UNIVERSITY OF NOTRE DAME INSPIRATION MARS DESIGN TEAM

Transcription

1 UNIVERSITY OF NOTRE DAME INSPIRATION MARS DESIGN TEAM

2 University of Notre Dame Inspiration Mars Design Team 2. UNIVERSITY OF NOTRE DAME INSPIRATION MARS DESIGN TEAM MISSION ARCHITECTURE PROPOSAL Sarah Jackson Matthew Kudija Sebastian Ortega Brian Quinn Ryan VanDeCasteele March 15,

3 University of Notre Dame Inspiration Mars Design Team 3 Abstract The Notre Dame Inspiration Mars Design Team, consisting of five undergraduate students in the Department of Aerospace & Mechanical Engineering, proposes the mission architecture for a free-return Mars flyby mission. The mission is to launch in late 2017 to take advantage of the low energy free-return trajectory published by the Inspiration Mars Foundation and carry two crew members within 100 km of the surface of Mars before returning them safely to Earth. The Team builds o the work of the Inspiration Mars Foundation and proposes a mission architecture that relies less heavily on NASA support to reduce mission schedule risk as compared to the IM proposal. The mission launch architecture uses a modified SpaceX Dragon capsule to deliver the crew to orbit and ensure their safe reentry, a modified Orbital Sciences Cygnus module for crew habitation, and a newly designed upper stage for trans-mars injection. These vehicles are launched aboard SpaceX Falcon 9 and Falcon Heavy rockets currently in production or under development. The crew selection and training process is carefully analyzed to ensure the crew can meet the intense physiological and psychological demands of an extended deep space mission. Their work during the mission includes a number of science experiments to further our understanding of technologies required to eventually settle and terraform Mars. This mission architecture study provides a comprehensive vision for successfully completing the objectives put forward by the Inspiration Mars Foundation to design a Mars flyby architecture that is as low cost, safe, and operationally simple as possible. Future work for this Team will further analyze this mission architecture proposal and examine this proposal in the larger context of future manned space exploration.

4 University of Notre Dame Inspiration Mars Design Team 4 Contents 1 Nomenclature 6 2 Project Background Inspiration Mars Proposal Mission Purpose Notre Dame Design Philosophy The Notre Dame Approach Risks of Using Commercial Launch Providers The Notre Dame Team Proposed Mission Architecture Overview & Concept of Operations LEO Test Mission Flyby Mission Launch Vehicles Capsule Selection Habitation Module (HAB) Selection Upper Stage Design Reentry Reentry Burn Feasibility Landing & Recovery Spacecraft Systems Attitude Control Environmental Control Power Solar Flare Protection Navigation Communication Crew Considerations Selection & Training Crew Duties Pre-Flight On-Orbit Safety Crew Launch Safety On-Orbit Safety Science Mission 25

5 University of Notre Dame Inspiration Mars Design Team 5 7 Mission Summary Documents V Budget Mission Specifications Summary Mission Cost Estimates Vehicle Development & Integration Time-line Hardware and Vehicle Development Testing and Verification Mission Execution Other Ideas 31 9 Future Work Mission Graphics Acknowledgments 37 List of Figures 1 Mission Risk Analysis: SLS Launch Mission Risk Analysis: Non-SLS Launch Plot of Predicted Sun Spot Count for Cycle 24 (NASA Image) Mission Vehicle Stack Mission Launch Sequence Mission Map List of Tables 1 Mission Architecture Launch Summary V Budget Component Masses Launch Vehicle Specifications Dragon Capsule Specifications Cygnus Habitation Module Specifications Notre Dame Team-Designed Upper Stage Specifications Mission Cost Estimates Vehicle Development and Integration Time-line

6 University of Notre Dame Inspiration Mars Design Team 6 1 Nomenclature Abbreviations and Acronyms ACS Attitude Control System ATV Automated Transfer Vehicle ECASS Enhanced Course Analog Sun Sensor ESA European Space Agency EI Entry Interface F9 Falcon 9 Rocket, Developed by Space Exploration Technologies Corp. FH Falcon Heavy Rocket, Developed by Space Exploration Technologies Corp. HAB Habitation Module IM Inspiration Mars IMF Inspiration Mars Foundation ISS International Space Station L/D Lift to Drag Ratio LEO Low Earth Orbit NASA National Aeronautics and Space Administration Orbital Orbital Sciences Corporation PLA Payload Adapter RF Radio Frequency SLS NASA Space Launch System SpaceX Space Exploration Technologies Corp. STEM Science, Technology, Engineering, and Math TMI Trans-Mars Injection Burn TPS Thermal Protection System US Upper Stage VCD Vapor Compression and Distillation Greek V Change in Velocity (Delta-V)

