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ASE 387 N ssion Analysis and Design Mals InSitu Propellant Production Sample Return Msslon Architecture Andreas Mogensen Monday 3 May 2004 Mars TnSitu Pronellant Production Mi ion Andrea Morgen en Executive Summary This mission demonstrates the insitu propellant production ISPP techniques and the launch technologies that will be a vital part of any manned Mars mission The ISPP technology is demonstrated in the context of an unmanned Mars sample return mission The mission is scheduled for launch during the October 2009 launch opportunity A Delta 7925H launch vehicle will launch the 1135 kg spacecraft The spacecraft which consists of a lander a rover and an Earth Return Vehicle ERV will follow a Type 2 minimum energy transfer trajectory The spacecraft will arrive at Mars in August 2010 and perform a ballistic entry using parachutes retrorockets and airbags Surface operations will include 90 sols of sample collection using a rover and 210 sols of insitu propellant production A total quantity of 205 kg of ethyleneoxygen bipropellant will be produced by reacting a small amount of imported hydrogen feedstock with carbon dioxide extracted from the Martian atmosphere In April 2011 when propellant production is complete the ERV will launch from the surface and enter a 600 km parking orbit around Mars When the launch window for Earth return opens in July 2011 the ERV will follow a Type 2 transfer trajectory back to Earth where it will arrive in July 2012 with approximately 500 g of Martian rock and soil samples Mars lnSitu Pronellant Production Mi ion Table of Contents Section 1 0 Introduction 20 Mission Synopsis 40 60 21 Mission Pro le 22 Mission Systems Mission Goals and Success Criteria Driving Requirements 41 InSitu Propellant Production 42 Core Samples 43 Other Science Goals 44 Planetary Protection Key Trade Studies 51 InSitu Propellant Production Techniques 52 Mission Architecture Mission Systems Description 61 Lander 611 InSitu Propellant Production Plant 612 Power System 613 Communications System 62 Rover Andrea Morgen en Page 20 20 Mars TnSitu Pronellant Production Mi ion 70 80 90 621 Science Instruments 622 Communications System 63 Earth Return Vehicle 631 Mars Ascent Vehicle 632 Earth Entry Vehicle Mission Pro le 71 Mass Schedule 72 Trajectories 73 Mission Phases and Timeline 731 Launch 732 Trans Mars Cruise 733 Mars Entry 734 Surface Operations 735 Mars Ascent 736 Trans Earth Cruise 737 Earth Entry Operational Concerns 81 Mars Planetary Protection 82 Earth Planetary Protection References Appendix A List of Abbreviations Andrea Morgen en 21 21 22 23 25 27 27 28 29 30 30 32 33 33 33 34 35 35 35 37 38 Mars TnSitu Pronellant Production Mi ion Andrea Morgen en 10 Introduction The president s new space exploration initiative has emphasized manned exploration of the moon and Mars The initiative will combine the synergies of robotic and manned missions to create a full program of exploration of the moon and Mars Too often in the past it has been a question of either robotic or manned missions It is important to recognize that some missions are more suited for robots while others are more suited for man The current eet of robotic rovers landers and orbiters to Mars has been hugely successful and has returned a wealth of information that has once again changed our perception of Mars Most recently the Mars Exploration Rovers have found credible evidence for the existence of a large body of water on Mars at some point in its past If the evidence is con rmed then the exploration of Mars will move from the current phase of searching for water to a new phase of searching for life The search for life either extinct or living is a mission best suited for man and his analytical abilities It is evident that any manned Mars mission in the near future will rely on insitu propellant production ISPP on Mars to produce part or all of the fuel needed for the return journey to Earth The use of ISPP in manned Mars missions significantly reduces the launch mass on Earth to the point where a heavy lift vehicle comparable to the Saturn V can be used and assembly of parts in low Earth orbit avoided ISPP is the crucial technology which makes a manned Mars mission affordable and achievable in the near future It was rst proposed by Dr Robert Zubrin as part of the Mars Direct Plan and later incorporated into the NASA Design Reference Mission The primary goal of the mission is to demonstrate this crucial technology while the secondary goal is to return a sample of Mars to the Earth The mission would encompass a lander a rover and an Earth Return Vehicle ERV Once on the surface of Mars the ISPP plant would begin production of the ethyleneoxygen bipropellant which would be used as rocket fuel for the return journey The ISPP process is a fairly simple chemical reaction which uses hydrogen feedstock brought from the Earth to convert carbon dioxide from the Martian atmosphere to ethylene and water The water is electrolyzed into oxygen and hydrogen which is recycled back into the chemical reaction Once fuel production is complete and a suitable Mars sample has been collected the ERV will launch from Mars and return to Earth Demonstrating this crucial technology on Mars will be a vital step on the road to manned Mars missions Mars TnSitu Pronellant Production Mi ion Andrea Morgen en 20 Mission Synopsis The mission synopsis will introduce the important ideas and elements of the mission A more detailed and in depth discussion of the mission elements will follow later in the report The following synopsis is dependent on the completion of the trade studies described in the Key Trade Studies section of this report It represents the best mission architecture at the current time 21 Mission Profile The mission is nominally scheduled for launch from Cape Canaveral during the October 2009 launch opportunity It is proposed that an intermediateclass launch vehicle such as the Boeing Delta 2 will launch the spacecraft on a Type 2 transfer trajectory to Mars The spacecraft which consists of a lander a rover an Earth Return Vehicle ERV and a small amount of liquid hydrogen feedstock will arrive at Mars in late August 2010 The proposed landing site is currently Terra Meridiani but any equatorial landing site between 5 0N and 5 0S latitude is possible On arrival at Mars the spacecraft will perform a landing that is very similar to the landing used on both the Mars Exploration Rover MER mission and the Mars Path nder mission First the spacecraft will use an aeroshell and a supersonic parachute to decelerate through the Martian atmosphere Then retrorockets and airbags will be used to cushion the lander at surface impact After touchdown the lander will immediately start a small chemical processing plant The plant will pump carbon dioxide from the Martian atmosphere into a reaction vessel where it will be reacted with a small amount of imported hydrogen in what is known as the reverse water gas shift RWGS reaction The reaction products include water carbon monoxide and hydrogen The water is condensed out and electrolyzed to produce oxygen which is stored and hydrogen which is recycled back into the RWGS reaction The remaining carbon monoxide and hydrogen is further reacted to produce ethylene which is stored and water which is recycled back into the electrolysis unit After approximately 210 days of operation sufficient ethyleneoxygen bipropellant will have been produced to fully fuel the ERV While