International Student Design Competition for Inspiration Mars Mission Report Summary (Team Kanau)
Jeff Stuart1, Ashwati Das1, Max Fagin1, Kshitij Mall1, Shota Iino2, Eriko Moriyama3, Ayako Ono4, Takuya Ohgi5, Nick Gillin6, Yuri Aida1, Koki Tanaka1, Daichi Nakajima7 Professor Hiroyuki Miyajima8, Professor Michael Grant2 1
Purdue University (USA), 2Keio University (Japan), 3International Space University (Alumnus, France), 4Tohoku University Graduate School of Medicine (Alumnus, Japan), 5Nagoya University (Alumnus, Japan), 6Art Center College of Design (USA), 7Tokyo University of Agriculture and Technology (Japan), 8Tokyo Jogakkan College (Japan).
Team Kanau
Table of Contents 1
Introduction
1
2
Mission Overview
1
3
4
5
2.1
Vehicle Overview
2
2.2
Concept of Operations
3
2.3
Mission Cost
3
2.4
Safety Analysis and Design
6
2.5
Mission Schedule
10
Crew Factors
12
3.1
Inputs from Professionals
12
3.2
Kanau Spacecraft’s Interior Design
12
3.3
Facilities for Crew Physical Health
15
3.4
Facilities for Radiation Protection
16
3.5
Crew Selection from American Astronauts
17
3.6
Crew Daily Schedule
17
Detailed Mission and Vehicle Description
19
4.1
Trajectory Analysis and Launch Vehicle Selection
19
4.2
Attitude Determination and Control System (ADCS)
27
4.3
Environment Control and Life Support System (ECLSS)
28
48
4.4
Command & Data Handling Subsystem (C&DH)
31
4.5
Communications (COMMS)
32
4.6
Electrical Power Systems (EPS)
33
4.7
Thermal Control Systems (TCS)
36
4.8
Payload Mission
37
Outreach
38
5.1
Astronaut Fashion Competition
38
5.2
World Cancer Day Support
38
5.3
Student Mission Simulator
39
5.4
Thought Exchange
39
5.5
Further Design Competitions
39
6
Conclusion
39
7
Kanau Prospects for 2021 Mission Opportunity
40
8
Kanau Team Workflow, Website and Animation Video
41
9
Acknowledgments
42
10 References
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List of Figures FIGURE 1 RANKING OF ENGINEERING SPECIFICATIONS NEEDED TO ACCOMPLISH MISSION. .............................................1 FIGURE 2 KANAU VEHICLE STACK. ..................................................................................................................................2 FIGURE 3 CONCEPT OF OPERATIONS. ...............................................................................................................................3 FIGURE 4 RELATIONSHIP BETWEEN COST ESTIMATION METHODS AND MISSION PHASES. ................................................4 FIGURE 5 WORK BREAKDOWN STRUCTURE AND ASSOCIATED COSTS. .............................................................................5 FIGURE 6 COST SUMMARY. ..............................................................................................................................................6 FIGURE 7 COST DISTRIBUTION BY PHASE (M$). ...............................................................................................................6 FIGURE 8 COST DISTRIBUTION BY PHASE (%). .................................................................................................................6 FIGURE 9 SUBSYSTEM RELIABILITY FIGURES. ..................................................................................................................8 FIGURE 10 RISK ASSESSMENT PRIOR TO IMPLEMENTATION OF MITIGATION STRATEGIES ..............................................9 FIGURE 11 RISK ASSESSMENT AFTER IMPLEMENTATION OF MITIGATION STRATEGIES.................................................10 FIGURE 12 GANTT CHART FOR MISSION. ........................................................................................................................11 FIGURE 13 ASPECTS OF INTERIOR DESIGN. ....................................................................................................................13 FIGURE 14 ARTIFICIAL RAINBOW GENERATED BY SPOTLIGHT. "ROUND RAINBOW" BY OLAFUR ELIASSON, HIRSHHORN MUSEUM, USA. PICURE COPYRIGHT: AYAKO ONO 2008. .....................................................................................13 FIGURE 15 INSIDE THE SHELTER. ...................................................................................................................................14 FIGURE 16 INTERIOR DESIGN OF THE HABITAT. .............................................................................................................15 FIGURE 17 OUTLINE OF FACILITIES FOR CREW PHYSICAL HEALTH.................................................................................16 FIGURE 18 MORPHOLOGICAL CHART OF KANAU MISSION. ............................................................................................19 FIGURE 19 BALLISTIC FREE-RETURN TRAJECTORY WITH EARTH DEPARTURE IN JANUARY 2018, MARS FLYBY IN AUGUST 2018, AND EARTH RETURN IN MAY 2019. COLOR BAR INDICATES SPACECRAFT DISTANCE FROM EARTH. ..............................................................................................................................................................................21 FIGURE 20 MARS CLOSE APPROACH,. ............................................................................................................................21 FIGURE 21 KANAU VEHICLE STACK WITH EXPLODED VIEW OF PROPULSION MODULE. GREY CYLINDERS REPRESENT ADAPTERS BETWEEN THE BOOSTER STAGES AND THE VEHICLE. ............................................................................23