7 University of Notre Dame Inspiration Mars Design Team 7 2 Project Background 2.1 Inspiration Mars Proposal The Inspiration Mars (IM) competition, sponsored by the Mars Society and based on the Inspiration Mars Foundation s (IMF) Architecture Study Report Summary [1], serves as the foundation for the Notre Dame team s mission design and feasibility study. In the Foundation s proposal, a Cygnus-derived habitation module (HAB), a modified Orion Multi-Purpose Crew Vehicle, and a currently undeveloped upper stage, are boosted to orbit aboard a single NASA Space Launch System (SLS) launch vehicle. Shortly thereafter, the mission crew launches aboard a commercial capsule and docks with the vehicle stack in Low Earth Orbit (LEO) before completing the trans-mars injection (TMI) burn. Mission termination occurs after the Orion capsule reenters Earth s atmosphere at 14.2 km/s and splashes down. In addition to the detailed analysis of how to accomplish the flyby mission, the Foundation also went to great lengths to describe why such a mission is beneficial both to the manned space program and the human race. Recognizing the value of this analysis, the Notre Dame team concentrated its e orts on improving the architecture presented to increase the operational simplicity of the mission, the safety of the crew, and the cost e ectiveness of such an endeavor. 2.2 Mission Purpose The Inspiration Mars mission is an important step in the e orts of the United States to pursue manned exploration of our solar system s celestial bodies and an essential opportunity to test many of the technologies and procedures that will be required for future landings on Mars, currently forecasted to take place in the early 2030s. Long duration manned spaceflight remains a largely untested frontier of human experience waiting to be explored and, as with Apollo, lessons learned in simply reaching Mars with a human crew are transferable to many aspects of subsequent landing attempts. Perhaps as important as paving the way for landings on Mars, the successful completion of the IM mission would bolster the American spirit and economy in a multitude of ways. As scheduled, all launch vehicles and spacecraft modules are to be designed and manufactured in the United States, increasing the need for skilled work in the space sector that has diminished with the retirement of the Space Shuttle. Additionally, the mission provides a unique platform from which NASA s O ce of Education can continue to promote the benefits of working and training in the STEM professions. Ambitious missions such as this provide the inspiration for school children to further pursue their technical studies and will spawn the next generation of driven leaders in the space industry, continuing the push by the Inspiration Mars Society and NASA to send humans farther from Earth than they have ever been before.

8 University of Notre Dame Inspiration Mars Design Team Notre Dame Design Philosophy Paramount to the design philosophy employed by the Notre Dame Team was to build on the strengths of the existing IM architecture proposal while reducing its reliance on NASA resources in certain strategic areas to mitigate launch schedule risk. As outlined in Subsection 2.1, the IM proposal relies heavily the NASA-funded Orion Multi-Purpose Crew Vehicle and Space Launch System. Although the current development schedule calls for completion of the SLS by 2017, this leaves only a small window of opportunity in which to conduct tests before the launch of the IM mission. Requested NASA support also extends to areas such as heat shield development, mission operations facilities and infrastructure, and significant monetary provisions. These place tremendous pressure on the agency to redistribute personnel and resources from its other chartered programs to meet deadlines for the Inspiration Mars mission. The SLS s extended development period is a consequence of NASA s unstable political position in recent years. Decreased funding from Congress and an inconsistent vision for the future have made strategic resource allocations challenging. Because of this, NASA engineers have had to design the SLS to accommodate a wide range of possible mission profiles, complicating its construction and delaying its flight readiness date. Stepping outside of specific hardware development, NASA s response to the original IM proposal has left little doubt that the expected level of involvement from the agency is unfeasible for this mission. Shortly after that proposal NASA Associate Administrator for the O ce of Communications David Weaver stated, The agency is willing to share technical and programmatic expertise with Inspiration Mars, but is unable to commit to sharing expenses with them. However, we remain open to further collaboration as their proposal and plans for a later mission develop. 1 If NASA is unable to share expenses for this mission the cost of an SLS launch itself would likely put the mission cost above an acceptable level The Notre Dame Approach The mission architecture proposed by the Notre Dame Team tempers the expectations of what mission planners could hope to gain from a NASA partnership and structures its proposal in a way that provides mutual benefits to both the Inspiration Mars Foundation and NASA without over-taxing the resources of either organization. The Notre Dame proposal specifically omits use of the SLS and Orion, replacing these with hardware from the Space Exploration Technologies Corp. (SpaceX). Despite the move away from using NASA vehicles, the proposal still relies heavily on NASA s operational infrastructure and technical support. These are areas where NASA is uniquely positioned to support this mission, and expenses accrued in these areas are more easily assimilated into the agency s yearly operating budget. As part of this agreement, mission planners would gain resources and support for heat shield development, crew training, and flight operations. 1 Smith, Marcia S. Tito Now Wants NASA Funding for Inspiration Mars - Here s NASA s Response, (November 21, 2013).

