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Deorbit Maneuver and Targeting Strategy for Unmanned Mars Landers

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Deorbit Maneuver and Targeting Strategy for Unmanned Mars Landers The Space Congress® Proceedings 1968 (5th) The Challenge of the 1970's Apr 1st, 8:00 AM Deorbit Maneuver and Targeting Strategy for Unmanned Mars Landers W. J. Pragluski Martin Marietta Corporation D. Marquet Follow this and additional works at: https://commons.erau.edu/space-congress-proceedings Scholarly Commons Citation Pragluski, W. J. and Marquet, D., "Deorbit Maneuver and Targeting Strategy for Unmanned Mars Landers" (1968). The Space Congress® Proceedings. 3. https://commons.erau.edu/space-congress-proceedings/proceedings-1968-5th/session-6/3 This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. DEORBIT MANEUVER AND TARGETING STRATEGY FOR UNMANNED MARS LANDERS by W. J, Pragluski, D. Marquet Martin Marietta Corporation, Denver, Colorado SUMMARY Several deorbit maneuver strategies for unman­ certainties. This aspect of the error analysis ned Mars soft landers are evaluated in terms of pro­ leads to the definition of an entry corridor and pulsive efficiency, targeting capability, communica­ selection of potential landing areas for 'early tion link geometry, and sensitivity to system uncer­ missions. tainties. These strategies include (1) minimum de- orbit impulse, (2) minimum entry condition uncertain­ The final aspect of the analysis includes the ties, (3) minimum variation in the lander antenna effect of targeting flexibility on the post land­ aspect angle during entry, (4) minimum communication ing communication link characteristics between the range from the lander to the orbiting spacecraft dur­ lander and orbiting spacecraft. The influence of ing entry. The selected maneuver strategy is a com­ both aspects of the error analysis on the post bination of the above which restricts orbiter lead landing communication link is discussed. angles to the range between 0 deg and -10 deg. The analysis covers a range of elliptical orbits with ' INTRODtJCTION periapsis altitudes of 500 to 2500 KM and apoapsis altitudes of 10,000 to 20,000 KM. The nominal The analysis of deorbit maneuver strategy orbit which is selected in order to best satisfy only be meaningful if it is made within the bound­ all mission considerations is 1300 KM by 12,500 KM aries imposed by scientific objectives of the mis­ (hp x hA) . sion, system capabilities, and constraints imposed on the mission profile by subsystem design, consid­ Mission considerations include orbit orienta­ erations. To establish a deorbit maneuver strategy tion of the orbiting spacecraft for planet surface all of the flight operations in- the vicinity of mapping, landing site location between 15 to 30 Ma r s - mu st be considered i n c lud i ng or b i t, s e 1 ec t 1on 9 degrees from the terminator for entry TV imaging, deorbit impulse size.,, guidance subsystem accura­ and orbit characteristics which insure a 50-year cies, co iranu n i c a t ion subsystem 1 i nk geome t r i e s , and, orbit lifetime and non-occultation of either the last, but hardly least, the ability to place the Sun or Star Canopus for 30 days after encounter. lander at scientifically interesting landing areas. These constraints, coupled with the uncertainties The analysis presented 'below touches on many of introduced by trans-Mars navigation uncertainty and the more important aspects which must be considered. orbit insertion maneuver uncertainties, are used in It includes propulsion subsystem considerations for the definition of minimum targeting flexibility re­ orbit insertion, orbit trim, and deorbit, the error quired to land at a preselected location (latitude analysis associated with these maneuvers, targeting, and longitude) and the regions of Mars which can be capability, and communication subsystem link geom­ selected for landing. The selection of a landing etry 'characteristics. These analyses lead to the site after surveillance from orbit where as much selection of a deorbit maneuver strategy and range targeting flexibility as possible is desired is of suggested orbits. also considered. Before proceeding with the analysis, a brief Two aspects of the error analysis are consid­ description of the total mission, profile is present­ ered. The first deals with the range of orbits ed below to provide, the proper perspective. This relative to a preselected nominal which might be includes the more important mission, constraints experienced. The sources leading to this range assumed in this analysis. Only the operations in uncertainty include cruise navigation, orbit in­ the vicinity of Mars are presented. In any detail* sertion maneuver, and orbit trim maneuver. The deorbit maneuver strategy must be capable of com­ MISSION PROFILE AMD COMSimilTS pensating for these orbit uncertainties if a pre­ selected landing site is to be