the chemical processing plant is operating the lander will deploy a surface rover The rover will conduct a minimum of 90 sols of surface operations such as surveying the local terrain extracting core samples from selected rocks and conducting insitu measurements of the local environment so that the context of the core samples is well understood The rover will deliver the core samples back to the lander where they are placed in the sample return capsule in the ERV Once surface operations have been completed the ERV will launch from Mars and enter a 600 km parking orbit around Mars When the launch window for the Earth return opportunity opens in late July 2011 the ERV will follow a Type2 transfer trajectory to Earth On arrival at Earth in July 2012 the ERV will release the Earth Entry Vehicle Mfr In m M inn EEV and perform a de ection maneuver to miss the Eanh The EEV will follow a ballistic reentry trajectory and land in the continental United States The sample retum capsule in the EEV will be retrieved and taken to a secure containment facility where the p39 39quotL quotAn 39 39 39 39 39 quot 39 L inFigurel Am Jul on Jan Aer Jul 01 usquot nor Jul on m nor rut on Jan ll ll ll ll ll llllll ll ll ll ll 2am H 2 am Jzn l l l l l 2315 512 EanhLauncn 7 us Cm Surhce Openhulls ERVL MarsDrhll H m Eanhcmlsz Ea hRemm V Figure 1 Missiun Timeline 22 Mission Systems The mission will require the use ofthree major ight system elements This section will present ahighlevel overview ofthe general characteristics of these s stems in order to ilitate the description of the mission design A detailed description of each ight system elemem 39 39 quot 39 39 39 39 39 quot 39 39 entitled Mission Systems Description Lander At 450 kg the lander is the largest ofthe three ight system elements Its design derives from the lander used in the MER missions though to accommodate the larger payload the shape of the deck is a modi ed hexagonal rathert an tr39 1 e maximum dimension ofthe lander in cruise con guration is 27 m allowing the lander to t within the payload fairing ofthe Delta 2 launch vehicle Two ofthe six sides ofthe hexagonal position on the deck The remaining four s39des are covered in solar arrays All six sides e owered by a motorhinge arrangement that is capable of uprighting the lander to its correct position a er touchdown on Mars The payload is mounted on the deck ofthe lander and includes the ERV the rover the ISPP plant and the communications system The ERV is stored horizontally within the A I 439 439 39 39 39 39 39 lidin rail 39 39 length ofthe ERV This system will raise the ERV into avertical position priorto laimch rover is positioned over the ERV for the duration of the Mars cruise The rover Mars TnSitu Pronellant Production Mi ion Andrea Morgen en returns to this position when transferring samples to the ERV The ISPP plant is located on one side of the deck and the communications systems on the other side Rover The Athena rover is a smaller version of the MER rovers The sixwheeled rover has a mass of approximately 70 kg and measures 131 m long 110 m wide and 150 m tall in its deployed con guration Ref 8 The wheels are 20 cm in diameter and are attached to a rockerbogie type suspension which provides a ground clearance of 25 cm The rover is solar powered and can achieve a top speed of about 1 meter per minute Total accumulated driving distance during the initial 90 sol mission is about 15 km The rover carries the Athena science payload which consists of a panoramic camera a miniature thermal emissions spectrometer a miniature corer an alpha proton xray spectrometer a Moessbauer spectrometer and a Raman spectrometer Sample collection will be performed using the miniature corer which is mounted on the bottom of the rover The baseline procedure will involve driving the rover over a selected rock drilling into the rock and then depositing the core sample into a small cache box next to the drill When appropriate the rover will return to the lander and regress to a xed position above the Earth Return Vehicle on the deck of the lander The samples will then be transferred directly to the sample return capsule in the Earth Entry Vehicle Earth Return Vehicle The Earth Return Vehicle ERV consists of a Mars Ascent Vehicle MAV and an Earth Entry Vehicle EEV The MAV is a 19 m long and 06 m diameter twostage rocket capable of returning the collected samples to Earth Both stages of the MAV are liquid fueled using ethyleneoxygen bipropellant produced on Mars The rst stage is used to launch the ERV into a 600 km parking orbit while the second stage is used to perform the transEarth injection burn The second stage also provides the utility systems such as power communication and guidance navigation and control of the spacecraft for the duration of the Earth cruise The EEV which is the payload of the MAV is a small capsule approximately 03 m in diameter It consists of a secure and isolated sample return capsule SRC surrounded by a crushable honeycomb matrix and a heat shield The SRC holds the sample cores collected on Mars and provides the structural strength and integrity to protect the samples The heat shield protects the SRC during the ballistic reentry while the crushable honeycomb matrix absorbs the landing impact The EEV also includes a beacon which guides the search and retrieval party to its location The ERV is stored in a horizontal position on the deck of the lander between the wheels of the rover for the duration of the Mars cruise During surface operations the rover will transfer the collected samples directly to the SRC in the EEV by driving onto the deck of the lander and positioning itself over the ERV Mars TnSitu Pronellant Production Mi ion Andrea Morgen en The ERV arrives at Mars unfueled except for a small amount of hydrogen feedstock As surface operations continue and the hydrogen is consumed the ERV is slowly fueled with ethyleneoxygen bipropellant The ERV is encased in a thermal unit which provides sufficient refrigeration to liquefy the propellant After approximately 210 days the ERV is fully fueled Just prior to launch the ERV is raised into a vertical position using a system of sliding rails Mars TnSitu Pronellant Production Mi ion Andrea Morgen en 30 Mission Goals and Success Criteria The mission serves as both a technologydemonstration mission and a scienti c mission As such the mission serves two distinct purposes which are re ected in the overall mission goals The primary technological goals of the mission are 1 To demonstrate the feasibility of insitu propellant production by producing a sufficient quantity of ethyleneoxygen bipropellant on Mars using insitu resources 2 To launch a vehicle from the surface of Mars and back to Earth using the propellant produced on Mars The primary scienti c goals of the mission are 1 To return approximately 500 grams of sufficiently diverse rock and soil samples from Mars 2 To conduct a detailed geographical geological and atmospheric survey of the landing site area In support of the primary goals the secondary goals of the mission are 1 To operate a rover on Mars for a minimum of 90 days while traveling a minimum combined distance of 15 kilometers The technological and scienti c goals of the mission are related Clearly the scienti c goals of the mission can not be achieved without also achieving the technological goals As a result the following success criteria have been established in order of relative importance Criteria Result Approximately 500 g of