9 University of Notre Dame Inspiration Mars Design Team Risks of Using Commercial Launch Providers Because the Notre Dame proposal relies exclusively on commercial spacecraft hardware particularly from SpaceX it is therefore necessary to analyze the risks associated with SpaceX as a mission partner to ensure the possible challenges it may face in delivering the required hardware on schedule will not surpass those expected of NASA s SLS development track. The first potential obstacle to a SpaceX partnership with the Inspiration Mars mission is the interest of CEO and Chief Designer Elon Musk in such an endeavor. In the recent past, the SpaceX name has become associated with multiple ambitious proposals of untested feasibility. Mars One 2 and StratoLaunch 3 are two such examples. Both of these particular missions would require extensive modification of SpaceX hardware and do not necessarily fit into Musk s vision for SpaceX. Cognizant of these examples, the Notre Dame Team believes that the Inspiration Mars mission objective aligns more closely with Musk s vision because it is a tangible step toward the future colonization of Mars, a venture that Musk has publicly identified as his primary goal for SpaceX. 4, 5 An equally aspect of commercial partnership to consider is profitability, which must be achieved to justify involvement in the project. To increase the appeal of the IM mission to SpaceX, significant e orts have been made by the Team to reduce the necessary modifications to the Dragon capsule and Falcon family of launch vehicles, both of which are employed in the Notre Dame Team s mission design. In particular, the goal of the Notre Dame Team is to require no more extensive modification to the Falcon launch vehicles than would be required for a satellite launch. Dragon will require a new heat shield and thermal protection systems to withstand reentry, but this development will support future e orts supporting the long term vision of Mars colonization. This places the requirement on the Inspiration Mars Foundation to secure the necessary funding to complete these design changes, but this requirement has already been well defined in the IM architecture proposal. 2.4 The Notre Dame Team An understanding of the goals of the Notre Dame Team (the Team) also puts the proposed mission architecture into perspective. The Team consists of five aerospace engineering undergraduate students, led by Matthew Kudija and advised by Professor Thomas Corke. As undergraduates interested in pursuing careers in manned spaceflight, the primary goal of this study is self-education. Specifically, the Team s goals for this project are twofold: 1. To gain knowledge and experience in human space mission design. 2. To provide analysis to enhance the feasibility of the Inspiration Mars mission by proposing architecture that optimizes the cost, safety, and simplicity of the mission. 2 Mars One: 3 Space News, Orbital Sciences Replaces SpaceX on StratoLaunch Project, (December 3, 2012). 4 Coppinger, Rob. Huge Mars Colony Eyed by SpaceX Founder Elon Musk, (November 23, 2012). 5 Metzger, Philip. How Feasible Is Elon Musk s Idea To Establish A Colony On Mars In The 2020s? (January 17, 2014).

10 University of Notre Dame Inspiration Mars Design Team 10 Despite its general lack of resources in experience, time, and expertise compared to the Inspiration Mars Foundation and NASA, the Notre Dame Team is confident in its ability to critically assess the Inspiration Mars mission proposal with an eye for identifying risks and providing solutions that had not been addressed. These e orts take their focus from the main design drivers to make the mission as cost e ective, safe, and operationally simple as possible. The Team has no formal organizational structure and meetings are designed to foster collaboration. While all decisions were finalized as a team, each member specialized in specific aspects of the mission architecture design as detailed below: Sarah Jackson: trajectory, reentry, capsule assessment Matthew Kudija: overall architecture and team organization, report, graphics Sebastian Ortega: upper stage analysis, subsystems, science mission, website Brian Quinn: crew training, schedule, architecture concepts, website Ryan VanDeCasteele: HAB and upper stage assessment, science mission, graphics The Team website summarizes the contents of this report and is available here: With education as the Team s primary goal, a substantial fraction of the early design process was dedicated to research into human spaceflight and the details of the IM proposal. From there the Team conducted a risk analysis of the IM plan to identify the most significant barriers to mission success, all of which are presented in Figure 1. The abscissa represents level of di culty for the implementation of a solution to each risk and the ordinate, the severity of that risk. The dotted line represents the limit condition, above which mission failure occurs. As the analysis indicates, many issues related to launch aboard an SLS vehicle have the potential to pose significant risk to mission completion. Following its own approach, involving launch without use of the SLS, the Notre Dame Team conducted an identical risk analysis, the results of which are given in Figure 2. As the graphic clearly illustrates, the ease of solution for many mission-critical barriers are vastly improved with a non-sls mission architecture. This graphic was consulted throughout the entire design process to ensure that the primary drivers for mission success were addressed.

11 University of Notre Dame Inspiration Mars Design Team 11 Figure 1: Mission Risk Analysis: SLS Launch. Figure 2: Mission Risk Analysis: Non-SLS Launch.

