Chapter 4 - Sample Return Mission

Chapter – Sample Return Mission Sample Return Mission 4.1 Mission Overview Fig 4.1 Created By Ben Toleman 4.1.1 Introduction – Matt Maier One of the most important scientific tasks we are conducting during this mission will be the return of a Martian soil and rock sample This mission lasts for the duration of the astronauts stay in Martian orbit The astronauts have the task of using the two different rovers to collect the sample and perform other scientific measurements Note that the rovers are located on opposite sides of the planet for maximum communication time This mission will play a significant role AAE 450 Senior Spacecraft Design Spring 2004 167 Chapter – Sample Return Mission for future manned missions to the surface of mars We would like for a human mission to use a maximum amount of resources found on Mars, this would reduce the mass and cost of putting the first human on Mars In conjunction with the data collected from the Mars Exploration and Pathfinder missions the soil and rock data from our robotic missions will help choose an appropriate landing site for such a mission Another benefit we gain from the sample return mission is the demonstration of producing the required propellant for a Mars to orbit launch This is a very important technology that must be proven before a human landing is possible Other technological benefits such as precision landing will are also demonstrated in our rover missions The rock and soil samples once returned to Earth will provide researchers with data that would have taken numerous Mars rover mission to accomplish 4.1.2 Mission Timeline – Matt Maier The two rover landers are launched shortly after the aero-capture maneuver for the spacecraft has been completed Two landers are sent to the surface to ensure the success of the sample return mission in the event that one fails These failures include but are not limited to unsuccessful landing, improper rover or sample return vehicle (SRV) deployment, complications in propellant production or unfavorable weather conditions We target the landers at two different landing sites on different sides of the planet We need two landing locations for two different reasons; variety of samples and communication In the event that both sample return missions are successful it is beneficial to future missions to have very in-depth analysis of two different landing sites Placing the rovers on opposite sides of the planet allows for the design of our spacecraft’s orbit to ensure that the astronauts are always in contact with at least one landing site After the landers touchdown and deploy the rovers a subsystem of the lander starts producing the propellant for the sample return vehicle using in-situ production processes (section 4.6.4) It is necessary that the SRV should employ this technology not only for the reduction in mass but also to prove these techniques for future manned missions During this time the astronaut controlled rovers collect up to ten kilograms of samples and perform other important scientific duties Once the SRV has been fueled it is launched to rendezvous with the spacecraft orbiting The rest of this chapter discusses the details of these components and procedures AAE 450 Senior Spacecraft Design Spring 2004 168 Chapter – Sample Return Mission 4.2 Launch of Rovers 4.2.1 Release of Landers – Allison Bahnsen After the Transport Vehicle performs aerocapture and the periapsis-raise maneuver, and prior to the apo-twist maneuver, we release the two landers that venture to the surface of Mars The side, cross-sectional profile of the landers in the Transport Vehicle is shown on the left of Error: Reference source not found As we can see, the landers are housed within the body of the main spacecraft The image on the right shows a top view of the Transport Vehicle The protective, hexagonal doors covering the two landers are indicated by arrows Prior to release, these doors slide open to reveal the landers Protective Doors Fig 4.2 Side View of Landers in Transport and Protective Doors Created by Ben Toleman and David Goedtel We release the landers when the Transport Vehicle is traveling as slowly as possible to reduce propellant costs The slowest point in the Transport Vehicle’s orbit, seen in blue in Fig 2, occurs at apoapsis, where we release the first lander This release at apoapsis costs 1.05 m/s, AAE 450 Senior Spacecraft Design Spring 2004 169 Chapter – Sample Return Mission and places the first lander on the green trajectory in Fig with periapsis altitude at 100 km Since the second landing