rock and soil samples are returned to the Complete success Earth The ERV is successfully launched from Mars but fails to reach Partial success Earth with the rock and soil samples run 1 391 J quot produced but the Partial success ERV suffers launch failure The rover conducts 90 days of surface operations traveling a Partial success total combined distance of 15 kilometers Table 1 Success Criteria Mars InSitu Pronellant Production Mi ion Andrea Morgen en 40 Driving Requirements The goals of the mission impose several requirements which are driving the mission design The main driving requirements are described below 41 In Situ Propellant Production The lander must include a chemical processing plant that is capable of producing sufficient ethyleneoxygen bipropellant to launch the ERV with the core samples from Mars and return them to Earth The plant must be simple and robust in other words it must be capable of operating for the duration of the mission without any maintenance and it must be capable of surviving launch from Earth and entry and landing on Mars In addition the power requirements of the chemical processing plant must be minimal while the rate of propellant production must be sufficiently high to limit the overall mission time The propellant also imposes storage requirements such as refrigeration to minimize boiloff 42 Core Samples Core samples totaling approximately 500 grams must be returned to Earth The core samples should be of sufficient diversity to characterize the geology of the landing site area Photos and additional insitu measurements should be collected to properly understand the context of the core samples The core samples should be isolated and kept below 50 0C No Earth biological contamination can be returned in the core samples in order to avoid ambiguity in the analysis of the material 43 Other Science Goals The rover should include a suite of instruments to properly characterize the landing site area in terms of geography geology and atmosphere Although smaller than the Mars Exploration Rovers the rover should be capable of many of the same types of measurements In addition the rover should be capable of performing an extended mission of exploration after the core samples have been collected 44 Planetary Protection Both forward contamination of Mars and backward contamination of Earth must be avoided Mars must be protected from contamination by Earth organisms which could potentially travel with the spacecraft to Mars and contaminate the landing site and the core samples In addition Earth must be protected from the uncontrolled release of any unsterilized Martian material Mars TnSitu Pronellant Production Mi ion Andrea Mogen en 50 Key Trade Studies Several trade studies are currently underway which will determine the optimum mission design 51 In situ Propellant Production Techniques In choosing an appropriate insitu propellant production technique several requirements were imposed on the production process and the propellant combinations The requirements are listed in Table 2 below Process Propellant Requirements Simple and robust No maintenance required High specific impulse Shock and vibration tolerant Able to Storable on Mars with minimum withstand launch and landing loads refrigeration requirements Low power requirements High production rate to minimize mission time Table 2 ISPP Process and Propellant Requirements Several potential chemical processes were considered The possible production processes and propellant combinations are summarized from Ref 11 and listed in Table 3 The rst potential ISPP process listed in Table 3 involves the direct electrolysis of water into its constituent components of hydrogen and oxygen The electrolysis process is very simple and robust and the technology is mature Rugged electrolysis units capable of withstanding depth charge attacks are commonly used on nuclear submarines The hydrogenoxygen propellant combination is also very suitable However the process is not suitable for a Mars sample return mission because of the dif culty of obtaining insitu water Although we now believe that water is relatively abundant on Mars it would either have to be condensed out of the atmosphere mined out of the permafrost or drilled out of a subsurface reservoir None of these options are feasible The second potential ISPP process involves the direct dissociation of atmospheric carbon dioxide into carbon monoxide and oxygen using zirconia tubes The carbon dioxide is heated to about 1000 0C where it partially dissociates into carbon monoxide and oxygen The oxygen is separated out by the zirconia tubes which are porous to oxygen transport The process is simple but not robust The zirconia tubes are brittle ceramic tubes and are susceptible to cracking In addition the carbon monoxideoxygen propellant is not optimal The propellant requires a very high ame temperature and only has a modest speci c impulse of about 270 s Mars TnSitu Pronellant Production Mi ion Andrea Morgen en The third potential ISPP process involves the Sabatier reaction in which hydrogen and carbon dioxide are reacted in two steps to produce methane and oxygen The rst step involves the reaction of atmospheric carbon dioxide and a small amount of imported hydrogen to produce methane and water The water is electrolyzed to produce hydrogen which is recycled and oxygen which is stored The process is simple and robust and the methaneoxygen bipropellant is suitable However the process is not the optimal process ISPP Process Propellant Comm ents Direct electrolysis of insitu H20 HZOz Not feasible Direct dissociation of insitu COZ COOz Not feasible Sabatier reaction of H2 and C02 CH4Oz Feasible Reverse water gas shift with ethylene QH4Oz Optimum production Reverse water gas shift with methanol cHgOHOz Optimum production Table 3 Possible ISPP Processes and Propellant Combinations The fourth and fifth ISPP processes listed in Table 3 are the optimum processes They are very similar in that they both involve the reverse water gas shift RWGS reaction The RWGS reaction produces carbon monoxide hydrogen and water using atmospheric carbon dioxide and a small amount of imported hydrogen The hydrogen and carbon dioxide are further reacted either in the presence of an iron FischerTropsch catalyst to produce ethylene or in the presence of a copperzinc catalyst to produce methanol The water is electrolyzed to produce oxygen which is stored and hydrogen which is recycled back into the RWGS reaction Both ethyleneoxygen and methanoloxygen are suitable propellant combinations The two processes are both suitable for the Mars sample return mission The advantage of producing ethylene is that it minimizes the amount of hydrogen that needs to be transported to Mars to support insitu propellant production whereas the advantage of producing methanol is that it minimizes the power requirements of the insitu propellant production system Ref 11 It is thus the recommendation of this report that the mission uses the ethyleneoxygen propellant combination for the following reasons 1 The amount of imported hydrogen is minimized which minimizes the launch mass on Earth 2 The speci c impulse of ethyleneoxygen is 376 s which is higher than the speci c impulse of methanoloxygen which is 353 s Mars TnSitu Pronellant Production Mi ion Andrea Morgen en It should be noted that the imported hydrogen could in theory be acquired on Mars by electrolyzing insitu water However the processes involved in acquiring the water on Mars either