehlcle is only just capaDle of performing a rendezvous mission with the easIest-reached target, 1982 HR. A chemical-electric vehicle~ as proposed here, has a much larger performance envelope and allows far more flexibility in target selectlon. Gi.ven the launch as a secondary payload, this flQxibility IS vital.
An all-electriC yehtCle would be capable (at first Inspection) of performlng mlSSlons of even Wider scope. Howevl!'r~ in this case, the thruster llfetlme, rather than propellant supply, becomes the limiting parameter. Escape from GTO using electric propulSion would be a nightmare from an attItude control point of vleW, wouid Incur substantial radlation damage, and would take over one year (USlng the two UK-to thrusters with 2kW of available power.) Further, since a slow, spIral escape leaves the craft with a hyperboliC departure velocity of zero, the actual propellant mass savlng for a typical III1$S10n (wlth delta-v-one equal to about 4 km/s) will be very small. - -
Accordingly, the combined chemical/electric mission is retained.
ConfIguration The configuratlon driver~ are principally the volumetric constraint imposed - by the launch vehlcle, and the cruise pointing requirements of the ion thrusters, solar arrays and communlcatlons antenna. A number of posslble configuratlons were trIed but reJectea on the grounds that they reqUIred complIcated mechanlsms. -
)he configurahon that finally evolved has a body-fi:~ed payload, chemical motor and communlcatlons antenna. The vehlcle 1S an octagonal prism, sized to fIt the envelope 1n flgure :. A cruciform solar array 15 deployed by - unfolding four four-panel 'wings' from the sides of the probe. These arrays are body-fixed and do not artlculate about any rotahng jOlnts. (In some respects, the probe resembles the Mariner Mars probes.) The only mechanisms - used are booms for two of the Instruments. gImbals for the ion thrusters, and one-shot mechanlsms to deploy the solar arrays and antenna feed. ChemIcal PropulSion, - Imtially the concept of the mission was to use the electric propulsion system to the full. Since a spiral escape from GTO 15 not easy, It was inltlally proposed to use a SOlId motor to e~cape from the Earth. However, - ~nen the escape manoeuvre was considered in more detail, bearing in mind that the launch date and orbit parameters would be dictated by the primary passenger, a more fle~:ible propulsion sy~tem. capable of multiple burns, was reqUIred. ThIS necessltated a storable bipropellant system: thiS comprises a 500N motor, (the Leres 1 engine manufactured by Royal Ordnance) I'll th four titanlum 1(1) litre tanks with capillary propellant management devices
The noz::le of the 500N motor projects from the centre of the 937mm adaptwr ring an the nomlnal antl-sun face of the spacecraft. The volume allocatIon for the nozzle would have to be negotIated with the launch authol'"lt/' The tankage is mounted inslde the adaptor rIng, such thaL the chemIcal propuls10n system \e;~cept for the attitude control thrusters! can be kept separate from the rest of the vehicle untIl final integration. The ITiodular - construction of the probe 15 shown in figure 8.
AnalYSlS Indicates that for the spacecraft mass considered here, 500 N IS adequate thrust to prOVIde Sufflcisnt delt3.-v aver the permItted thrust arc. Lower thrust levels would extend the required arc to the pOInt where pOIntIng losses durlng the spln-st.abI11sed burn would degrade the ,USSlon performance unacceptaoly.
El~ctric Propulsion
The electrIC propulslon system is built around two UK-tO :(enon '00 thrusters (ref.a). In order to eli.llinate disturbance torques generated by - 4 -
thruster mIsalIgnment and movement of the spacecraft centre of mass during the ffilssion the twa thrusters are mounted on gimbals which allow the thrust vector of each to be tIlted +/- 4.2 degrees. Additionally thrust vectoring over a wideI"' range ( ... /- 20 degrees) is possible about one a:ds by dlfferentially throttling the two thrusters - this allows optimal thrustIng WhIle maintainIng earth-pointing. This configUration also allows at least a minimal mi 3 sion to be accomplished even if one of the thrusters fails.