FIGURE 22 SCHEMATIC OF THE EARTH DEPARTURE STAGES..........................................................................................23 FIGURE 23 EARTH RETURN HYPERBOLA AND AEROCAPTURE ELLIPSE. ..........................................................................25 FIGURE 24 VARIOUS PLOTS FOR LIFTING AND BALLISTIC RE-ENTRIES. ..........................................................................26 FIGURE 25 FUNCTIONS FOR EACH ECLSS MANAGEMENT SUBSYSTEM..........................................................................28 FIGURE 26 LETTUCE HOME CULTIVATION KIT46. ............................................................................................................30 FIGURE 27 MATERIAL FLOW THROUGH ECLSS SIMULATOR FOR CLOSED LIFE, SICLE...............................................30 FIGURE 28 MATERIAL AMOUNT IN TANKAGES BY SICLE SIMULATION. ........................................................................31 FIGURE 29 KANAU SPACECRAFT COMMUNICATION LAYOUT. ........................................................................................33 FIGURE 30 THE ESTIMATED INSOLATION LEVELS ON THE SPACECRAFT FOR THE INTERPLANETARY TRAJECTORY. .......35 FIGURE 31 EXAMPLES OF PAYLOAD MISSION.................................................................................................................37
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List of Tables TABLE 1 ESTIMATED SUBSYSTEM MASS AND CONTINGENCY MASS OF THE VEHICLE. ......................................................2 TABLE 2 SUGGESTIONS BY PROFESSIONALS ON THIS MISSION. ......................................................................................12 TABLE 3 RADIATION DOSAGES FOR KANAU MISSION. ...................................................................................................17 TABLE 4 NOMINAL DAILY SCHEDULE ABOARD KANAU SPACECRAFT. ...........................................................................18 TABLE 5 CRITICAL EPOCHS OF THE EARTH-MARS-EARTH FREE-RETURN TRAJECTORY. ...............................................20 TABLE 6 VELOCITIES AND PERIAPSIS ALTITUDES OF THE DEPARTURE, FLY-BY, AND RETURN HYPERBOLIC TRAJECTORIES. ......................................................................................................................................................20
TABLE 7 POTENTIAL LAUNCH SITES...............................................................................................................................22 TABLE 8 PROPULSION MODULE PERFORMANCE AND DEPARTURE STAGING ORBITS.......................................................23 TABLE 9 LAUNCH TIMELINE AND MANIFEST. .................................................................................................................25 TABLE 10 BOUNDARY VALUE CONDITIONS FOR LIFTING RE-ENTRY. .............................................................................25 TABLE 11 COMPARISON BETWEEN DIFFERENT AERO ASSIST RE-ENTRIES OF ERV. .......................................................26 TABLE 12 SPECIFICATIONS FOR ATTITUDE CONTROL SYSTEM. ......................................................................................27 TABLE 13 MATERIAL CONSUMPTION OF HUMAN CREW (KG/CREW MEMBER-DAY).......................................................28 TABLE 14 MATERIAL PRODUCTION OF HUMAN CREW (KG/CM-DAY). ...........................................................................29 TABLE 15 MODULAR FUNCTIONS AND THEIR CONFIGURATION FOR DCDMS. ..............................................................32 TABLE 16 C&DH BUDGETS. ..........................................................................................................................................32 TABLE 17 COMMUNICATION FOR A MARS FLYBY MISSION. ...........................................................................................32 TABLE 18 COMMUNICATION ANTENNAS IN A NUTSHELL. ..............................................................................................33 TABLE 19 CRITICAL CONSIDERATIONS FOR POWER SYSTEM. .........................................................................................34 TABLE 20 THE VEHICLE POWER BUDGET. ......................................................................................................................35 TABLE 21 MAJOR TASKS PERFORMED BY INDIVIDUAL TEAM MEMBERS. .......................................................................41 TABLE 22 MAJOR ROLES OF TEAM'S ADVISORS..............................................................................................................41