involve mining the permafrost or drilling into a subsurface reservoir Neither is suitable for a sample return mission and it is more cost effective to import a small amount of hydrogen feedstock 52 Mission Architecture There are several potential mission architectures that could successfully be used to meet the mission requirements These are summarized in Table 4 Mission Architecture No of Launches Comments Direct return from Mars 1 Simplest Mars orbital rendezvous with the original 1 Difficult spacecraft Mars orbital rendezvous with a second 2 Difficult spacecraft Table 4 Possible Mars Sample Return Mission Architectures The mission architecture involving a direct return from Mars is the simplest architecture from a technological standpoint For this reason it is the architecture adopted for this mission The entire spacecraft is landed on the surface of Mars and sufficient propellant is produced on Mars to launch the ERV directly back to Earth Although this architecture requires the maximum amount of propellant to be produced on Mars the ISPP process is fully testable on Earth This means that although the ISPP technology represents new hardware the technology can be matured to the desired level prior to mission launch by prolonged testing The two remaining mission architectures both rely on Mars orbital rendezvous MOR In the first case the mission uses a single launch to place the spacecraft on a Mars transfer trajectory The spacecraft consists of an orbiter which remains in Mars orbit and a lander rover and MAV which land on Mars Sufficient propellant to launch the MAV into low Mars orbit is produced on Mars The remaining propellant for the transEarth injection TEI is provided by the orbiter which rendezvous with the MAV in Mars orbit The amount of propellant required to perform the MOR and TEI is significant Consequently the mass of the orbiter is significant which means that a heavylift vehicle is required to launch the mission in a single launch The third mission architecture avoids the need for a heavylift vehicle by splitting the mission into two separate launches One launch sends the lander rover and MAV to the surface of Mars while the second launch sends the orbiter into Mars orbit Apart from this the mission architecture is identical to the MOR architecture described above Mars TnSitu Pronellant Production Mi ion Andrea Morgen en The MOR mission architecture is technologically difficult An orbital rendezvous by two unmanned spacecraft around another planet has never been tried before and unlike the ISPP technology the MOR technology is not fully testable on Earth prior to launch This means that a significant mission risk is introduced by the MOR architecture For this reason a direct return architecture has been adopted for the sample return mission Mars InSitu Pronellant Production Mission Andreas Mogensen 60 Mission Systems Description This section provides a detailed description of the three main ight system elements and their subsystems The functional requirements that each ight system must satisfy are discussed and an initial conceptual design of a system which meets these requirements is offered The conceptual design provides a rst estimate of the system mass and dimensions A more detailed design of each system will be carried out in the preliminary design stage once the mission has been accepted 6 1 Lander The lander must satisfy several key functional requirements First the lander must deliver a payload that supports sample collection operations and ISPP operations Second the lander must form a base of operations on Mars which can support at least 210 sols of surface activities Finally the lander must enable the launch of the ERV The lander is not expected to survive launch The resulting design derives from the lander design of the MER mission and the experiences gained during that mission The deck of the lander is a modi ed hexagonal shape that is slightly elongated to accommodate the length of the ERV The Width of the deck has been reduced to allow the lander to t Within the payload fairing of the Delta 2 launch vehicle Six petals enclose the deck of the lander during cruise forming a pyramidtype shape The front and rear rectangular petals form the main structure of the lander in its cruise con guration as shown in Figure 2 The four remaining petals are triangular in shape and form an enclosed structure for the protection of the payload 215m 27m 27 m Figure 2 The cruise con guration ofthe lander Upon landing the six petals are opened to reveal the deck of the lander The front and rear petals serve as or off access ramps for the rover The four remaining petals are Mm Tn im 39 Mi sinn covered in solar arrays providing power when the lander is deployed and the petals are opened All six petals are attached to the deck via a motorhinge arrangement that allows the lander to upright itself after touchdown and deploy the remaining petals Four separate jackup legs can be pushed into the ground to level the deck of the lander and provide additional stability a er the petals are deployed A schematic diagram of the deployed lander is shown in Figure 3 F Communlcatlons l l system Solar arrays Rear access 49 m ramp Solar arrays Front access ramp ISPP Plant 675 m Figure 3 The deployed con guration or the lander The ERV is stored horizontally within the deck of the lander and is encased in a thermal regulation unit which provides refrigeration of the propellant to minimize boiloff It should be noted that only the oxygen component of the propellant requires refrigeration since ethylene is storable under Mars ambient temperatures at a pressure ofjust a few bar A sliding rail arrangement runs along the length of the ERV This system will raise the V into a vertical position prior to launch The jackup legs will provide additional stability to the lander when the ERV is in the vertical position The rover is positioned over the ERV for the duration of the Mars cruise A clearance of about 5 cm between the ERV and the rover allows the rover to straddle the ERV A er touchdown and deployment of the petals the rover egresses from the lander via either the front or rear ramps When the rover has collected sufficient core samples it returns to the lander where it uses either the front or rear ramp to regress to its position above the ERV The core samples are then transferred directly from the rover to the sample return capsule in the EEV The mass of the lander is estimated to be 450 kg which is about 100 kg more than the lander of e MER mission The increased mass accounts for the larger structure of the lander and the additional payload systems such as the thermal regulation system of the ERV the railing system used to raise the ERV into a vertical position the ISPP plant the solar arrays and the communications systems Note that the structural mass of the ERV and the rover are not included in this estimate Mars TnSitu Pronellant Production Mi ion Andrea Mogen en 611 In Situ Propellant Production Plant A schematic diagram of the ISPP process is shown in Figure 4 The ISPP plant draws in carbon dioxide from the atmosphere and lters and puri es it The carbon dioxide is then reacted with the imported hydrogen feedstock in the reverse water gas shift RWGS reaction The reaction which is mildly endothermic will occur rapidly in the presence of a catalyst at temperatures of 400 0C or greater producing carbon monoxide hydrogen and water The reaction must be narrowly catalyzed to avoid the alternative exothermic