The thrusters are run an Xenon stored near its critical pOint: the 50kg propellant capacity requires 100 litres of storage; unfortunately the tanks used far the bipropellant do nat have t~e pressure ratIng requIred land indeed anether tank of this si=e would be difficult to fit In the avaIlable volume) so the Xenon would be stored in two or feur sll',aller tanks.
'The Ion thrusters are run with an acceleration voltaJ;;e of 940V - thIS gl'19S a speCIfic Impulse of 3160 seconds. The thrusters can be throttled over a range 10-:25 mN each, WIth corresponding power consumption of 275-001) W. Some of thIS power is diSSIpated in the thruster power condltl00l,'g Unlts~ wnich have an area on the anti-sun face of the craft to radIate away thIS waste heat.
Power
The power demands of the spacecraft are somewhat hIgher than usual tor Its SIze thIS is clearly due to the power required by the Ion propulSIon system. To slI1lplify construction (and to avoid potenbal elastIcl ty problems c.f. Hubble Space TeleSCOpe) the solar arrays are of the rIgId toldout type. ThlS also allows the possibility of withst~nding shocks due to lanCIng on the asterOId surface if this IS to be attempted.
The four arrays are each four panels long, each panel equal in area to one of the eight sides of the spacecraft. Taklng a packIng fractIon of ~.8~ thl~ - gIves room for ':040') 2:::cm cells. Gallium Arsemde cells were s=lec~ed for their hIgher effIciency and Illore Important~v for thetr higher r~dIat1Qf1 tolerance.
The arrays prOVIde 1940W of power at BOL and 1AU; allOWIng tor pOIntIng canditloniny losses ,il;nd a 107. margin, this leaves 140()W~ permi.ttIng both Ion thrusters to be run at full power continuously. At 1.S AU, taken to be the design maximum solar distance, the available power IS about 4i"::W suffIcient to run one ion thruster at low power and run the other spacecraft systems.
In Earth orbit, the arrays are folded against the sides of the craft. WIth only the outermost panel exposed to the sun. Depending on solar aspect angle, up to 140W can be generated. However, power requ1rements 1n E~rth orbit are low.
When the power storage requirements (about 450 W-hours) are conSloereo, drtven mainly by ecllpse duration 1n orbit about the e~rth or the asterOId, it is found that Nickel-CadmIum batteries are uncomfortably maSSIve, :;0 Nickel-Hydrogen cells are selected instead. The cells are arranged In two batteries of twelve celis. Each battery has a DC-DC converter to boost the battery voltage (nomInally 14V) to the bus voltage (28V). Should any cell - 5 fail, it can be switched out of circuit and th~ battery wIll contInue to function: the lower battery voltage can be accommodated by the DC-OC converter. ThIS combi~atlon of redundancy (: batterIes) and the graceful - degradatIon of the batteries themselves offers e~{tremely hIgh relIabIlity . • Since SUbstantial e:,:cess power is developed by the arrays if the Ion thrusters are not fIrIng, a shent dissipator is reqUIred: thIS IS sItuated on the anti-sun face of the spacecra~t. Communicationli - The targets e):amined for the mission have aphelia up to 1.8 Astronomicai Units, so in principle, the Earth could be up to 2.8 AU dIstant from the spacecraft. However, this would make the line-of-sight pass very close to - the sun resulhng In a poor link. An arbItrary limIt on the sightllne-to-sLln angle in the ecliptic plane of 20 degrees was set: this f1::es the ma:nmLlm design communications range at ~.5 AU.