site is on the opposite side of the planet we wait half a sol (half a Martian day) to release the second lander Now that the spacecraft is no longer at apoapsis, we must find the orbit that intersects the current location of the Transport Vehicle and has a periapsis altitude of 100 km We can see this trajectory in red in Fig 2, and can transfer to it for a cost of 1.17 m/s These results are obtained using the MATLAB code in Appendix G Fig 4.2 Lander Trajectory For the above calculations we assume that the Transport Vehicle is in the same plane as the landing sites: the equatorial plane In actuality the Transport Vehicle is in the ecliptic plane when we releases the landers, and thus we would have to wait until the entry point in the atmosphere above the landing site is a node between the ecliptic and equatorial planes This plane change would cause the release ∆v’s (or change in velocity) to be three-dimensional, but their magnitudes would not be much larger than those of the aforementioned values Solving for these three-dimensional ∆v’s and the implementation times of hitting the two selected landing sites are not trivial matters by any means, and thus are out of the scope of this study This in no way affects the feasibility of the mission, it simply adds to the complexity of time lining the overall mission when it comes to fruition 4.2.2 Cruise Stage – Andy Kacmar The “cruise stage” is the configuration of the aeroshell for transport between the spacecraft and Mars Fig shows the full cruise configuration The cruise stage resembles AAE 450 Senior Spacecraft Design Spring 2004 170 Chapter – Sample Return Mission the Mars Pathfinder and Exploration mission designs The major differences that arise come from the fact that these missions were designed for the system to travel from Earth to Mars where our stage is only going to transport the aeroshell to the Martian atmosphere and ensure the proper entry point The structure affixed to the aeroshell is approximately 5.0 m in diameter and m thick With a mass of 235 kg, the structure consists of a basic aluminum frame with an inner and outer ring for support The top surface of the stage is lined with solar panels to supply power once detached from the spacecraft and the outer ring is lined with radiators to dissipate any heat build up from the Fig 4.3 Full Cruise Configuration Created by Ben Toleman solar radiation and electronics on board For navigation, there are three sun sensors (for redundancy), one star scanner, and an onboard positioning system coupled with the antenna to relay position and information back to the HAB For correctional maneuvers, the maximum Δv the system needs is less than 2.0 m/s This amounts to about kg of fuel when accounting for departure from the spacecraft, the correctional maneuver, and excess propellant left in the two aluminum lined tanks The cruise stage consists of two thruster clusters of four thrusters each running off of hydrazine propellant running through a catalyst bed The clusters allow for corrections in any direction to ensure a safe insertion into the Martian atmosphere Fig 4.4 Two-D Drawing of Cruise Stage – Created by Rebecca Karnes AAE 450 Senior Spacecraft Design Spring 2004 171 Chapter – Sample Return Mission 4.3 Atmospheric Entry / Touchdown 4.3.1 Landing Sites – Allison Bahnsen Nomenclature MER MGS TES = Mars Exploration Rover = Mars Global Surveyor = Thermal Emission Spectrometer One of the main scientific objectives of this mission is to return a Martian rock sample back to Earth for analysis hopefully leading to many new discoveries, including if life once inhabited Mars A major indication that water, the building-block of life, once existed on this hostile planet is the presence of an iron oxide mineral called hematite On Earth this mineral is usually formed in a large body of water in which iron is dissolved and gradually oxidized into hematite This insoluble mineral is then precipitated out and mixes in with the lake bottom sediment which eventually hardens into rock The hematite deposits on Earth are also one of the best rocks to serve as home to microscopic fossils of microbes that were trapped in the sediment before it hardened into rock.1 The presence of crystalline gray hematite on Mars was first observed by scientists analyzing the Thermal Emission Spectrometer (TES) data obtained from early phases of the Mars Global Surveyor (MGS) mission.2 Knowing that finding hematite could be the next step to discovering if life once existed on Mars, the presence of this mineral in the landing sites is a necessity The first