2005 RASCAL POSTERS

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Lunar South Pole Sample Return Mission Presenters: Matt Bartone, Paul Carter, Steve Isley, John Moore, Ted Shreve Other Contributors: C. Adams, N. Anding, J. Buller, J. Coltrain, S. Conners, V. Dang, A. Harper, N. Holtz, R. Hosn, K. Ito, B. Karlson, M. Leibrant , C. Lloyd, A. Oakes, M. Rodriguez, A. Senta, B. Shelley, N. Trinh, T. Yamasaki, S. Yuen Faculty Advisor: A. T. Mattick, Associate Professor Abstract As part of NASA’s response to President Bush’s “A Renewed Spirit of Discovery,” the University of Washington Undergraduate Space Design Team has been designing a mission to evaluate lunar resources and implement emerging technologies of exploration as part of returning humans to the moon. The primary goal of the Lunar South Pole Sample Return Mission is to collect subsurface regolith in permanently dark areas near the Lunar South Pole to verify previous missions findings and to provide insight into the concentration and form of the ice in the regolith. The purpose of this project is to design a spacecraft to travel to the moon, collect core samples from regolith in the polar region, and return the samples to Earth in 2010 for analysis. Mission Criteria To improve the chances of finding water, the samples were to be taken from areas of permanent darkness. The core samples was to be stored in a hermetically sealed container to insure that the core samples were not contaminated. The samples were to be kept below 200 K, such that the water does not sublime and lose its original structure. The Lunar Prospector mission detected water down to 1.5 meters in depth; thus the core samples were to be collected from the surface to 2 meters in depth at three different sites. Each collection site was to be at least 50 meters away from one another to introduce variation. Mission Concept To obtain cores samples, two unmanned mission concepts were considered: a land roving vehicle and a “hopping” vehicle. The major influencing factor was the permanent darkness and how that effected the power source for mobility in each of the concepts. Without sunlight, the rover would be massive and complex due to its large power supply. The hopper would use the same structure that lands for surface maneuvering; the additional fuel required is only ~1% of the fuel required to descend. The difference of mass between these two concepts was on the order of ~1000 kg. Most controls issues can be simplified and this mission can be used as a “test bed” for future missions involving flight maneuvering. New Frontiers: Moonrise A proposal to the New Frontiers Program in 2003, the “Moonrise: Lunar South Pole-Aitken Basin Sample Return Mission,” also seeks to return lunar samples utilizing robotic means. In contrast, Moonrise seeks mantel samples left over from the deep impact crater. The “Moonrise” mission also seeks to use two rovers, situated at separate locations, to collect samples. These rovers would also be operating in sunlight, enabling Moonrise to use solar power. Mission Progression: Timeline The line along the bottom of each poster shows the progression of this mission highlighting each stage. The first three stages involve orbital maneuvering to position the spacecraft for lunar descent Stage I: Launch [0 hrs] Delta IV Medium + is used to launch the spacecraft directly into Trans Lunar Cruise Stage II: Trans Lunar Cruise [0.05 hrs] Three course corrections are initiated to keep course to the moon. The spacecraft is exposed to solar radiation, but requires only MLI– the only “radiators” needed are elements of the structure. Stage IV: Lunar Orbit [67 hrs] Upon arrival, the first burn occurs to inject the spacecraft into an elliptical orbit around the moon. A communications satellite detaches from the landing section of the craft (the Lander) while the Lander prepares for descent, so that data and commands can be relayed from the Lander to Earth Above: The surface maneuvering concept, or “Hopper”. The communications satellite, used to relay data and commands from the Lander to Earth, is not shown. Above, Top: The spacecraft for the rover concept. Above, Bottom: The rover for the rover concept. Below: Table of initial masses and the time at the end of each stage. Each of the stages are described in the timeline at the bottom, as well.