The primary communications link uses a 1.65m parabolIC S-cand antenna~ SIted inside the 19:0mm aoaptor nng an the nOmlnal sunSlde of the craft. A feee is mounted or. an arm which swings Into positIon (using a one-shot mechanIsm) after the escape burn. During thQ maIn cruise phase, the sun-spacecraft earth phase angle is found to be less than about 30 degrees, so that the craft can be held Earth-pOinting with a 1O~~ or lesss COSIne lass in solar power generatIon. -
Using a lOW transmitter and convolutional slgnal encoding, the link can support 400 bIts per second downlink at a dIstance of :.5 AU. L,nlo: - performance is lncre~sed at shorter ranges, hIgher transmItter power and with more elaborate coding. The nominal link budget is shown in table ~ During the Earth orbIt phase an omnidirectional antenna is used. -
Immediately after Earth de~arture. the sun-probe-earth angle IS found to be much greater than 3(1 degrees: the spacecraft is then sun-pennted a.nd communIcations ~ont1nue with the omni antenna. The ornni antenna is capable - of supporting 400 bps until the Earth-probe distance exceeds 0.1 AU, by WhICh time the phase angle has reduced to a point where the hIgh-gain dlsh can be used wi thout incurri ng a:{cessi ve sol ar array Ini spOlntl ng 10SS9S. - Attitude Control
DUrlng the Eal'th orbit phase~ the spacecraft is spin-stablllsed. Even wit~ the solar arrays folded up against the sides of the craft~ the moments 0+ Inertla are suited for spin stabilisatlon, which serves to Mold the atc.l"tude rIgid durIng motor fIrings.
Attltude determlnatlon is by sun and earth sensors in this phase. In the event of a failure of any of th~se components, some aadIt.anai attItude informatton could be provided by the star sensors (If the spIn rate, ~robably 10-15 rpm, is low enough for the sensors to c~pe) and a magnetometer carrIed as payload. Spin rate adJustments and slew manoeuvres are performed by monopropellant hydra.:lne thrusters. - After depa.rture, t!'le craft is despun, the arrays are deployed and sun-lac:"
6 - - -
achieved. Thereafter the craft is three-axis stabIlised.
The dom1nant torque in the cruise phase turns aut to be the ion thrusters. While the hydrazlne reqUirement to combat thiS torque IS not exceSSlve (although large) the number of thruster firings would exceed the rated life of the thrusters, so gimbals have to be used to orlent the lon propulSIon thrust vector to elimlnate disturbance torq~es. The gImbals are based around an e:nstlng antenna pOinting mechanism (by BADG) wlth a mass of 4.: kg and a power consumption of 8W (ref.9). Remember that gimbaillng WIll only be necessary when the ion thrusters are hring, when 'there w111 be plenty of power available.
The use of reaction wheels was Investigated: these would ellminate propellant consumption during limit-cycling in cruise and sleWIng for imaging (the payload is body-fixed) at the target asterOld. Whether a reactlon wheel sUlte is less massive than the propellant otherwise reqUIred depends on the number of slews reqUired (the limit-cycllng fuel requIrement 1S very low) - break-even occurs at about 2000 s1e ... s. A flgure of ::::500 slews was taken, maklng wheels slightly better in mass terms: however~ this ~rea needs more careful anal ysi s and whether the 5kg mass savi ng I s worth the addltll:lnal comple:dty 1S decatable. Nominally. then. no wheels are earned.
Attltu~e determInation durlng cruise and rende=vous is accompllshed wlth star sensors (based on CCD cameras- ref. to) and sun sensors. -",ddl1:ior,al Information, 1n the event of sensor blinding, IS provlded by attltude references which could be based on a number of low-mass technologies (e.g. gas gyros. fibre-optic gyros or solid-state gyros.)
It is acknowledged that the attitude control of a vehlcle on thlS sort ,~-t mission is lndeed complex and will demand substantlal on-board processlng capability. It is conjectured that the development effort involved would be considerable.