landing site we select is located in the Terra Meridiani region of Mars, with the exact coordinates of 1.98° S, 6.18° W and a landing ellipse with dimensions of 81.5 km by 11.5 km Error: Reference source not found shows a photo mosaic of this region from Viking which is superimposed with data from the MGS TES We can see that the exact site, marked with an arrow, is located in an area with approximately 15% hematite This landing site is also the location of the Mars Exploration Rover (MER) Opportunity, which landed there on January 26, 2004 Already a few months into the mission, this site has proven to be a jackpot in the eyes of scientists containing the largest concentration of hematite that they have ever seen.4 AAE 450 Senior Spacecraft Design Spring 2004 172 Chapter – Sample Return Mission Fig 4.6 Hematite Distribution Map3 Aside from having a large distribution of hematite, this site also boasts low wind shear, a low abundance of boulders and low slope angles in the craters, all of which are positive attributes when looking to land and operate a rover The low wind shear in combination with the relatively low amounts of dust compared to other parts of the planet make this site not only a very good scientific candidate, but also very environmentally appealing We select the second landing site on the opposite side of the planet in the Athabasca Valles at 8.92° N, 205.21° W One of the main reasons to choose the second site to be on the opposite side of Mars is for communication issues This guarantees that one rover will always be on the side of the planet that is facing the Transport Vehicle, which gives the astronauts the maximum time to control the rovers This site was also one of the back-up sites for the MER mission.Error: Reference source not found We can see the site along with its landing ellipse, with dimensions of 152 km x 16 km, in Error: Reference source not found.6 AAE 450 Senior Spacecraft Design Spring 2004 173 Chapter – Sample Return Mission Fig 4.7 Map of Second Landing Site6 2nd Landing Site In addition to the presence of hematite, this site is appealing because as we can see in the elevation map in Error: Reference source not found, the site is in a large channel system that could have possibly been cut out by catastrophic floods or some other type of flowing water This location is also the seed of a great debate between geologists concerning the age Some think it is a geologically young site, while others think it is an ancient site that has just recently been exhumed.7 Therefore, obtaining a rock sample from this site could settle the dispute We can see both of the landing sites on a map of Mars in Fig 58 During the design process, concerns were expressed with regards to communication and the difference in inclination between the equatorial landing sites and the 63.4° inclined Transport Vehicle orbit These concerns have been addressed and dispelled in full in Appendix G AAE 450 Senior Spacecraft Design Spring 2004 174 Chapter – Sample Return Mission Fig 4.5 Landing SitesError: Reference source not found 4.3.2 Entry Trajectory 4.3.2.1 Mission Timeline – Ayu Abdullah We present our mission timeline for the Aeroshell containing the Mars Lander and Rover in Table below This timeline begins at first point of entry into the atmosphere, taken to begin at 100 km altitude A graphic timeline is also provided in Error: Reference source not found AAE 450 Senior Spacecraft Design Spring 2004 175 Chapter – Sample Return Mission Table 4.2 Mission Timeline Time (sec) Altitude (km) 0.0 100.0 261.9 262.2 267.2 272.2 272.8 9.0 8.9 7.9 6.9 6.8 368.9 1.7 408.9 0.0 Event Aeroshell with rover enters the atmosphere of Mars at 4.896 km/s and begins the landing sequence of events Entry, descent and landing (EDL) takes approximately 6.8 minutes Drogue deploys (304m/s) Drogue fills Aeroshell bolts are fired (200m/s) Heat shield separates Parachute attached to lander deploys, releasing it from backshell Parachute fills Lander altimeter returns information on altitude, rocket-assisted deceleration engines (retro-rockets) fire (85m/s) Bridle cable is cut Rover lands softly on surface of Mars Fig 4.9 Mission timeline 4.3.2.2 Aerocapture – Ryan Whitley The equations of motion delineated in Section 3.5.1 also apply to a more general reentry Thus, we propagate an entry trajectory for the Lander using these same equations The Lander’s desirable trajectory ends at parachute deploy altitude, and will hit the ground if left unchecked We choose a parachute deploy altitude of km Thus, an optimized trajectory contains an initial AAE 450 