12 University of Notre Dame Inspiration Mars Design Team 12 3 Proposed Mission Architecture The Team s mission architecture has two primary phases: (1) a long-duration LEO test mission and (2) the final Mars flyby mission. In keeping with the design philosophy outlined in Section 2.3, the proposed vehicle launch architecture specifically does not include launch aboard the NASA SLS. Instead, the launches for both the LEO test mission and the flyby mission are spread over a number of launches on SpaceX Falcon 9 and Falcon Heavy vehicles. This mission architecture requires several automatic docking operations but significantly reduces total program costs while building in a long-duration LEO test mission and providing more opportunities for vehicle component checkouts prior to departure to Mars. 3.1 Overview & Concept of Operations LEO Test Mission The LEO test mission allows for testing of vehicle components, automatic vehicle assembly operations in orbit, and mission operations procedures. Furthermore, this allows a crew to test various procedures and examine the e ects of the flyby mission in a simulated environment with reduced risk, as the crew can reenter from LEO at any point in case of an emergency. The distances traveled by the flyby crew will cause the communications delay to increase as the spacecraft leaves Earth. Delayed communications will be simulated on the LEO test mission to match what the flyby crew would experience. The crew will train heavily in failure scenarios and operational procedures to allow them to operate with little assistance or entirely autonomously from the mission controllers on the ground in case an emergency occurs at a point when communications are significantly delayed. The windows of the HAB will be covered with displays during the LEO test mission. These project the view the crew would have at any point in the actual flyby mission to simulate the isolation, distance from Earth, and excitement upon arriving at Mars they flyby crew will experience in reality. For the LEO test mission, a dummy upper stage is used. This dummy upper stage contains the major structural elements of the upper stage and the docking mechanism for HAB attachment, but has no propellant or engine, which is not necessary for the objectives of the LEO test mission since no TMI burn is included. First, the dummy upper stage is launched to a parking orbit. When in place, the Cygnus HAB is launched. After completing system checkouts in orbit, Cygnus automatically docks with the dummy upper stage. Finally, the Dragon capsule with the LEO test crew launches, performs checkouts, and docks to the HAB. This final vehicle stack then proceeds on the LEO test mission, after which the crew reenters in Dragon. All three of these launches are aboard a SpaceX Falcon Flyby Mission The flyby mission is the culmination of the program to send astronauts within 100 km of the Martian surface. The LEO test mission architecture is designed to closely match the launch architecture of the final flyby mission to test all mission operations. Again, the upper stage is launched first into a parking orbit. For the flyby mission, this upper stage launches aboard a Falcon

13 University of Notre Dame Inspiration Mars Design Team 13 Heavy due to its increased mass with full propellant for TMI. This is again followed by Cygnus and Dragon launches, checkouts, and docking to complete assembly of the vehicle components. Once fully assembled and checked out, the crew initiates the trans-mars injection burn on January 5, The upper stage is jettisoned after completion of the TMI burn, and the crew is on its way to Mars. Work begins immediately on the science experiments included on the mission, detailed in Section 6. Anticipation builds as the crew approaches the Mars flyby. During the flyby, the crew revels in being the first two humans in history to enjoy a close up view of Mars from their tandem cupola windows in the Cygnus module. Work resumes on the science experiments after Mars fades from view and for the duration of the trip back to Earth. Shortly before reentry, the crew begins final preparations. These include jettisoning the Cygnus module filled with all spent consumables as well as the Dragon trunk, carried up until this point to house solar panels and provide protection for the heat shield. The Dragon capsule reenters and completes a water landing near waiting recover vessels. Recovery of the crew successfully completes this flyby mission. Complete illustrations of the mission operations sequence is shown in Section Launch Vehicles The estimated cost of the SLS ranges from $500 million to $1.2 billion per launch. Coupled with the additional crew launch aboard a commercial vehicle, as proposed in the IM architecture, total launch costs of the flyby mission fall between $600 million and $2 billion. By distributing the orbital insertion of mission vehicles over several launches on the less expensive SpaceX family of vehicles, the total launch costs can be reduced drastically to approximately $383 million. More importantly, this architecture allows for testing of vehicle components in LEO before the final flyby mission. A summary of the launches required, including payload, vehicle, and scheduled dates, are given in Table 1. Table 1: Mission Architecture Launch Summary Launch Payload Launch Vehicle Launch Date LEO Test Mission Dummy Upper Stage Falcon 9 12/12/15 Cygnus Falcon 9 2/1/16 Dragon Falcon 9 2/21/16 Flyby Mission Upper Stage Falcon Heavy 10/15/17 Cygnus Falcon 9 11/8/17 Dragon Falcon 9 11/29/19

14 University of Notre Dame Inspiration Mars Design Team Capsule Selection The dual purpose of the capsule is to deliver the crew to the HAB in LEO at the start of the mission and to protect the crew through reentry and recovery. There are many challenges associated with selecting a capsule, specifically one that will be able to withstand the intense heating during Earth reentry. Reentry speeds are expected to be as high as 14.2 km/s, much higher than any previously attempted reentry by a man-made object. The primary concern with this high velocity reentry is the heating on the capsule. The level of heating depends on the vehicle shape, entry speed and flight trajectory, atmospheric composition, the thermal protection system (TPS) material composition, and surface properties of the heat shield. Taking these factors into consideration, the team researched two possible options for the capsule: (1) the Orion Multi-Purpose Crew Vehicle, under development by Lockheed Martin and NASA, and (2) the SpaceX Dragon. These two capsules di er in mass and shape. The Dragon is about half the mass of the Orion and has a much steeper sidewall angle. The sidewall angle of a capsule has an impact on several aspects of capsule performance. A steeper sidewall angle improves the packing e ciency of the capsule, increasing the internal volume by more closely approximating a cylindrical cross section. The sidewall angle also drives the stability of a capsule in angle of attack. The more shallow sidewall angle of Orion creates two aerodynamically stabilizing positions: nose forward, and heat shield forward. When the heat shield is pointing forward, the capsule will stabilize in this position for reentry. The steep sidewall angle of Dragon simplifies this stability concern as the capsule strongly favors the heat shield forward position. Flight control of the capsule comes by controlling its lift-to-drag ratio, L/D. Non-zero L/D is achieved by moving the center of gravity away from the capsule geometric centerline, causing the vehicle to trim at a particular angle of attack. For flight control the capsule is rolled around its velocity vector, which alters the lift vector, and thus allows the capsule to be steered. The ability to steer the capsule provides the required degree of flight control which is crucial for accurately reaching the landing zone and ensuring a safe recovery by waiting recovery vessels. The steeper sidewalls of Dragon will experience more heating during reentry than the shallower sidewalls of Orion. This will require additional TPS on the sidewalls as well as the main ablative shield, slightly increasing capsule mass. The team selected the Dragon capsule for this crew launch and reentry for the mission. The notable advantages of a lighter capsule and the stability of the steeper sidewalls outweigh the detriment of the increased TPS mass. To the capsules benefit, the weightier Dragon will still be lighter than Orion, simplifying mission operations and reducing costs. Another important advantage of the Dragon capsule is its integration with a proven launch vehicle, the Falcon 9 and its unmanned variant s own successes in LEO. In contrast, Orion has yet to complete a spaceflight, manned or unmanned. With human rating modifications currently underway, and a suitable time-line for the additional thermal protection needed for a reentry from Mars, the Notre Dame Team believes Dragon is the best option for this mission.