Payload
While the main alm of the study was to see what can De achleved wLth avallable technology, clearly the misslon 1n practlce Will be dr!'1en to " large e:{tent by scientific objectives. These are to determlne
1. Global charactenstIcs: shape. si=e, mass distnbution, rotation a,:lS and penod Surface morphology: regolith structure and depth~ craterlng record e~c. ~. Elemental and Mineral CompOSItion 4. Dust and Plasma enVIronment; magnetic field if any
A survey of previous asteroid misSlon proposals Irefs. 6,11,1:,1:::, 14i was undertaken to see what kind of instruments are carrIed to mee't these objectives. The instrument payload selected as baseilne resembles those of other missions:
1. Optlcal C.9.meras based on CeD Imagers, ilrobably falrly rudlmen'tary by planetary e::plorAtlon standards. The e:{act design of the. optlCS was oeyono the scope of thIS study and would depend on the final mission plan. No'te that whIle provlding hlghly lmportant sCience data from orb1t about the asterOld. the cam~ras would also be ~ed to .S»ISt 1n the nav1gatlon oi the
7 spacecraft just prior to rendezvous. The cameras are fixed to the s.de of - the spacecraft,. which 15 slewed round to aim them • 2. Gamma Ray Spectrometer - to identify the elemental composition of the - asterOld: resolution 15 of the order of the spacecraft-asteroid distance. The instrument is deployed at the end of a boom to reduce contamlnatlon of the measurements by the spacecraft material itself. - 3. Dust Detector - to investigate the dust partlcle ditribution about the asteroid. The detectors could use any spare 'real estate' on the spacecraft surface. - 4. Magnetometer to investigate any possible remnant magnetism. ThiS Instrument too would be mounted at the end of a boom to mlnlmlse contamination. - 5. Imaging Infrared Spectrometer - to identify surface mineral dIstrlbutlon. This lf1strument IS optlonai ~ subject to available mass! power and data budgets. LIke the cameras, the instrument is body fixed. - 6. Secondary Ion Mass Spectrometer - to study surface mineral and elemental composlt1on'. An e;,~periment of this typs was Eown (but nat used before the craft was lost) On the Soviet Phobo$ 2 mIssion: an Ion beam IS used to - sputter materIal from the asteroid surface for analysis an-board In a mass spectrometer. Here an ion thruster would be used to generate the primary Ion beam (ref.15). This 1nstrument, too, is optional. - The science obJect1ves met by the varIOUS instruments are 5ummarlsed 1n table ~. - Although international collaboration is attractIve, most of these instruments could be sourced In the UK. In partIcular, the teams at Kent (Dust Detectors) and ImperIal College, London (Magnetometers) have excellent reputatIons. - Groundstation - Overall programme costs can sOQr if missions are not managed correctly. It is proposed that most spacecraft operations be conducted autonomously, ana that only one groundstatl0n is used. Use of ESTRAC,< or DSN is not conSIdered 1n ttns report (except in contingenCIes) as they would be hideously - e:~pensive and, In the antICipated timeframe of the mIssion - the late 1'?9(ls - these facilities w1l1 be In great demand for the many planetar,' and salar terrestrlal phYSiCS programmes taklng place. - A single~ ded1cated ground5tatlon would be expenslve to construct from scratch. However, there is a 10m ~ntenna and a control faCIlity at the Rutherford Appleton Laboratory In Odol"dshlre (figure 91: these have been - unused Slnce the US/Dutch/UK IRAS mission. Although the dish 1S old, and construct10n worK has severed the cables connecting the d1Sh to the controi room, the station could be restored to operation far about.f: mllllon. - Conclusions A .... iable asteroid rendezvous mission can be conducted with a spacecraft -
8 -
launched as a secondary payload. In the case examined here, a combIned chemIcal/electrIc propulsion system IS requIred to provIde adequate performance: 'or other mlSSlons or launCh orbits either all-chemical or al1- electrIC vehIcles may be better.
Whil e the vehi cl e deSCribed is 1 ndeed campI e:( and uses many new technologIes, no 'magic' IS required. Unlike a number of other proposals for planetary mISSIons using small electrIcally-propelled vehIcles (refs.:.16i the technologies and components suggested are In manufacture (or at least on the workbench) rather than being slmply on paper.
While programme costs are notoriously dIfficult to estImate. launch as a secondary payload more than halves the p~lce of the launch. Whlle mast components 1'1111 have flown at least once befo~e the IIIlSSlon takes place, the technologies a~e still ~elatively new and favou~able te~ms could p~obably be negotiated. The ~eally difficult pa~t to estimate lS the cost of ope~atlons. softwa~e development and systems integration. If the prog~amme IS managed appropriately. wlth universlties undertaking muc:h of the work, overall p~ogramme costs could be kept down to the point whe~e even the Uf:'o mlght be able to affo~d the mission.