Senior Spacecraft Design Spring 2004 176 Chapter – Sample Return Mission Fig 4.24 How mass varies as the pressure increases, when R = 0.25 m 4.6.3 Power - Ben Phillips The sample return vehicle (SRV), which will carry the Martian rock sample back to the astronauts in the habitat module, requires power for approximately seven days This power is needed for the vehicle’s navigation while performing rendezvous maneuvers and eventual docking The relatively long mission lifetime of this return vehicle limits the number of options available for power There are only two viable options, either a solar array and battery combination could be used or a radio-isotope (RTG) power system The difficulty that arises with the solar array/battery combination is determining where the solar array would be placed on the return vehicle The use of a RTG power system is a good choice in this aspect because the entire system can be placed inside the return vehicle The effectiveness of a solar power array can be seen in the appendix The power that a solar array decreases at a rate of one over the distance to the sun squared This means that at Mars a solar array can only produce about half as much power as it could at Earth The mass for a RTG power system needed for the SRV is relatively small The radioisotope system would have a mass of approximately 15 kg 49 This value is miniscule when compared to the approximate 170 kg mass that a battery power system would require to provide AAE 450 Senior Spacecraft Design Spring 2004 217 Chapter – Sample Return Mission power to the SRV for its seven-day lifetime A fuel cell that could provide the power would have a mass of about 500 kg These mass comparisons show the savings that using a RTG power system gives in this case A radio-isotope system would also require less volume, in this case a cylinder with a diameter of 0.3 meters and a length of 0.4 meters Once the SRV has docked with the manned habitat module, the RTG power system will not present a radiation problem The chamber that in which the radioactive material is kept protects the astronauts from radiation, as does the EVA spacesuit that they would wear Once the astronaut has retrieved the Martian rock sample, the SRV will be jettisoned along with the radioisotope power system In conclusion, the radio-isotope power system for the SRV offers several advantages two of which are savings in volume and mass The increase in reliability that a radio-isotope system gives is also a major advantage These savings help keep the SRV payload mass to a minimum and thus keeping the sample return rocket to a minimum mass The difficulties in placing solar arrays on the SRV are also avoided 4.6.4 Propellant Production - Matt Maier Nomenclature CH4 H2O O2 H2 CO CO2 = = = = = = chemical formula for methane chemical formula for water chemical formula for oxygen chemical formula for hydrogen chemical formula for carbon monoxide chemical formula for carbon dioxide For this mission it is necessary to employ the techniques of in-situ propellant production This is an important process that must be proven before landing a man on Mars In this process we produce oxygen and methane from a supply of hydrogen and the Martian atmosphere (95% CO2) One of the benefits we gain from this process is reducing the mass needed to be taken from Earth For this mission it might seem like a small mass savings, but it is a technology that we need to be prove for human missions where the propellant produced would be used not only to bring the astronauts back to Earth but to provide ground based power and other various resources.50,51 AAE 450 Senior Spacecraft Design Spring 2004 218 Chapter – Sample Return Mission This process requires a supply of H and since it is not readily available on Mars we must import it from Earth The first step in producing our fuel is a process known as the Sabatier reaction (Equation –7) H + CO2 → CH + H 2O 4–7 For this analysis we apply the water-gas shift reaction (Eq 4-8) in conjunction with the Sabtier reaction to produce (Eq 4-9) The reaction converts carbon to methane and water by reacting it with the imported hydrogen; there is also an excess amount of carbon monoxide produced which we release into the Martian atmosphere This equation is exothermic therefore in the presence of a catalyst the reaction requires no net input of power to operate; therefore this is a favorable method.Error: Reference source not found,Error: Reference source not found,52,53 CO2 + H → CO + H 2O 4–8 3CO2 + H → CH + H 2O + 2CO 4–9 From here the CH4 is stored cryogenically and the water