reactions which will produce methane or methanol RWGS Reaction Ethylene Reaction Stored 2COZ 6H2 2C0 4H2 ZHZO 2C0 4H2 C2H4 2HZO C2H4 Electrolysis Stored 4HZO 4H2 202 gt 202 Figure 4 A schematic diagram of the ISPP process The equilibrium constant of the RWGS reaction is very low Hence in order to drive the reaction to completion the RWGS process is operated with an excess of hydrogen The water that is produced by the reaction is condensed and electrolyzed while the carbon monoxide and hydrogen mixture is fed into an ethylene reactor The ethylene reactor uses an iron FischerTropsch catalyst to produce ethylene and water The ethylene reaction is highly exothermic and thus can be used as a heat source to provide the energy needed to drive the endothermic RWGS reaction In addition the equilibrium constant is very high making high yields of ethylene possible However if the reaction is not narrowly catalyzed the reaction will have side reactions producing methanol CH3OH and propylene C3H5 The three hydrocarbons are all miscible and a mixture of the three would still constitute an acceptable propellant combination However the amount of imported hydrogen would need to be increased in this case The water that is produced in both the RWGS reaction and the ethylene reaction is electrolyzed to produce hydrogen and oxygen The hydrogen is recycled back to the RWGS reaction while the oxygen is liquefied and stored The specific impulse of the ethyleneoxygen bipropellant is 376 s Other important physical and chemical properties of ethylene oxygen and hydrogen are summarized in Table 5 The boiling point M1 1 u Andreas Mu mseh of ethylene is relatively elose to Mars ambient temperatures A a e 5 sh 39 which average about 55 c s F gur ows under a pressure ofa few bar ethylene 1s 510m eonMars 39 39 39 quot 39 L4 391 39 the need for refrigeration Only che oxygen pmpenam and the hydrogen feedstock will I n needre 39igeraho chemiml Cnmpnund Liquid Density Builinanint Mnlenllar Weight k m l rc mnl Ethylene 56792 71 28 I54 Oxygen 1141 7183 31999 Hydmgeh 7U 97 7253 2 mm Table 5 Chaniml and phys39tzl prnper zs nnhe prhpeuams VapuurF ressurE1Jaruv 1 MF E m w m we wn m we 2m 2n ymmmehmmmmmmm Tempernlme m rempsyahmrc Figures Vapm pressure nf ah Mars TnSitu Pronellant Production Mi ion Andrea Morgen en A total of 205 kg of ethyleneoxygen bipropellant will need to be produced on Mars to provide sufficient fuel for the MAV 185 kg is needed by the first stage to reach a 600 km parking orbit and 20 kg is needed by the second stage to perform the transEarth injection With a propellant production rate of 10 kgday it will take the ISPP plant 205 days to produce the required amount of fuel The mass and power requirements of an ISPP plant with this production rate can be estimated from Ref 11 and are shown in Table 6 Note that the power requirements correspond to 12 hour daytime power The ISPP plant including the reaction vessels the sorption pumps the controls the lines and valves and the refrigerator is located on one side of the deck of the lander ISPP Plant Element Mass kg Power W Sorption pumps 4 30 Chemical synthesis 5 300 Controls 3 40 Lines valves misc 4 0 Refrigerator 3 120 Total 19 490 Table 6 Mass and power requirements of the ISPP plant 612 Power System Power for the ISPP system the communications system and all the auxiliary systems is provided by solar arrays The most energy intensive of these systems by far is the ISPP system which requires 490 watts per day The solar arrays are sized out from this power requirement At the time of arrival at Mars the position of Mars in its orbit measured from the vernal equinox is 142 This corresponds to late summer and consequently the sunlight intensity should be good The equatorial landing site 5 0N 5 0S latitude at Terra Meridiani should also ensure that the sunlight intensity remains good throughout the mission time spent on the surface of Mars Consequently we can assume an average direct solar incidence of 500 Wmz If we further assume that the conversion efficiency of the solar arrays is 25 then the requirement becomes a combined area of 4 m of solar arrays Then the mass of the solar arrays can be estimated to be 12 kg based on an average mass distribution of 3 kgm The 4 m2 of solar arrays will be fitted to the four triangular side petals shown previously in Figure 3 Each side petal has an area of 1125 ml yielding a total combined area of 45 m2 Consequently it should be possible to generate the required power from solar arrays located on the four side petals of the lander Mars TnSitu Pronellant Production Mi ion Andreas Mogensen It should be noted that the ef ciency of the solar arrays will decrease with time as they become covered in dust At this point it is not known if the solar arrays will be capable of functioning for the duration of the surface operations which is nominally 210 days It may be necessary too include some mechanism by which the dust can be removed from the solar arrays If this is not possible and if the rate at which the ef ciency decreases is too high then it may be necessary to use a radioisotope thermoelectric generator RTG instead of the solar arrays 613 Communications System The lander carries three separate communications systems The rst system is a directto Earth Xband link which provides the main communications system between the surface elements and the command and control center on Earth The second system is an Sband radio system between the lander and the rover This system handles command and telemetry functions on the rover and is capable of a rate of 256 kbps in telemetry transmission from rover to lander and a rate of 8 kbps in command transmissions from lander to rover The third and nal communications system is a UHF antenna capable of providing a link with Mars orbiting spacecraft This link provides a backup link in case the Xband link fails 62 Rover The rover must satisfy the following functional requirements First the rover must collect 500 g of rock and soil samples from a range of geologically diverse sites Each sample must be documented so that the context of the sample is well understood Second the rover must perform a range of measurements to characterize the landing site area Third the rover must transfer the samples to the ERV Finally the rover must operate for a minimum of 90 sols traveling a total combined distance of at least 15 km The rover design is based on the Athena class rover which has previously been suggested for use in a Mars sample return mission Ref 8 The 70 kg rover is 131 m long 11 m wide and 15 m tall in its deployed con guration The rover carries the Athena science payload which is described in detail in the next section Power for the rover is supplied by a 12 m2 solar panel Surface operations will nominally proceed as follows The rover obtains a panoramic image and a spectral measurement of the prospective site which are sent to ground operations and analyzed An assessment of the images leads to a selection of potential targets and the rover is commanded to drive to one of the targets At the site the rover collects a core sample The core sample is then imaged and analyzed Its mineralogy is compared with other previously collected samples in order to asses the diversity of the current sample The process is repeated until suf cient samples have been collected to warrant a return trip to the lander At the lander the rover drives up the access ramps and positions itself over the ERV The samples are then transferred