15 University of Notre Dame Inspiration Mars Design Team Habitation Module (HAB) Selection The habitation module (HAB) serves three primary purposes: (1) it houses the primary vehicle systems detailed in Section 4; (2) it holds necessary cargo for the mission including both consumable and non-consumable items; and (3) it provides the required crew volume for executing science experiments and performing other daily tasks during the 501 day mission. The Team analyzed three di erent options for the HAB module derived from existing or planned spacecraft: the European Space Agency Automated Transfer Vehicle (ATV), the Bigelow BA 330 inflatable vehicle, and the Orbital Sciences Cygnus. The Bigelow concept of an inflatable habitation module is attractive for reducing spacecraft volume at launch while still providing the required habitable volume. Bigelow claims that the inflatable, nonmetallic structure o ers enhanced radiation protection. 6 Unfortunately, the BA 330 has not been tested and at 20,000 kg is much too large for this mission. The ATV has a large habitable volume and proven automatic docking capability at the ISS. The ATV is human rated, but not designed for a mission of this duration. Because the ATV is operated by the European Space Agency, procuring and paying for the vehicle could be challenging, particularly since ESA participation in the mission would likely look to existing NASA support in providing vehicles. The ATV also has a relatively large external volume bringing up concerns about fairing encapsulation. Finally, the dry mass of 10,500 kg is significantly greater than Cygnus and would require a much larger upper stage to provide the required V for TMI. Arriving at a similar conclusion as the IM Architecture Study[1], the Cygnus module presents a promising spacecraft from which to develop the final HAB module. The current enhanced Cygnus is slightly smaller than what will be needed to store consumable cargo, life support systems, and have space for the crew to live and perform experiments throughout their 501 day mission, meaning the module would be slightly enlarged to a final length of 6.9 m. The Cygnus has proven its ability to successfully dock with the ISS. Being produced by Orbital Sciences Corporation, a publicly traded American company, there are no procurement concerns. Compared to other options, the development and procurement costs are reasonable. In addition to the existing docking port on Cygnus on the forward end, which will dock with the Dragon capsule, the modified Cygnus HAB will have a structural docking mechanism on the aft end. This serves as a structural connection for mating with the Falcon 9payloadadapterduringlaunchandtheupperstagewhenassembledinorbit. The heavier modified Cygnus will launch aboard the Falcon 9, which has a higher launch capacity than the Antares. This mission will be the closest human eyes have come to Mars. The modified Cygnus will have two cupola windows installed to provide the crew unprecedented views of Mars when they fly within 100 km of the surface. 3.4 Upper Stage Design The upper stage is responsible for the trans-mars injection burn, which places the spacecraft on its free return trajectory around Mars. The primary design constraint for the upper 6 Bigelow Aerospace:

16 University of Notre Dame Inspiration Mars Design Team 16 stage is to provide a V of 3.7 km/s to the 12,302 lb fully loaded combination of Dragon and Cygnus. After an analysis of existing upper stages, the team decided to forgo current designs and create a conceptual design for an entirely new upper stage. Detailed design and development of the upper stage is to be completed by a private contractor. Preliminary sizing of the upper stage was completed using data from existing upper stages. The stages examined included a variety of propellant types, including solid, liquid hydrogen (LH2-LOX), and kerosene (RP1-LOX). Since the upper stage is launched on a Falcon Heavy with relatively large mass budget, the fairing volume is the primary constraint driving propellant selection. Therefore, RP1-LOX is desirable because of its higher density. Similarly, candidate engines for the upper stage include the Rocketdyne J-2X, the Rocketdyne RL-10, and the SpaceX Merlin 1D-Vac. Engine selection depends on development schedules, availability, and the propellant type chosen by the development firm. The upper stage is designed to fully fill the Falcon Heavy payload fairing. The payload adapter joining the aft end of the upper stage to the launch vehicle integrates with the upper stage engine mounting structure. This payload adapter is jettisoned prior to the TMI burn. A docking mechanism on the forward end of the upper stage will allow the vehicle to dock with the Cygnus HAB in orbit. The upper stage has reaction wheels for attitude control while awaiting Cygnus and Dragon in LEO. Integrated vehicle subsystems are discussed in detail in Section 4. Detailed design and fabrication of the upper stage will be contracted to an established spacecraft manufacturer who can guarantee delivery according to the mission integration schedule given in Section Reentry One of the major unknown elements for the mission is Earth reentry because the capsule will be returning at speeds of up to 14.2 km/s, which is faster than any man-made object has ever attempted to reenter the atmosphere. The two major di culties associated with the high-speed reentry are the significant heating on the capsule and g-forces on the crew. The Notre Dame team explored several possible scenarios to mitigate these issues. The first possible technique is an aerocapture, with the intent that this would reduce the g-loads on the capsule and crew. One disadvantage of aerocapture was that it could extend the mission up to ten days. Because the capsule power system is battery powered, this could in the worst-case scenario result in a loss of power, making it impossible for the crew to make it through reentry. Lastly, a significant disadvantage was that to date, no aerocapture of this type has been attempted. This is a major concern because if any anomaly arises, the crew could be lost. A second option for reentry is the standard direct reentry. Manned vehicles in the history of the space program have all used direct reentry. The major concern with direct reentry is the heating on the vehicle. Because the Mars mission will be returning at speeds as high as 14.2 km/s, about 2 km/s faster than any other man-made vehicle has attempted to return, there needs to be more development work on a heat shield able to withstand the heating associated with reentry at these speeds. There would have to be major developments in heat shield technology by 2017 for it to be plausible to have a heat shield of appropriate design. Another major concern of the direct reentry is the g-loads the capsule and crew

17 University of Notre Dame Inspiration Mars Design Team 17 would experience. The team decided the best reentry approach would be a skip entry. This scenario involves one or more skips o of the atmosphere to slow the vehicle down before final entry. The Apollo capsules and the Space Shuttle had the capability to perform this maneuver, but it was never tested. One advantage to a skip entry is that it will allow the capsule to bleed o energy and slow down while imposing a lower g-load on the capsule and crew. Another advantage to the skip entry is that many algorithms have been studied and tested in simulators, giving knowledge of how best to integrate and implement this technique. Askipreentryrequirespreciseguidance,whereashallowentryanglewouldresult in the capsule bouncing o of the atmosphere and a steep entry angle would result in the capsule enduring excess heating. With the advances in guidance technology and the history of simulator tests, the Team determined that a skip reentry is feasible for the 2018 mission. Due to the precision required for this entry, most of the Dragon propellant will be reserved for contingency maneuvering before and during reentry. Dragon s hypergolic propulsion system ensures that it will be fully operational even at the end of the mission Reentry Burn Feasibility Amajorcomplicationtoreentryisheatgenerationduetoexcessivelyhighspeed. Inorder to alleviate the stress during reentry the Team investigated the possibility of storing reserve fuel to perform a retrograde burn, enabling the capsule to slow down during reentry. As the Team plans to design its own upper stage, the overall size is variable and could, theoretically, expand in order to store fuel for the purpose of reducing speed upon Earth reentry. The team calculated the total propellant mass and upper stage height required to achieve desired Vs of -2 km/s and -3 km/s. The resulting masses were 41,000 kg and 60,000 kg of propellant, respectively, with required upper stage heights of 16 m and 24 m, respectively. These configurations would require the internal volume of the Falcon Heavy fairing to nearly double in order to accommodate a large enough upper stage engine, forcing the team to abandon plans for a retrograde reentry burn. 3.6 Landing & Recovery The Dragon capsule is targeted to land in the Pacific Ocean o of the coast of Baja California, descending under drogue parachutes deployed at 14,000 m. The three main parachutes will be deployed at 3,000 m to further slow down Dragon. 7 Aprivatecompanywillbecontracted to conduct recovery operations. The capsule will be recovered from the water by a crane and brought back to shore on a barge. This recovery procedure is used for both the LEO test mission and the flyby mission. 7 SpaceX COTS 2 Mission Press Kit:

18 University of Notre Dame Inspiration Mars Design Team 18 4 Spacecraft Systems The Team s work was primarily focused on the mission architecture (see Section 3). The Team analyzed spacecraft systems at a high level to provide guidance as these systems are developed for the final design. Certain aspects of these systems are likely to change in detailed design. For the most part the Team envisions detailed system design being the responsibility of the manufacturers for the vehicle module the system resides in. 4.1 Attitude Control The spacecraft Attitude Control System (ACS) will be housed in the Cygnus HAB module. Errors in roll, pitch, and yaw will be assigned to the X, Y, and Z-axis, defined using a conventional coordinate system[2]. Attitude error will be measured in reference to the an on board Star Tracker and Enhanced Course Alignment Sun Sensor (ECASS). 8 The navigation system is detailed in Section 4.5. Attitude correction will be conducted using thruster firings for course adjustments and reaction wheels for fine adjustments. A measure of the rate of correction will be determined using gyroscopes installed on the spacecraft. Thrusters will work in pairs to cause rotations in the spacecraft. Thrusters will fire in opposite directions around a given directional axis to create the desired attitude adjustment, requiring six thrusters. However, each element of the attitude control system (thrusters and reaction wheels) will have a secondary system in case of system failure. Therefore twelve thrusters will be installed, with only one set of six active thrusters operating at a time. Reaction wheels will be used for a fine adjustment of spacecraft position. Three reaction wheels will be needed to generate the fine correction in roll, pitch, and yaw. Since the momentum generated by a rotating disk will act along a unit normal vector extending from the center axis of rotation, one reaction wheel will be enough per direction. To account for positive and negative errors, the polarity of rotation can be switched to cause positive or negative momentum vectors. 4.2 Environmental Control The Team considered the four basic divisions of environmental control: (1) atmosphere, (2) water, (3) waste, and (4) food. Given the duration of this mission, a closed loop system will be implemented for each to reduce the total mass of consumables. The atmosphere within the spacecraft will be composed of a nitrogen-oxygen mix. To reduce the risk of fire on board the vessel, oxygen will be kept below 30 % of the nitrogenoxygen mixture. In order for the crew to be able to breath and function a p O 2 will be kept between 13.4 kpa and 19 kpa[2]. Furthermore, a ventilation system will be implemented within the vessel to filter contaminants from the air and ensure continuous mixing. Water will be recycled using two filtration systems: a phase change system, and a filter based system. A phase change system will be implemented for body waste fluids such as urine. A phase change system will be used for these fluids due their contamination and salt 8 Loral Patent:

19 University of Notre Dame Inspiration Mars Design Team 19 content. The technology adapted for this system will use vapor compression and distillation (VCD). This system is already in use by the United States and has the capability to recover more than 96% of the water within these waste fluids[2]. Other water waste generated from less contaminated sources such as shower water will be recycled using conventional filters. Replacement filters will be carried on board for the duration of the mission. Do to the complexity of recycling solid waste, solid waste will be compressed and stored in air tight containers. These containers of waste, especially food waste, will be used to provide additional radiation shielding. In regards to food regeneration, majority of the food that will be consumed during the mission will be carried from the initial launch. A small portion of the food may be grown during the mission as part of the science experiments included, detailed in Section Power Solar panels on the Cygnus HAB and Dragon Trunk provide electrical power for the duration of the mission. These are supplemented with lithium ion batteries for backup power storage, as well as to power the Dragon capsule during reentry after separation from the Dragon Trunk. Due to the required power and the length of the trip to Mars, a Gallium Arsenide [GaAs] photo cell was chosen over the conventional silicon photo cells. GaAs photocells have a e ciency of available cells of 18.5 % as compared with 14.8 % for silicon photo cells. Detailed power system design is included in the modifications to Cygnus for the IM mission. 4.4 Solar Flare Protection With the launch date of the spacecraft on November 29, 2017 and the anticipated return of the crew on May 21, 2019, the mission will take place near the end of Solar Cycle 24. Sun spot count is predicted to reach a low during the time of the IM mission as highlighted in Figure 3. The crews sleeping quarters is designed for additional radiation protection. Significant radiation protection can be added in the sleep quarters at a relatively minor mass penalty due to their small size. With 8 hours of sleep per day, the crew will spend one third of their total time in their sleep quarters. If radiation exposure can be reduced to minimal levels during this period, a higher dose of radiation (with the associated reduction in radiation shielding mass in the rest of the module) will result in an average radiation exposure level over the course of the mission that is at acceptable levels. In addition to the creation of a radiation safe haven around the sleeping area of the crew, the Team also proposes the use of the water storage and filtration system as a means of additional radiation protection by residing in the spacecraft walls. Hydrogen based substances, in this case water, are e ective at blocking neutron radiation. The use of the water system in this manner of radiation shielding will make up for the ine ciency of lead in blocking neutron radiation. Since lead is much more dense than water, uncharged neutrons can pass through the material and expose the crew to a di erent form of radiation. 9 9 Thomasnet:

20 University of Notre Dame Inspiration Mars Design Team 20 Figure 3: Plot of Predicted Sun Spot Count for Cycle 24 (NASA Image). 4.5 Navigation To ensure that the spacecraft remains on the selected trajectory, a Star Tracker and an Enhanced Course Analog Sun Sensor (ECASS) system will be installed. The primary sensor is the Star Tracker, which will be installed with a redundant unit in case of system failure. The system computers will be programmed with the stars that the each star tracker is to view during the duration of the spacecraft journey. Using Quaternion Algebra, the observed star pattern from the mounted Star Tracker will be compared to the position of the predicted star pattern programmed before launch. The error between these two star patterns will be calculated and translated into roll, pitch, and yaw error by on-board computers. Star Trackers are currently being used in satellites from companies such as Boeing and SSL as well as other spacecraft. The backup to the star tracker will be the ECASS developed by SSL in California. The ECASS uses photo cells mounted aboard the spacecraft to detect the solar radiation emitted from the sun. The photo cells are mounted in a perpendicular L shape arrangement in order to determine the error in roll, pitch, and yaw. All positions required for these instruments are to be computed before launch. 4.6 Communication Communications systems ensure contact between the crew and ground controllers throughout the mission. The communication delay experienced as the crew departs Earth will be