is then reacted using electrolysis (Eq 4-10) The O2 is then cryogenically stored and the H is reacted again as in Equation – We repeat this process until the H2 supply is exhausted producing CH This is a very economical process that ideally consumes all of the H in producing CH4 and also produces O2 This system has been proven to be very efficient, some designs operated at 99% efficiency.Error: Reference source not found,Error: Reference source not found,Error: Reference source not found,Error: Reference source not found H 2O → H + O2 AAE 450 Senior Spacecraft Design Spring 2004 219 4–10 Chapter – Sample Return Mission With this combination of chemical processes we produce kilograms of CH and 16 kilograms of O2 for every kilogram of H This is a 20:1 mass savings This produces a mixture ratio (Φ) of Recalling from section 4.6.1 a Φ of 2.99 is required for our rocket engines; therefore this is an acceptable method of acquiring the needed propellant.Error: Reference source not found,Error: Reference source not found,Error: Reference source not found,Error: Reference source not found We also note that the components applied for this process are quite simple (Fig 25) This is a very attractive quality of the production process The Sabatier reactor is basically steel pipes containing a catalyst Fig 4.25 Prolellant Production Unit bed, exercised to jump start the By Toleman reaction, and a required filter, to keep Martian dust out of our propellant The reaction occurs spontaneously if the catalyst is nickel or ruthenium (noble metals) Other necessary components include a cryogenic cooling system This is the main source of the energy requirements for the system The electrolysis reaction is the other process that requires a significant amount of energy A summary of the systems components is seen in Table 19.Error: Reference source not found,Error: Reference source not found,Error: Reference source not found,Error: Reference source not found Table 4.19 Propellant Production Summary Methane Mass 185 [kg] Needed Production 616 [kg/day] Rate Time 300 [days] Oxygen Mass 550 [kg] Needed Production 2.46 [kg/day] Rate Time 223 [days] Component Required 47 [kg] Hydrogen Production 20 [kg] Equipment Power 400 [kw] Required AAE 450 Senior Spacecraft Design Spring 2004 220 Chapter – Sample Return Mission 4.6.5 Optimizing the Launch of the SRV – Allison Bahnsen Nomenclature EOM FBD γ SRV TPBVP = = = = = Equations of Motion Free body diagram Flight Path Angle Sample Return Vehicle Two-Point Boundary Value Problem In order to simulate the launch of the SRV we first set up the basic FBD, which we can see in Error: Reference source not found From this FBD we can obtain the EOMs Fig 4.34 FBD of SRV by breaking the acceleration of the rocket into x and y components, where the flight path angle is denoted as γ We can see these components in the first four equations of Eq – We know that we want the rocket to start from zero altitude and velocity and hit a certain speed at a certain altitude This speed and altitude correspond to periapsis of the Hohmann-like transfer that travels out to the apoapsis of the Transport Vehicle orbit We can see this illustrated in Error: Reference source not found with the black curve representing the launch, the red ellipse AAE 450 Senior Spacecraft Design Spring 2004 221 Chapter – Sample Return Mission being the Hohmann-like transfer, and the blue ellipse being the orbit of the Transport Vehicle Fig 4.35 Launch into Hohmann-like Transfer Since we have EOMs and boundary conditions, this problem lends itself nicely to functional optimization and solving a TPBVP In this problem we want to minimize the launch time to orbit, which in turn minimizes propellant We set up the TPBVP and solve it via a MATLAB code written by Professor Marc Williams through following a tutorial written by Belinda Marchand.54 Marchand also wrote a second tutorial55 that details how to set up an optimization of a launch off of the moon, and we will follow her example The full details of this analysis can be found in Appendix G Below is our well-defined TPBVP Eq 4-11 shows the differential equations, where x and y , are the traditional EOMs obtained from breaking the acceleration into components and γ and β are the co-states obtained from the Euler Lagrange Equations AAE 450 Senior Spacecraft Design Spring 2004 222 Chapter – Sample Return Mission x = v x y = v y 4–11 T cos(γ ) m T v y = sin(γ ) − g m  β = β sin(γ ) γ = − β cos(γ ) v x = Error: Reference source not found shows the boundary conditions: Table 4.21 Boundary Conditions Initial ConditionsFinal