Acknowledgements
The autho~ Wishes to record hlS thanks for the support of RAE Farnborough and the Di!pa~tment of Aeronautl c:s and AstronautICS, Un! verSl ty of Southampton. The support and enc:ou~agement of Dr Graham SWlne~d, who supervised the proJect, is gratefully acknowledged.
Reference!!
1. R D Loren:: 'DeSign of an Electrically-Prop@lled Aste~oid Rende::'/ous P~obe Launched as a Secondary Payload' Bacnelor of Enginee~ing Thlrd Year HonoL\~s P~oject, Southampton Uni ',ersi ty, May 1990
2. Arianespace presentation at the IAA Small Satellite seSSlon at the 40th Congress of the lntern~tional Astronautic~l FederatIon, Malaga, Spaln, 8-15 Odober 1989
3. A E Pietrass 'TraJectory Design for an Ion Drive Asteroid's Rendezvous Misslon Launchi!d into an Ariane Geostationary Transfe~ Orblt' AIAA-84-2058. AIAA/AAS Astrodynamics Conference, Seattle, August 1984
4. D F Bender 'Osculatlng Elements of the Asteroids' In T Gen~els led) , Asteroi as' , Universlty of ArlZona~ 19i8
5. CO Lau and N D Hulkower 'On the Accessibilty of the Near Earth Asteroids' AAS 85-352 AAS/AIAA Astrodynamics Specialist Conferenc;:!, Vall, Colorado, August 1985
6. J E Randolph and D A Baker 'MIssion Candidates for the Second Planeta~y Observer' AAS 85-306 AAS/AIAA Astrodynamics Speclalist Conference, Vall, Colorado, August 1985
7. F Bonneau 'Catalog of Asteroids or Comets In the Vlclnity of Earth and
9 - Mars' IAF-88-381 Presented at the 39th Congress of the InternatIonal Astronautical F;deratlon, Bangalore, India, October 1988
8. D G Fearn, A R Martin and P SmIth 'Ion Propulsion Research and - Development in the UK' IAF-B9-274 Presented at the 40th IAF Congress, Malaga, Spaln, October 1990 - 9, H M Briscoe 'MQchanlSms' chapter in Space Technology Course Notes. Southampton UniverSity. Easter 1988
10. P Butler and A Steln 'Deep Space Missions for Small Satellltes' - Proceedings of the 2nd AlAA/USU Conference on Small Satellites, Utah State U~lverslty, Logan, Utah, September 1988 - 11. A BraMlc 'The JOlnt WorkIng Group Asteroid M15510n' Advances In Space Research Vo1.5 Na.2 pp.97-106, 1985 12. Vanous 'Asteroid E;{ploration Mission (ASTEREX)' Assessment Study No.2 - ESA-SCI(81) 1. ESA May 1981
13. Various lVESTA - A Misslon to the Small Bodies Ot the Solar System' ESA SCI(88)6, ESA October 1988
14. K J Cole, H Feingold, J McAdams, M L Stancati and J R French 'Small Spacecraft for Low-Cost Planetary MiSSIons' Proceedings of the 3rd Annual AIAA/USU Conference on Small Satell i tes, Utah State Un! '1ersi ty, Septemtler - 1989