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1 Introduction There is strong potential for a near-future, human mission to Mars, with Inspiration Mars and the Mars Society in particular advocating for a two-person, manned flyby mission within the coming decade. This document presents Team Kanau’s response to the challenge posed by the Mars Society for student groups to plan a flyby of Mars in the year 2018 by a male and female astronaut pair (Kanau, the Japanese word for “collaboration” and “synergism”, was chosen to symbolize our team members hailing from different universities of US and Japan). Building upon analyses published by Inspiration Mars, we investigate the requirements for mission architecture, including: spacecraft design, crew life support, launch vehicle selection, the flyby trajectory, and Earth capture and re-entry. Our final design places heavy emphasis on modifying existing hardware rather than the design of new systems, given only 4 years of allotted time between now and the 2018 launch. We give particular attention to ensuring the physical and mental health of the two astronauts for the duration of the journey. Specifically, regenerative air scrubbers, the combination of pre-packaged and grown food, as well as 3-D printing technology support the human crew during the 501-day flight. A novel combination of a Falcon Heavy launch to LEO and a ULA DCSS-enabled Trans-Mars Injection places the crewed spacecraft on a free-return trajectory, while aerocapture using bank angle lift modulation upon Earth return enables re-entry with benign heat-rate and G-loads. The technologies and design concepts that we propose, while requiring further development prior to the stated 2018 mission opportunity, are powerful enabling factors for the crewed Mars flyby as well as other manned deep space missions.
2 Mission Overview We provide a brief summary of our mission architecture, including the concept of operations for the interplanetary flight, our proposed budget, risk mitigation analysis, and development timeline. We performed a preliminary systems analysis using quality function deployment (QFD) to rank the engineering specifications needed to meet the requirements of this mission. We obtained results as shown in Fig. 1. In all stages of the mission feasibility analysis, we focus on two primary goals: (i) the human flyby of Mars, and (ii) the safe return of the crew to Earth. The need to decrease mission cost and complexity as indicated by Fig. 1 is also kept in mind. Hardware reliability ranks as the topmost technical specification for this mission and is, therefore, the means to achieve safety. The detailed systems analysis is described in the online Appendix.
Figure 1 Ranking of engineering specifications needed to accomplish mission.
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2.1 Vehicle Overview The Kanau spacecraft vehicle stack is approximately 23 meters long, roughly 5 meters in diameter and consists of three primary spacecraft – the service module, habitat and ERV. The service module (SM) is a modified version of the Automated Transfer Vehicle (ATV) developed by the ESA. This module was chosen due to its reliable flight heritage since 2008, and its potential to house custom ADCS, ECLSS, thermal, and power systems that help satisfy the requirements of the proposed mission. The habitat module is a modified MultiPurpose Logistics Module (MPLM), a major pressurized vessel employed to ferry cargo to and from the ISS. Its large Figure 2 Kanau vehicle stack. habitable volume is an attractive option to offer the crew a more comfortable journey. However, as the original MPLM has a docking adapter at only one end, the version used here is modified to possess docking adapters at both ends. Furthermore, a custom mating adapter must be developed to interface the ATV with the MPLM. As the SpaceX developed Dragon’s unmanned flight heritage to human rated capabilities, its already existing heat shield is a crucial design feature in increasing crew survival rates during high speed Earth re-entry. Thus, Dragon serves as the ERV for the spacecraft. The spacecraft is 3-axis stabilized and generates power using four Ultra-Flex panels developed by ATK (two are hidden from view in Fig. 2). The radiators aid in maintaining desired thermal loads on the spacecraft and a high gain antenna is used for communications between the spacecraft and the Earth. A brief summary of the spacecraft mass budget is presented in Table 1: Table 1 Estimated subsystem mass and contingency mass of the vehicle.