Conditionstoyf = rc = 100 kmxovxf = vc = 4.91 km/syovyf = 0vxovyo Where rc is the desired altitude and vc is the desired speed at that altitude on the Hohmann Another parameter necessary for Professor Williams’ code is the specific thrust of the rocket, which is set at 4.54 as provided by Matt Maier in Section 4.6.1 The above well-defined TPBVP once inputted into Professor Williams’s code gives the optimal trajectory, which we see in Error: Reference source not found Error: Reference source not found shows the optimal steering law, which tells us that after launching vertically for a few seconds from the lander to avoid impacting any surroundings, the guidance rotates the rocket down to about 30° and will continue to angle the thrust downward until γ actually becomes negative While this seems counter-intuitive, as we can see in Error: Reference source not found the altitude continues to increase This decrease in γ is used to push the velocity in the y direction to zero, which is one of our final boundary conditions and a requirement to be at periapsis in a Hohmann transfer AAE 450 Senior Spacecraft Design Spring 2004 223 Chapter – Sample Return Mission Fig 4.36 Trajectory of Optimized SRV Launch Fig 4.37 Optimal Steering Law Error: Reference source not found highlights some of the optimized SRV parameters and the launch data AAE 450 Senior Spacecraft Design Spring 2004 224 Chapter – Sample Return Mission Table 4.22 Optimized Rocket Parameters ParameterNumeric ValueAltitude [km]100Range [km]732X-Velocity [km/s]4.91Hohmann speed at 100 km [km/s]4.91Burn Time [s]307Thrust [N]13,000 4.6.6 Docking of SRV and Retrieval of Mars Sample – Allison Bahnsen Nomenclature DART EVA ISS RCS RMS SRV = = = = = = Demonstration of Autonomous Rendezvous Technologies Extra Vehicular Activities International Space Station Reaction Control System Remote Manipulator System Sample Return Vehicle AAE 450 Senior Spacecraft Design Spring 2004 225 Chapter – Sample Return Mission Once the SRV launches off the surface of Mars following the optimized steering law detailed in Section 4.6.5, it begins its seven-day journey back to the Transport Vehicle When it closes within a few hundred kilometers of the Transport Vehicle, it is well on course due to continual course monitoring by the onboard guidance system and slight correctional inputs from the RCS jets It is at this time that the computer switches on the automated rendezvous software using technology obtained from DART.56 The DART technology includes collision avoidance software, and the system uses radar to determine the closing distances and relative speeds of the two spacecraft, similar to the proven Russian Kurs system used on the ISS 57 This software commands the RCS jets to fire until the relative velocity between the two spacecraft is negligible The software then switches over to the autonomous docking sequence which first ensures that the SRV is lined up in the general area of the docking receptacles The petals revealing the docking probe, seen Error: Reference source not found, have been opened and jettisoned along with the spent fuel tanks earlier in the mission Finally, the docking sequence uses the RCS jets to slowly insert the docking probe on the SRV into the cylindrical docking receptacles on the Transport Vehicle, which we can see in Error: Reference source not found The docking receptacles consist of three overlapping steel cables, each with one end attached to a fixed outer collar, and the other end attached to the movable inner collar, as we see in Error: Reference source not found Fig 4.38 Petals Opening and SRV Docking Mechanism – created by Ben Toleman and Matt Maier AAE 450 Senior Spacecraft Design Spring 2004 226 Chapter – Sample Return Mission Fig 4.39 Airlock (left) and Docking Receptacle with SRV mated (right)– created by David Goedtel and Ben Toleman AAE 450 Senior Spacecraft Design Spring 2004 227 Chapter – Sample Return Mission The lengths of the cables are the diameter of the cylinder, as we see on the left in Error: Fig 4.40 Reference source not found Prior Docking Receptacle, Collars to entry of the probe, the inner Inner Collar collar is rotated 60° causing the Outer Collar cables to go slack and allowing for the probe to enter We can see this configuration on the right of Error: Reference source not found Once the probe enters, the tip hits a push-button activator located in the back of the receptacle This activator releases a torsion spring between the two collars that then spins the inner collar back to its original position, thus securing the SRV to the