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Stage IV: Lunar Descent [68 hrs] Gravity turn trajectory is used to begin descent; at 500 m, a hazard avoidance algorithm coupled with LIDAR is used to safely land the Lander. Stage V: Lunar Operations [70 hrs] First set of core samples are taken at landing site. Stage VII: Ascent [280 hrs] With all of the samples stowed into the loading mechanism, the ascender detaches from the Lander to begin direct ascent to Earth. Stage VI: Hopping #2 [190 hrs] Second surface maneuver to travel to final drilling site. Third set of core samples are drilled and stowed. The Spacecraft The spacecraft, before launch include several stages and a communications satellite (Comsat). Since the target landing site is in permanent darkness, the Landing section of the craft will not be in direct line of sight of Earth and will need a relay for data and commands. Initially, the Comsat is mounted above the re-entry vehicle with a detachable truss, which is released in lunar orbit. The Lander portion has legs, thrusters, and a drill to take samples and travel to drill sites. The center portion of the Lander is the Ascender, which is used to take the samples and re-entry vehicle back to Earth. Computer: Command and Data Handling Subsystem (Spectrum Astro) Payload data handling up to 960 Mbps Downlink data rates over 50 Mbps. Chosen for modularization and flight history. Antennas: To minimize pointing requirements, an array of wide beam-width, low gain antennas (120 degrees, gain 4) will relay between orbiter and lander. The antenna is compact, efficient, and compatible with NASA and ESA equipment. Lander battery pack: 51kg Voltage 29V Max current 32Amp Peak output 640W Energy 16000Wh Ascender battery pack: 15kg Voltage 29V Max current 10 Amp Peak output 290W Energy 5000Wh Power system: Primary Batteries (LiSO2) Lander Tanks: Propellant management device (PMD) Pressure Regulated (Helium) Main Thrusters: Three 4.0 kN Thruster (Aerojet) MON/MMH Fuel type was chosen for its storability and flight history. Attitude Control System (ACS): Star Tracker Sun Sensors Inertial Measurement Unit (IMU) 16 10 N Thrusters (Aerojet) Ascender Tanks: Propellant management device (PMD) Blowdown (Helium) Legs Dampened Feet design to distribute weight Similar to Apollo Robotic Arm: Two segements Used to move and orientate drill to take and stow samples. Frame (AlLi ): Static and dynamic tests at 10g (simulation) Small deflection (<1cm) Strees well below yield stress Stage VI: Hopping #1 [130 hrs] First surface maneuver to travel to second drilling site (low right). Second set of core samples are drilled and stowed. Hopping: Flight time ~ 20 seconds Fuel used ~ 20 kg Maximum acceleration ~ 7 m/s2 Low impact velocity ~ 3.4 m/s Sonic Drill: Developed by JPL Laboratories Low required axial force = 10 N Drill rate ~ 0.1 mm/s Interchangeable drill bits (length = 0.20 m, dia. = 0.01 m) Drill depth reached = 2 m Detailed diagram to the right. Sensors (on arm): CCD Camera w/strobe Neutron Spectrometer (HYDRA, compact) Contact sensor Thermocouples Light Detection and Ranging (LIDAR): Developed by JPL Laboratories (“LAMP” system) 3 Dimensional images Resolution ~ 0.1 m3 Scan rate of 10 kHz Re-Entry Vehicle (Lavochkin): Inflatable Re-entry Descent Technology (IRDT) -to the right. Adaptable storage container Loading Mechanism All bits and samples are stowed into re-entry vehicle container Accessible by drill and robotic arm Lunar South Pole Sample Return Mission

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Stage IX: Re-Entry [350 hrs] Re-entry vehicle detaches from Ascender prior to entering Earth’s atmosphere. The re-entry device inflates during re-entry and lands on terrain. Stage VIII: Trans Earth Cruise [280.5 hrs] A couple of course corrections are made. Stage X: Retrieval [350.1-374 hrs] The samples can remain cool for days while being retrieved. Cost Estimates Two methods of estimating costs of this mission were used: parametric and analogous. Parametric cost models were taken from the Unmanned Spacecraft Cost Model database and use to calculate the cost of each subsystem(below). A more general method, analogous cost modeling, was used to estimate the entire mission cost with a model from NASA. Outreach The initial mission architecture and design were already presented before faculty at the University of Washington and members of the surrounding industry, including employees from Aerojet and Tethers Unlimited, Inc. Later this year, each subgroups of the entire design team will select a high school to present the overall design and their respective parts. Each group was asked to not only present the material at a high school level, but present a lesson specific to the class and the design. Subgroups are formed from the different aspects of the mission (i.e. structures, lunar operations, thermal regulation, etc.) As an example, the structures subgroup may have decided to lecture in a physics class and give a lesson about moments and torque with respect to the robotic arm. Acknowledgments Members of the University of Washington Senior Design Team (2005) would like to thank the following people: The Lunar and Planetary Institute RASC-AL Forum Charlie Vaughn (Aerojet) Jeff Slostad (Tethers Unlimited, Inc.) Jon Walker (Goddard) Dan Toomey (Spectrum Astro) Larry Matthies (JPL) UW Faculty Conclusions In this mission to return lunar regolith samples for ice exploration, using a hopping vehicle to transport from site-to-site was ideal for the areas of permanent darkness. Hopping maneuvers were kept simple and the hopping trajectories were predetermined, making the maneuvering more predictable and safe. Successful hazard avoidance systems coupled with LIDAR exist for terrain applications and will need to be developed for future robotic missions to Mars—a mission to the moon using such a system would be an ideal testing situation. The complexity of the subsystems was reduced by using a simple power system (batteries) and having simple thermal management devices (MLI and electric heaters). Additionally, the cost estimates suggest that this mission is low cost ($574 million)—lower than NASA’s cost cap on the “Moonrise” proposal ($700 million). With a more detailed cost analysis, one might expect the costs to go down to account for the simpler electric and thermal systems. Future Work This is a continuing design course that ends in June 2005—much more work is needed. The following must be research and designed: Reliability evaluation, probability of failure Structure attaching the communications satellite to the spacecraft during cruise Launch vehicle adapter Additional work will be done on the current design: The descent can be optimized for hazard avoidance. Computer simulation of the entire spacecraft in motion. The lander frame could be reduced and optimized. Re-entry vehicle needs to be further adapted to this mission Temperature models can be improved for unsteady solutions Some work that has been done between the submission of the paper and the conference: Loading mechanism has been designed with more detail. The main thruster configuration was changed from three thruster to four, where three of the thrusters would remain on the Lander, while one thruster would be part of the ascender. The drill and drill bit designs have been refined. Right: The ascender with re-entry vehicle, tanks, ACS, and thrusters--originally center section of the Lander, pyrobolts are used to decouple the two frames prior to ascent. Right: A updated drawing of the loading mechanism with the bit/sample container with guide rods below the re-entry vehicle. Above: Details of the bit design, including (starting upper left, going clockwise) the inner texture of the bit, the bit interface with the drill, and the bit face. Lunar South Pole Sample Return Mission