Subsystem
CBE Mass (mT)
Contingency
MEV Mass (mT)
% of Dry Mass
Crew Systems ECLSS EPS TCS C&DH Comm. ADCS SM Habitat ERV Mating adapter
4.31 8.63 1.45 0.47 0.08 0.08 1.09 5.12 4.08 4.2 0.85
20 % 20 % 30% 30% 30% 30% 10% 10% 10% 10% 30%
5.17 10.35 1.89 0.61 0.10 0.11 1.20 5.63 4.49 4.62 1.10 34.88
14.8% 29.7% 4.3% 1.7% 0.3% 0.3% 3.4% 16.1% 12.9% 13.2% 3.2%
Dry Mass
Note that CBE stands for Current Best Estimate and MEV stands for Maximum Expected Value. Mass contingencies are applied for each subsystem. A contingency of 30% is applied to smaller subsystems whereas a growth margin of 20% is assumed for Crew Systems and ECLSS due to
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less uncertainty arising from their more definite heritage on the ISS. The lower margin of 10% for ADCS reflects the choice of specified hardware: the mass margin is for connectors and supports for these devices of known mass. Finally, the masses of the individual vehicles comprising the spacecraft are known, however a margin of 10% is applied to reflect possible modifications and reinforcements. Note that the dry mass of the SM / ATV is lower than vendor supplied values as we remove large portions of the existing ATV systems and replace them with our custom designed systems for the Mars flyby mission.
2.2 Concept of Operations The mission timeline is divided into nine phases as shown in Fig. 3 and detailed below: I. Delta Cryogenic Second Stage (DCSS) launch atop SpaceX Falcon Heavy II. DCSS + Service Module (SM) launch [Falcon Heavy] III. DCSS + Crew Habitat + Earth Return Vehicle (ERV) launch [Falcon Heavy] IV. Vehicle stack assembly in staging LEO V. Earth departure and Trans-Mars Injection (TMI) VI. Mars flyby encounter VII. Earth return cruise and approach VIII.Jettison habitat and service module IX. Aerocapture, entry, descent and landing, and crew recovery
Figure 3 Concept of Operations.
2.3 Mission Cost The rapid growth and diversity of space missions, especially in the commercial and small satellites sector provide an opportunity for the development of cost models that help estimate the cost of a mission. Trivailo et al.1 provide the following overview of different types of methods that are best suited to different phases of a mission:
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Figure 4 Relationship between cost estimation methods and mission phases.
As per Fig. 4, the ‘pre-phase A’ timeline of this study benefits from cost estimation and assessment using a combination of methods such as ROM, Parametric Modeling and Loose/Close Analogy. The analogy-based model was used whenever cost values were found for components/processes with a similar scope of function/performance. It is important to note that heritage items or modifications to existing technology, especially those pertaining to crew life support systems have been incorporated wherever possible to lower the risks and associated costs. The parametric modeling method is incorporated when cost numbers are challenging to find for analogous systems and the parametric measures required for the Cost Estimating Relationships (CERs) are available. Various tools exist for performing such parametric cost estimates. Many of them are formulated based on near-Earth, small satellite and robotic mission data alone. So, extreme care must be taken to understand the limits where the outputs of these tools are invalid for crewed missions. Some of these tools are the Advanced Missions Cost Model (AMCM) developed by Johnson Space Center, Project Cost Estimating Capability (PCEC) developed by the Marshall Space Flight Center, Small Satellite Cost Model (SSCM) developed by the Aerospace Corporation and Unmanned Space Vehicle Cost Model (USCM) developed by the U.S. Airforce. The cost breakdown for Team Kanau was approached using NASA’s work breakdown structure (WBS) as found in Ref. [2] and shown in Fig. 5:
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Figure 5 Work breakdown structure and associated costs.