15. R D Lorenz 'Remote Asteroid Surface Analysis Using E;;haust from an Ion Thruster' IAF-90-312, to be presented at the 41st Congress of the International Astronautical Federation, Dresden, GDR. October 1790 10. W K Daniel and V C Gordon 'Small Satellite Design for Inner Solar System - E:-:p 1orati on , Proceedings ot the 3rd AIAA/USU Conference on Small Satellites. Utah State UniverSity, September 1989 - - -
-
10 - ) I I ) I
, .. , ...... f""'" "'''' c.m l"-:ll .,," ''''' """ 3'2
...... ,,~ M"~ ... ,,'
l .... O,...."" __
/', I
.8 p,.~,~" ",,,"") I l • ~i,. --~--"------r -, ' !---~---"jl ------T ,!o IH'O ----, \~OO f'" 2 o '" TABLE IIELTA-V kEqU! irEI1ENTS FOir RENIIEZVOU:> ~ .. ll'h&~u '1 J ...... ~,_ 01' .~"t&II.""",\",c. ,(I;UI;) "0. NAltE , , Perl Api, dVl dV,u, 1235 SCIlOItItIA 1.910 O. 154 2S.0 2 ;'0 6129 61119 12328 1 1685 tORO 1.368 0.436 9.4 o. "'n 4939 2:nl '/310 1741 WRIGHT 1.709 O. 110 21.4 '" 5114 5101',,1 11031 1139 ,\TAltl 1.94.7 0.255 n . . "4~i 2.44'" son 3"H3 9286 1621 IVAX . 00' 0.3,:n ,. , 1. 12 2 60 6199 IS03 Ha2 IS19 CLEHEMCE I. 936 o 095 I'). 3 1. 15 2. 12 54'n 5539 11032 IB62 ArOLLO 1.470 0.560 0.65 5598 2S01 B 102 1566 lCAItUS .017 ° 621 "," 0.19 '" IZOI12 89S5 21031 2063 BACCHuS .OTI 0_ :;49 0.10 " 3342 3;)0/ 6649 2100 RA-SHALOII 0.832 o 4:H 15.8" 0.4"1 " 4';00 6811 11318 IS43 AMTEROS 1.430 o 256 ,. , , '"eo 4250 2015 6325 20GI ANZA 2.260 O.5J·1 3 ., .00'" 3.4"l "1303 -1910 2606 SENECA 2.480 0.586 15.6 ." 3. ~13 !l409 1845'" 10254 EROS 1.460 0.223 '" , I. 79 n?1 2105 1026 '" ALI NIIA 2.500 0.555 9.2 " 3.89 1983 Iwa 9191 1665'" CERBE~IIS .ono o 4()"I >c O.'" Sf! !>H 5131 4913 10051 2102 TANTALUS 1_290 0.298 (,4.0 0.91 I b I L2436 14'140 21116 IB63 AMTlNOIJS 2.260 o 61)6 '" , O.fI') 3 83 9599 :1126 13025 '06' ATEN O.StiG o !O3 Hl.9 0.19 3145 (JOGO HHI~l 1221 AHOIt .920 o 4-J4 " , I. 09 2.'" "/5 6659 2008 tI!:>w, 2101 ADONIS I.B'15 o '164 , , 0.44 3.31 "l1!:>6 3213 IOJ6tl 1620 GEOGRArliOS .245 0.335 " , 0.83 , " 4390 3288 '/616 193"1 liS .639 '-' (J21 e , 1).62 2.66 634L 2451 tl19S 1950 DA 888 1).':>02 12 2 ,eo 2.53 !HI!-' 2163 W,/I 1980 AA . tl~ I U 141 , , 1 O'j 2. '/3 6314 Til 1041 '-....."" .... ;;: • .. <- ...... T." ..... 1954 II.A o 11"1 0.345 , , o ~1 1044 'j:j(;2 640/ (~'-T" .. e"" .... tll\...F .. Itf\,,o 1964 QA 0.9Ga 0.168 , , 0.53 I "'45 3691 4026 8518 n"",.. ~''''..... <,,,,,,;5: I1E~CU~Y o Jill, I! ,,(If, , () 31 0 41 f,jl:lCJ 124~)cJ IfJtlfltl VENUS o "/26 U.U()I o 12 0 /.J 249J 3084 SSHI • f,()(, I HARS ':;"" " ()~U , 44 L. '14 3191 21(;'1 A sea, adjU' ilX,S 'All' E ecce .. " ,<:lty LII<:I,,,.,,o" ,,, .. ,-11)"'" Ide~J ferL r"r"'~I"'" (Alii Aph Al'hel,o" <11111 dVI - E~"h 1"'1"" 'u'" ,1<-1", v (~'SI dV2 - ~""de2.v""s d.,I'~ v '11,-;1 107. .01 ,,,,-1 ,"~"o" "h"ng~ po' dV,,,, '" .I~"!' ~l"" (' I I ) I I I I I I I I ) I I I I