21 University of Notre Dame Inspiration Mars Design Team 21 simulated during the LEO test mission. Due to the short development window traditional RF communications systems will be used. Future missions to Mars may look into LASERCOM designs in order to accommodate higher data rates to improve items such as higher resolution video. The fixed variable for communications system development is the antenna size. Antenna size needs to be fixed in order to maintain the mass budget required for the launch sequences of the Team s proposed mission architecture. Detailed design of the antenna will balance the transmission power and data rate given a fixed antenna size[2]. 5 Crew Considerations 5.1 Selection & Training The proposed Mars flyby mission although unique in its objectives, distances traveled, and certain aspects of the space environment is a direct analog to current long-duration spaceflights aboard the International Space Station in terms of many aspects of crew training. The major constraint of the mission architecture, however, is the limited time period in which crews must train for such a mission before launch in late Current NASA astronaut training includes approximately two years of intense basic skills training before placement on active status, and an additional three years of missionspecific training for ISS crew members. Although likely to be reduced slightly given the relative simplicity of HAB operations compared to those of the ISS, the time requirements for selecting a new class of NASA astronauts and training them fully for the flyby mission leaves only a few months before selection must occur. Crew volunteers for the mission will be solicited first from the existing astronaut corps to take advantage of their experience. Understanding that candidates may need to be solicited outside of the astronaut corps, and therefore require more training, the selection process begins immediately. Although the original proposal calls for a married couple with a strong bond[1], the Notre Dame Team believes that a pair of astronauts with an excellent working relationship will exhibit the same qualities desired for the flyby mission crew. Geared primarily towards increasing crew preparedness, the training window expansion also opens launch opportunities for a crewed test mission in LEO. This gives both ground controllers and crew an opportunity to familiarize themselves with vehicle operations as well as assessing the functionality of the flight configuration with full abort capabilities available to the crew at all times in flight. This full duration LEO mission also serves to evaluate many of the safety and human factors concerns for the Mars flyby mission. With these qualifying aspects of the mission in mind, crew selection preferences and requirements were established. Again, the crew is selected from the current pool of active astronauts, eliminating the time and expense usually spent advertising the position and verifying the general qualification of candidates. From this group, strong preference will be given to those who have served on previous long-duration missions aboard the ISS. Although this designation does not preclude selection of non-flown astronauts, the Notre Dame Team is of the opinion that an astronaut who is a veteran of an ISS Expedition crew will be more able to identify systematic issues with flight processes or control configurations early on

22 University of Notre Dame Inspiration Mars Design Team 22 in the hardware development process, preventing issues during the mission for the test or flyby crews. As in the early years of the United States space program, astronaut training will coincide with flight vehicle development and modifications, allowing members of each crew to specialize in and supervise development of specific subsystems to meet their needs. Therefore experience with similar hardware is desired. Such candidates would also be ideal based on their experience occupying free time in orbit, which is a valuable resource given that both crews will likely face extended periods of unscheduled time. There are certain desirable career areas which would give candidates preference in the selection process. For instance, submariners have experience with long-duration missions in small enclosed spaces. Any candidate with exceptional performance in similar conditions to the flyby flight environment will be given preference due to their familiarity with mission-like environments. 5.2 Crew Duties Pre-Flight Critical to the crew s training and understanding of vehicle systems is crew involvement in the final design and testing processes of the vehicle components. All vehicle components require some additional level of development giving the crew a unique opportunity to train while the process continues. Similar to the Apollo program, the eight astronauts selected for the LEO and flyby missions will be split into four crews of two, with both a primary and a backup crew for each mission. One member from each mission crew will select one of the flight s habitation vehicles, Dragon or HAB, and will specialize their knowledge and serve as a crew liaison to the individual manufacturing companies, ensuring the vehicle designs and layouts are comfortable and accessible for the long duration missions. As is the goal for the entire mission, this individualized training will allow the crew to essentially train on four separate tracks in the first quarter of their mission preparation time-line, allowing them more opportunity to learn and practice more mission-critical scenarios. Since the LEO mission must fly at least two years in advance of the flyby mission, the two crews assigned to the flight will begin mission-specific procedural training approximately eight months before their launch in February Borrowing this estimate from training schedules of Space Shuttle crews, this will allow ample time for the crews to simulate their flights and various failure scenarios while also allowing them to become familiar with their on-orbit accommodations. As the LEO mission has the constant option to abort the mission and return home, many of the failure scenarios that would need to be simulated for the flyby cruise would be irrelevant in low Earth orbit, again saving on time and increasing the feasibility of the dress-rehearsal mission. However, training for these scenarios will continue for them while in orbit. While the LEO crew trains, the flyby crew will continue supporting hardware development and integration to ensure final vehicle readiness in time for launch. Once this process is completed, the crew will move into its own mission readiness training, primarily concerned with navigational emergencies and communication methods. Additionally, the Mars trajectory leaves only a narrow window of opportunity for the crew to declare an emergency