directly from the rover to the sample return capsule in the ERV When the transfer is complete the rover returns to the Martian surface to resume its sample collection mission 20 AndrensMn msen 4mm Figure 6 Athena class mer R1212 a 621 Science Instruments The rover earnes the Athena sclence payload wlnen eonslsts oftne mstruments llsted m able 7 lnslrnmen nurse ermrnnre emerr cnnretenze tn nrmnnnngs Mrmrtnre tnerrnn ernrsnrns cnnretenze tn nrmnnnngs sneetmrneter Mrmrtnre carer ootnntn care srrnnle Almrnmtrnrny speernneter cnnretenze tn rnrnernrgy awn srrnnle Mressorner spectrrrneter cnnretenze tn rnrnernrgy awn srrnnle ennrn spectrrrneter cnnretenze tn rnrnernrgy awn srrnnle Mnrnrnrger lrnrge tn srrnnle Tzh le 7 Allle science inslrumems 622 Cnmmnniean nnsSystem y telemetry funetaons on the rover and ls eapable of a rate of 256 kbps m telemetry txansmlsslon from royer to lander and a rate of 8 kbps m eommand transmrssrons from lander to royer Thls lmkl allnerofrslght system Henee m order to allow for overrther honzon operanons mdependent of the lander a seeondary U39HF eommunreatrons system 21 Mr In lhl Mt inn 4 provides a link for rover to orbiter relay communications This system is also capable of quot 39 quot r oorbiter 39 f 392 39 co and transmissions from orbiter to rover Ref 8 63 Earth Return Vehicle The functional requirements of the Earth Return Vehicle ERV are to return the samples to Earth and to keep the samples isolated and secure until the samples have been retrieved More speci cally the ERV must first launch the samples into a 600 km parking orbit with an inclination of 6 The requirement for the orbit inclination results from the declination of the launch azimuth which at the eparture ate is 6 Second when the launch window opens the ERV must inject the payload onto a transfer trajectory to Earth The ERV consists ofaMars Ascent Vehicle MAV and an Earth Entry Vehicle EEV The MAV is a twostage liquid fueled rocket and the EEV is the payload e EEV contains the Sample Retum Capsule SRC which isolates and protects the core samples re 7 A schematic diagram ofthe ERV is shown in Ergu Earth Entry Vehicle Stage 2 Len 35 cm Length 65 cm gm Radlus 30 cm Stage 1 Length 125 cm Radlus 60 cm Figure 7 Earth Remm Vehicle The ERV is stored horizontally on the deck of the lander for the duration of the surface operations When the ERV is fully fueled it is raised into avertical position and launched into a600 m parking or it by the first stage o e MAV e rrst stage is jettisoned on e s u as p navigation and control for the EEV during the Earth cruise phase Upon arrival at Earth the EEV is targeted for a ballistic reentry approach and released from the ERV The Mars Tn8itu Pronellant Production Mi ion Andrea Morgen en ERV then performs a de ection maneuver which causes it to miss the Earth It is also possible that the ERV which at this point only consists of the second stage of the MAV is sufficiently small that it can be allowed to reenter the Earth s atmosphere and burn up A more detailed assessment of the risks associated with this will have to be performed 631 Mars Ascent Vehicle The MAV is a twostage rocket where the first stage is capable of launching the ERV into a 600 km parking orbit and the second stage is capable of injecting the ERV onto an Earthbound transfer trajectory This translates into the following requirements The first stage must provide a AV for launch of 403 kms while the second stage must provide a AV for the transEarth injection burn of 207 kms in order to obtain the required value of C3 which for the July 2011 launch opportunity is 70 kmzs2 Ref 1 A first approximation of the mass and size of the MAV is obtained from an analysis of the variation of the mass fractions shown in Figure 8 Note that the payload of the MAV is the EEV which has an estimated mass of 8 kg Ref 11 An analysis of the data presented in Figure 8 reveals that reasonable estimates for the step structural factors for the first and second stages of the MAV are 02 and 048 respectively This leads to the mass break down shown in Table 8 for the entire ERV 7 Structural 7 Structure 7 Propellant 7 Stage 1 Mass kg Stage 1 Mass Fraction 01 02 03 0 01 02 03 Stage 1 Step Structural Factor Stage 1 Step Structural Factor 0 V Stage 2 Mass kg Stage 2 Mass Fraction 01 01 02 03 04 05 0 01 02 03 04 05 Stage 2 Step Structural Factor Stage 2 Step Structural Factor Figure 8 Variations in mass with step structural factor 23 Mars TnSitu Pronellant Production Mi ion Andrea Mogen en Mass kg Stage 2 Stage 1 Payload mass 8 46 Propellant mass 20 185 Structural mass 18 45 Gross initial mass 46 276 Table 8 Mass breakdown for the ERV Table 8 shows that the total dry mass of the ERV is 71 kg and that the gross initial mass of the ERV is 276 kg of which 205 kg are propellant An analysis of the chemical reactions involved in the ISPP process reveals the mass and volume of the ethyleneoxygen bipropellant for each stage This information is summarized in Table 9 The relevant values of the molecular weights and liquid densities of the ethyleneoxygen bipropellant that where used in the calculations where given previously in Table 5 MassVolume of EthyleneOxygen Stage 2 Stage 1 Mass of ethylene 610 5638 Mass of oxygen 1390 12862 Total Mass kg 20 185 Volume of ethylene 00107 00993 Volume of oxygen 00122 01127 Total Volume m3 00229 02120 Table 9 Mass and volume breakdown for the ethylene and oxygen propellant The size of the propellant tanks can be estimated from the total propellant volumes listed in Table 9 If the radius of the ERV is 03 m then the length of the propellant tanks for the first and second stages are 075 m and 01 m respectively This leads to the initial sizing estimate of the ERV shown in Figure 9 The total length of the ERV is 19 m and the maximum radius is 03 m A similar analysis can be performed on the hydrogen feedstock that is imported to Mars The analysis reveals that the total amount of hydrogen feedstock required to produce 205 kg of ethyleneoxygen bipropellant is approximately 9 kg By allowing for 20 boiloff during the cruise to Mars approximately 11 kg of hydrogen will have to be included at the time of launch from Earth This amount of hydrogen will occupy 01518 m3 which can easily be stored in the propellant tanks Note that the amount of hydrogen will slowly decrease as the amount of ethyleneoxygen bipropellant is produced on Mars 24 Mfr Tm rm Mi ion Element Length in Earth Entry Vehicle 0 35 m Auxlllary systems 0 20 m Propellant tanks 0 10m Engines 0 15 m Auxlllary systems 0 15 m Prop ellant tanks 0 75 m Engines 0 20 m Figure 9 Length at the elements at the ERV 632 Earth Entry Vehicle The EEV is a blunt shaped conical structure approximately 03 m in diameter with a low ballistic coelficient for low terminal velocity It consists of a secure and isolated l 39 at he SRC during the ballistic reentry hile the crushable honeycomb matrix absorbs the landing impact The EEV also includes abeacon which guides the search and retrieval party to its location The mass of the EEV is approximately 8 kg Ref 11 Mars and provides the structural strength and integrity to protect the samples The heat shield protects t w Summ shield Abrams l Figure 10 Earth Entry Vehicle Ra m Figure 11 Sample Lransfer Ret s1 Mars TnSitu Pronellant Production Mi ion Andrea Morgen en It should be noted that the design of the heat shield has not been determined at this point in the mission design and Figure 10 is only a possible solution It may be sufficient