acquired. Total Mission Profile The second aspect of the error analysis deals The mission profile considered is shown In with the flight conditions at entry and landing Figure 1. The planetary vehicle, orbit ing space­ site acquisition accuracy. The uncertainty sources craft and lander , is inje.ct.ed onto a trans-Mars • considered here include orbit ephemeris determina­ trajectory from an Earth parking orbit* A Sun- tion accuracy, deorbit maneuver, and atmosphere un­ Ganopus attitude control system is employed during 6.3-1 the cruise mode. At least two midcourse maneuvers representative opportunity to discuss some of the might be required during the trip of about 200 days orbit insertion and orbit geometry considerations. duration. Based on Earth tracking and the desired The energy contours for a Type I mission are pre­ landing site, the orbit insertion command is calcu- sented in Figure 2. The relationship between _ lated and sent to the planetary vehicle. At a pro­ Earth departure energy, C~ 9 and Mars approach grammed time the orbit insertion maneuver is execu­ energy, V « 3 is illustrated. Two constraint bound­ ted, At least four orbits are assumed to determine aries are also shown. The lower limit on encounter the ephemeris of the resulting orbit from Earth- date is established by limiting the ¥„„ to less than based tracking. An orbit trim maneuver may be re­ 3,25 KM/SEC. The declination of the departure asymp- quired to compensate for the off-nominal orbit has been constrained tote, § HE® to be less than resulting from approach and orbit uncertainties, 35 deg from a range safety consideration. The cross Potential landing sites may be viewed from orbit hatched region is where the most favorable pay load and a site selected. Four to five days after orbit margins occur,- This is the combination of insertion, depending on the desired landing site which results in near maximum pay loads in orbit in?rill longitude, the lander is separated from the orbiting about Mars, Three circled points of interest are 'Spacecraft, The lander orients itself for deorbit shown. Points (1) and, (2) axe used for the discus­ firing during a ha If -hour coast phase required to sion of periapsis shift requirements presented below. separate the two vehicles by at least 300 meters, Point (3) is used for the example of occultation con­ The deorbit engine 'is fired and the lander then, straints. coasts to the desired atmospheric entry point. Be­ tween .deorbit firing and entry the lander is reor­ Constraints iented to provide a nominal entry angle of attack of aero degrees, A ballistic entry is followed by The mission constraints considered in this parachute deployment at an altitude of about 18,000 paper are given in Table I. below, along with a brief ft above the surface. Vernier ignition and para­ comment on their source and influence on the mission. chute jettison occur at an altitude of about 5000 ft, Vernier shutdown occurs 10 feet above the surface NEAR MRS. PROPULSIVE PHASES and the lander free-falls to the surface* The de­ tails of the terminal phase system are discussed in faother paper and are not considered here* The orbit insertion phase is discussed in terns of the impulse required and the associated errors. Only insertion at the periapsis of the approach fhe 1973 Mars mission is selected 'here as a hyperbola 'is discussed in this section., The influ-* mius i IHElJJgEICK 1», Orbit lifetime of at least 50 years Non- sterilisation of the orbit­ Size of orbits ing spacecraft 2. Orbit insertion A V fl¥ less than Reasonable weights in orbit $ize of orbits and 1,75 shifts allowable 3% Inclination of orbit to Hartian Orbiting spacecraft mapping Allowable landing site lati­ equator greater than 30 deg mission tudes 4* * £Ton*occultation of either the Sun Orblt ing spacecraft at11tude Al lowable Incl ina tions a t or Ganopus for 30 days after en­ control reference and, power Mars counter source 5* Landing site located 'between, 15 to Entry TV imaging Allowable range of targeting 30 deg from the terminator parameter 6* Orbiter sub-periapsis between sero Orbiter TV imaging Allowable range of targeting to 45 deg from the terminator parameter and. periapsis shift 7* Orbit insertion, in view of Gold- Mi s s ion operat ions. Time of arrival; orbit orien­ stone tation 8* A VD < |,00 I/SIC leu sonable lander weights Allowable range of targeting parameter t, fu* < 3*25 KM/SBC 'Reasonable weights in orbit Allowable range of orbits and perlapsIs shlft ence of periapsis shift requirements on Z-iV T is maneuver impulse of 150 nips is assumed here with maneuver uncertainties of 5% of impulse (3 <T ) later. Impulsive maneuvers are assumed. discussed both normal and parallel to the nominal AV. The period error after trim will be .081 hours Orbit Insertion Impu1se. The maximum Av O.I. (one sigma) for the 1000 x 15 ,000 KM orbit used as considered is 1.75 KM/SEC (Table I).
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