side of the Transport Vehicle The concept of using cables attached to fixed and moving collars to secure payloads has been proven; it is used in the end effector of the Canadian RMS arm on the Space Shuttle to securely grapple and transport large pieces of hardware.58 Fig 4.41 Docking Receptacle, Cables AAE 450 Senior Spacecraft Design Spring 2004 228 Chapter – Sample Return Mission After confirmation that the two SRV’s have successfully attached to the side of the Transport Vehicle, the astronauts ready themselves for the pre-scheduled EVA Depending on when the sample retrieval is placed in the EVA timeline, the astronauts make their way over to where the SRV’s are secured, which we can see in the overall view of the Transport Vehicle in Error: Reference source not found Opening the same hatch that the rover used to place the rock sample cartridges in the rocket, the astronauts carefully remove the cartridges and place them in carrying bags The SRV’s, having completed their mission, are left attached to the side of the Transport Vehicle Fig 4.42 Airlock and Docking Receptacle – created by David Goedtel Airlock Docking Receptacle AAE 450 Senior Spacecraft Design Spring 2004 229 Moomaw, Bruce, “Uncovering The Meridiani Formation,” Space Daily, 04/02/01, http://www.spacedaily.com/news/mars2003-01a3.html Martel, Linda, “Grey Iron Oxide 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Company, Hampton, Virginia 15 Barua, D., AAE 450, School of Aeronautics and Astronautics, Purdue University 16 Whitley, R and Manning, R., AAE 450, School of Aeronautics and Astronautics, Purdue University 17 Whitley, R., AAE 450, School of Aeronautics and Astronautics, Purdue University 18 Soddit Matlab code written by Damon Landau and modified by Matthew Branson 19 Sandia One-Dimensional Direct and Inverse Thermal Code (SODDIT), Sandia National Laboratories, Albuquerque, New Mexico, 1990 20 Professor Steven Schnider, Associate Professor Purdue University 21 http://encyclopedia.thefreedictionary.com/Heat%20shield 22 Charles D Brown, Elements of Spacecraft Design, AIAA Education Series, Castle Rock, CO, 2002 23 Humble, Ronald, W and Henry, G N., and Larson, W J., Space Propulsion Analysis and Design, McGraw-Hill, 1995, Chap 24 http://www.goodfellow.com/csp/active/static/A/C_41.HTML 25 Hexcel Composites 26 Larson, Wiley J., and Wertz, James R., Space Mission Analysis and Design, Microcosm Press, Torrence, California, pg 407-427 27 Larson, Wiley J., and Wertz, James R., Human Spaceflight Mission Analysis and Design, pg 643-665 28 Lee, Darlene S., “Design and Verification of the MER Primary Payload Mars Exploration Rover Primary Payload Design and Verification”, Spacecraft & Launch Vehicle Dynamics Environment Workshop Program, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 06/17/03 http://www.aero.org/conferences/sc-lv/pdfs/lee_mer_03.pdf 29 Rummel , John D., Race, Margaret S DeVincenzi, Donald L., Schad, P Jackson., Stabekis, Pericles D., Viso, Michel., and Acevedo, Sara E., NASA, Hanover, MD, October 2002, NASA/CP-2002-211842 30 Mahaffy, Paul R and 15 co-authors (2003), The Organic Contamination Science Steering Group, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 12/02/03 31 “Tyvek®”, DuPont, Wilmington, DE http://www.tyvek.com 32 “Planova ® filters are designed for virus removal”, 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Flat-Moon Problem,” Purdue University AAE 508: Optimization in Aerospace Engineering Course notes, 03/04/98 56 NASA Facts, “DART Demonstrator to Test Future Autonomous Rendezvous Technologies in Orbit,” Marshall Space Flight Center, 09/03, http://www1.msfc.nasa.gov/NEWSROOM/background/facts/dart.pdf 57 Golightly, Glen, “Docking Zvezda: Tricky Space Ballet Takes Practice,” Space.com, 07/12/00, http://www.space.com/news/spacestation/zvezda_docking_000712.html 58 Thomas, Linda, “EVA Contingency Operations Training Workbook: CONT OPS 2102,” NASA Johnson Space Center, 03/95, pp 4-22 to 4-28 ... of mass AAE 45 0 Senior Spacecraft Design Spring 20 04 191 Chapter – Sample Return Mission 4. 4 Lander 4. 4.1 Introduction - Dan Nakaima As part of the mission, we are to obtain and return up to... Drogue Lander 170 385 10 .4 16.7 48 48 16 23 5 48 48 Volume [m3] Total mass [kg] 021 17 039 32 AAE 45 0 Senior Spacecraft Design Spring 20 04 186 Chapter – Sample Return Mission 4. 3.1 Retro Rockets Nomenclature... Design Spring 20 04 201 Chapter – Sample Return Mission 4. 5 Rover Fig 4. 18 CAD Images of our Rover - By Ben Toleman 4. 5.1 Mission Design - Masaaki Atsuta The requirement of our mission is to collect