to simply coat the outside of the sphere with an ablative material rather than use a conical shaped shield After the rover has collected core samples on Mars they are transferred directly from the rover to the SRC The rover regresses to its xed position on the deck of the lander and transfers the samples through an opening in the MAV and into the SRC as illustrated in Figure 11 Ref 8 Once all the samples have been transferred the SRC is sealed by pyrotechnic welding for example The samples and the inside of the SRC which has been exposed to the Mars environment is now isolated The lid of the SRC which has also been exposed to the Mars environment is sterilized to break the chain of contact with the Mars environment and prevent the release of Martian contaminants into the Earth s atmosphere The rest of the SRC is sealed inside the MAV and never encounters the Martian atmosphere Ref 8 Upon arrival at Earth the EEV will be targeted for a ballistic reentry trajectory and released from the ERV The ballistic reentry trajectory avoids the large AV requirements and hence fuel requirements associated with an Earth orbit insertion The EEV will use a passive reentry system consisting of an ablative heat shield and a crushable material with high energy absorption characteristics This honeycombed structure will absorb the landing impact and prevent deforming loads from reaching the SRC A passive reentry system is favored over an active system such as parachutes because the parachute system does not have the incredibly high reliability necessary to meet the containment assurance requirements Consequently the EEV would still have to be designed to survive ground impact in the event of a parachute failure The ground impact in this case would be much more severe because the EEV would be signi cantly more massive in order to accommodate the parachute system Hence it is much more reasonable to use a passive reentry system It should also be noted that the scienti c value of the samples are not reduced by the impact loads associated with a passive reentry system Ref 8 Ground impact is favored over water impact even though water impact is more benign However inclement weather such as rough sea conditions and sinking introduces substantial risk in the recovery of the samples This not only represents mission failure but also a loss of sample containment Consequently the EEV is nominally targeted to land in the salt ats of Utah 26 Mars ln itu Pronellant Production Mi ion Andrea Morgen en 70 Mission Pro le This section will describe the mission pro le in detail The mass schedule of the mission will be presented rst followed by a discussion of the Marsbound and Earthbound trajectories Finally the timeline will be presented together with an indepth discussion of the sequence of events of the mission 71 Mass Schedule The total mass of all the surface elements to be landed on Mars is approximately 600 kg This includes the mass of the lander the rover the ERV and the hydrogen feedstock The 600 kg compares well with the total mass of the surface elements on the MER missions which was approximately 540 kg Although the MER missions consisted only of a lander and a rover the mass of the MER rover at 185 kg makes the total mass of the surface elements of the two missions comparable Consequently as a rst approximation it is assumed that the masses of the cruise systems and the entry landing and descent systems for this mission are of the same order as the systems on the MER mission The resulting mass schedule for the mission is shown in Table 10 Mission Element Launch Mass kg Lander 450 Structure and auxiliary systems 419 ISPP Plant 1 9 Solar arrays 12 Rover 70 ERV 71 MAV unfueled 63 EEV 8 Hydrogen feedstock l 1 Cruise stage 195 Back shell Parachute 210 Heat shield 80 Propellant 50 Total 1137 Table 10 Mission mass schedule 27 Mr In itu Ml mm 4 72 Trajectories The mission will utilize minimum energy trajectories for both the outbound and inbound transfer trajectones For the October 2009 Earth to Mars launch opportunity there are two minimum energy trajectories aType 2 and aType 4 trajectory The mission will use the Type 2 transfer trajectory shown in Figure 12 in order to avoid the excessive two year transfer time associated with the Type 4 trajectory The Type 2 trajectory has a transfer time of about 11 months and a c value oflll km s2 This is the maximum c3 value for the 20 day launch window It is wonh noting that the declination ofthe launch azimuth is less than 285 which means that the mission can be launched from Cape Canaveral Additional characteristics of this launch opportunity and transfer trajectory are summarized in Table 11 Ref 1 In Table 11 the quantity Ls refers to the position of 39n its orbit around the sun measured from v mox In addition the quantity DAP refers to the planetequatorial declination ofthe incoming asymptote that is ofthe V vector The DAP provides ameasure of the minimum possible latitude of impact for the Earth return Mars Arrival Augusl 2010 Mars Departure July 2011 Eath Reruru JV 1012 Eanh Lauuctr 39r OcluherlDUSr r in e transfer time is approximately 11 months and the maximum c3 value is 70 kasz The declination of the launch azimuth is 6 which constrains the inclination of the parking orbit and the latitude of the Mars landing site To minimize AV requirements and fue 28 Mr In mt Mt ton requirements the latitude ofthe landing site should be such that the ERV can be launched directly into a parking orbit with a 6 inclinations thereby avoiding expensive plane change maneuvers The DAP value at Earth arrival is suf ciently small that the EEV can be targeted to land anywhere in the continental USA Characteristics Outbound Inbuund Launch period 10042009 7 10232009 07292011 7 08182011 Trajectory type 2 2 c3 kHz52 11 1 7 0 Declination ofthe launch aztmuth lt 28 5 6 Arrival penod 082820107 09052010 07082012 vc at amval hms 2 5 3 5 15 at arrival 142 a Solar declination at amval 15 e Maxtmum latitude oflandlng site 50 N15 s 7 DAP 7 11 Table 11 Characteristics at the uutbuund and inbuund transfer trajecmries 73 Mission Phases and Timeline An overview ofthe mission timeline is given in Figure 13 below As the gure shows several distinct phases can be de ned for the mission Table 12 contains a summary of these phases and their dates assuming a launch at the rst opportunity of each launch window l l nts m on AM 01 1m Aer 1n Um 1m nor Jul 01 m nor Jul on m ltt ttlttlttltt ttlttlttltt ttlttlttltt ttlttlttl t 2 H t v EanhLaumh v cruise Sudcc unemttons ERV Launch V MarsDrhll mv summits sanhuentmv Figure 13 Missiun Timeline Mars Jn8itu Pronellant Production Mi ion Andrea Mogen en Mission Phase Dates Comments Launch 4 Oct 2009 Starts at liftoff and ends at time of spacecraft separation from launch vehicle Mars cruise 4 Oct 2009 to From launch vehicle separation to atmospheric 28 Aug 2010 entry at Mars radius of 35222 Entry descent landing 28 Aug 2009 From Mars entry to touchdown on surface Surface operations 28 Aug 2009 to From touchdown to launch of ERV mid Apr 201 1 Approximately 210 days Mars ascent mid April 2011 From surface launch to orbit insertion Launch date depends on surface conditions Mars orbit Mid Apr 2011 From insertion into parking orbit to transEarth 29 Jul 2011 injection burn Earth cruise 29 Jul 2011 to From transEarth injection burn to atmospheric 8 Jul 2012 entry at Earth Entry descent landing 8 Jul 2012 From Earth entry to touchdown Table 12 Mission phases 731 Launch The mission is nominally scheduled for launch from Cape Canaveral on 4 October 2009 The 20 day launch window spans from October 4 to October 23 and has a maximum C3 value of 111 kmzsz The total launch mass of the spacecraft is 1137 kg and will require the use of a Delta 7925H launch vehicle The heavy lift version of the Delta 7925 vehicle is necessary in order to achieve the maximum C3 value as shown in Figure 14 Ref 2 The maximum dimension of the spacecraft in the cruise con guration is 27 m Consequently either the 29 m or 30 m payload fairing can be used 732 Trans Mars Cruise The Mars cruise phase will last approximately 11 months with arrival at Mars at any time between 28 August 2010 and 5 September 2010 During the cruise the spacecraft will execute several cruise maneuvers to control its trajectory Maneuvers early in the phase will primarily correct for dispersions in the trajectory caused by launch injection errors while maneuvers late in the phase will ensure that the uncertainty in the ightpath angle at atmospheric entry is less than i 01quot Based on previous missions 60 ms of AV has been allocated for these maneuvers The cruise con guration of the spacecraft will be very similar to the cruise con guration of the MER mission which is shown in Figure 15 Ref 6 30 Mar In imPrnnpiiznthdnctinn Mi inn Payload kg Andrea Mn en en 29m 9541 Fairing Figure 14 Launch energy of the Delta 7925 and Delta 7925H vehicles Ref 2 l I r i rum3910quot rum 3 213 Mum c n u mu lam144mquot I u n IHr m nll nnwc Slug I sanmm m 4 mn Imm i M nus 32m n i u m quotmm111mm l a 2 rm Mum u pom 39uldc kcrmhcli aummm 4mm 39 nrlt Hm mummums Y imamu we ummnn m mm i39unA39i i nnlh Mm yaw Hmmmn Mum i o iulr 0 1M Figure 15 Cruise con guration of the MER mission spacecraft Ref 6 31 AndrasMu ensen 733 Mars Entry Landing and Descent The mlsslon wlll use an ungulded ballrstrc entry landrng and descent pro le that ls yery m ar e descent proflle of the Mars path nder mlsslon and the Mars Exploratron l R 6 per y s separatlon Just pnor to atrnosp enc entry A supersomc parachute ls then deployed to slow the spacecraft shortly alter thrs the heat shleld ls Jettlsoned and the lander ls separated from the backshell by lowenng rt on a tether or brldle The lander then beglns radar h 4 mm A rts horrzontal yelocrty Next the alrbags are rn ated and the retrorrockets are flred bnn In the lander arrzero vemcal yelocrty The bndle ls then cut and the lander a m Tumsxan L c l m rm Camale39mEyL 77 mm la 2 ulnsngesrmlen L lmln Awwsvhaminlry L 5mm undrele Vsmll rumple L 4m menl chem t us see sputum ls a my sawmm army ruesmlasemlen tam Houseman 1 n aasrawmlm see 2 m h smmma cm W ug cmllmi esAB alvm L Sass 2 an e A Lr ssc v m am 91mg L 225 V1kmamcmun V vslirAlhangalan L an an mm Ratva xkyang lesm mm lammurmur am at ace Mmbemaapmmmala l e l l are mm new Nahum V mama mansquot I armor FDHSUDKDV m in Rausm lam Jewrgscy J Lee 7 l gtci 7h N aags m loSEmn g k am ms m A I mLummlw V as gtm Figure 16 Entry descent and landing Sequnce men a a nalr oflowrgaln antennas These tones wlll slgnal the accornplrshrnent of crlucal tasks durlng the descent phase Dependrng on the orbrtal assets around Mars atthe trrne ofamyal the The landlng slte has nomlnally been selected as Tara Merrdlanl because of the A h However any equatonal landlng srte wth 5 quotN and 5 s ls posslble Mars TnSitu Pronellant Production Mi ion Andrea Mogen en 734 Surface Operations Surface operations will commence as soon as the lander rolls to a complete stop The rst step will be the retraction of the airbags and the deployment of the side petals The lander will determine its orientation and if necessary it will raise itself to an upright position by deploying the correct side petal With the deployment of the side petals the solar arrays will begin to recharge the batteries The panoramic camera mast on the rover will then be raised into its upright position and a rst panoramic image of the landing site area will be taken and transmitted to Earth During the following days communications with Earth will establish the health of the lander and its subsystems When all systems have been checked out and the lander is ready to commence surface operations the rover will be deployed via the front or rear access ramp At the same time the ISPP plant will begin propellant production The propellant production will proceed automatically with minimal supervision by the ground operations team However the rover will need signi cant supervision to guide it to appropriate sites where it will collect core samples It is envisioned that the rover will make three or four sorties of varying length to different areas around the landing site The rst sortie will likely be a very short drive and grab sortie to guard against possible rover contingencies The following sorties will be of longer length and duration During each sortie the rover will collect a range of samples which will then be returned to the ERV before the next sortie After approximately 90 sols the primary rover mission will be completed If possible an extended rover mission to sites further away from the lander will be performed 735 Mars Ascent Approximately 210 sols after touchdown the ISPP plant will have produced suf cient ethyleneoxygen bipropellant At this point there is approximately 3 months until the launch window opens at the end of July 2011 The ground operations team will decide on a suitable launch day and on the chosen day the ERV will be raised into a vertical position The rst stage of the MAV will launch the ERV into a 600 km parking orbit around Mars at which point the rst stage is jettisoned The ERV will remain in orbit until the launch window opens at which point the second stage will be used to initiate the transEarth injection burn The launch window for Earth return is from 29 July 2011 to 18 August 2011 736 Trans Earth Cruise The Earth cruise phase will last approximately 11 months with arrival at Earth on 8 July 2012 During the cruise the spacecraft will execute several cruise maneuvers to control its trajectory For planetary protection reasons the ERV will maintain a trajectory that misses the Earth Shortly before Earth encounter the ERV will be commanded to perform an entry targeting maneuver which will change its trajectory from a yby 33 Mars TnSitu Pronellant Production Mi ion Andrea Morgen en trajectory to a ballistic reentry trajectory Shortly after this the EEV will be released and the ERV which now consists of the second stage of the MAV will perform an Earth de ection maneuver of 15 ms Ref 5 to keep from following the EEV into the atmosphere During the Earth cruise phase auxiliary systems such as power communications and guidance navigation and control will be provided by the second stage of the MAV 737 Earth Entry The EEV will be spin ejected from the ERV l 7 4 days prior to atmospheric entry and will passively follow the reentry trajectory The ablative heat shield will protect the EEV during the reentry and the low ballistic coef cient will ensure that the terminal velocity is low and on the order of 20 ms Ref 11 At this velocity the impact loads will be suf ciently small to allow the shock absorbing properties of the crushable material to absorb the loads The landing site is currently chosen as the salt ats in Utah A beacon will guide the search and retrieval party to the location of the EEV The EEV will then be taken to a containment facility where the core samples will be retrieved from the SRC 34
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