Referencing historical mission design, the preliminary analysis phase (Phase A) is calculated to be approximately 1% of the total crewed mission costs, whereas, 5-10% is advised to be allocated to the design phase (Phase B) to prevent cost overruns in phases C&D.3 These third and fourth phases (C&D) account for costs associated with the design, development, test and evaluation stages of the mission. The USCM tool used to calculate some of the spacecraft subsystems accepted inputs for both new and modified components and also enquired as to whether government or private vendors would be utilized. Where possible, private vendors were selected as this resulted in cost savings. Additionally, even though most selected subsystem components have flight heritage, a conservative approach was taken and the average of the new and modified component cost estimates were calculated. The ECLSS and the vehicles consisting of the ERV, Habitat, and Service Module seem to be the cost drivers of the C&D mission design phases. The costs associated with the ECLSS subsystem obtained using parametric methods were also satisfactorily cross-checked with the scaled values for FY2018 from NASA’s budget for the ISS ECLSS subsystem in FY2005.5 In Phase E, WBS 07 (Mission Operations) has been budgeted for astronaut training and also pre-flight and post-flight ground operations. The calculated ground operations cost was derived from analogous Shuttle operations budgets for 2009. The ground systems costs cover the expenses associated with astronaut training facilities, mission unique facilities and communications facilities to name a few.2 In order to acquire a ROM on the cost of this system, it is estimated to be approximately 30% of the Launch Vehicles and Services costs. This estimate is based on the assumption that pre-existing crewed mission facilities would be exploited rather than building new ones. The Launch Vehicle and Services costs include the cost of three Falcon Heavy launches and the associated facility fees and logistics support. Two elements of the WBS, 4 (Science/Technology) and 5 (Payloads), are not incorporated into the total cost as these represent specialized equipment required to satisfy
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scientific or technology demonstration purposes. As the goal of the mission is the crewed Martian flyby using currently available technology, no advanced equipment is required. Also, the costs incurred for medical experiments/monitoring, plant cultivation and 3D printing experiments in space are already absorbed by the crew systems cost analysis. Furthermore, WBS 11 (Public Relations) is not currently specified as we expect no great advertisement will need to be made to garner public attention for the crewed flyby. However, some additional cost could be included, if desired. Some ideas for public outreach have been provided towards the end of this document. A more detailed breakdown of the budget may be found in the online Appendix. Kanau therefore proposes a total mission budget of approximately $4.363 billion for FY2018. Additional cost reserves have also been incorporated according to ground rules followed by NASA4 in order to prevent cost overruns. Introducing 20% cost reserves on Phases A-D and 10% cost reserves on Phase E result in a total of approximately $5.153 billion for FY2018. An overall summary of the total costs and cost breakdown by mission phases are presented in Figs. [6-8] .
Figure 6 Cost summary.
Figure 7 Cost distribution by phase (M$). Figure 8 Cost distribution by phase (%). 4
According to Larson and Pranke , the costs associated with Phase E (operations) for finite and short-termed crewed programs (e.g. Apollo) were less than 20% of the total mission costs. It is satisfying to see that the summed cost from the contributing factors of Phase E that have been calculated here closely agree with historical data. 2.4 Safety Analysis and Design The inclusion of human crew on the interplanetary flight and Mars close approach entails a need for rigorous analysis of the probability of mission success and safe crew return to Earth. The reliability of the proposed mission design is investigated and risk mitigation contingencies are discussed.
2.4.1 Reliability Assessment Technology that is sent into space has to endure heavy loads during launch, followed by drastically varying temperature gradients and harsh radiation effects in the space environment. Furthermore, microgravity as a work environment, diminished crew power and even technical expertise leads to very challenging repair/replacement situations once the rockets have fired. So,
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it is important to subject subsystem components to rigorous reliability analysis before they are approved. We present the definition of reliability as used by Larson and Wertz6 – “The probability that a device will function without failure over a specified time period or amount of usage”. The reliability of a system is computed as: 𝑅! = 𝑒 !
!"
Where, 𝜎 is the failure rate of a particular element of interest. In instances where the ‘failure rate’ parameter is challenging to quantify, for example the reliability associated with the ERV, other techniques such as the Bayesian probability of success7: 𝑅! =
𝑁𝑆 + 1 𝑇𝑁 + 2
Where TN is the total number of tests of a particular system and NS is the number of successful tests, are implemented to assess reliability instead. The confidence associated with Bayesian analysis increases with higher sample sizes and this has been recognized to be a shortcoming of this approach, particularly for space applications where relatively few identical units are manufactured. Reliability for systems operating in parallel as opposed to in-series are also dealt with differently: Reliability for systems 1, 2 & 3 operating in series: 𝑅!"#$% = 𝑅! 𝑅! 𝑅! . Reliability for systems 1, 2 & 3 operating in parallel: 𝑅!"#$% = 1 − 1 − 𝑅! 1 − 𝑅! 1 − 𝑅! . The appropriate mix of serial and parallel components describing the operation and potential failure modes of a particular system are determined via Fault Tree Analysis. In this manner, the reliability associated with high-level components from each subsystem are assessed as a first-cut for the early design phase decisions. The reliability results are displayed in Fig. 9, where the overall probability of mission success is assessed as roughly 80% (product of all subsystem reliability measures acting in series). Note however, the relative reliability of each individual phase of the mission. Systems with reliabilities highlighted in dark green (system reliability greater than 98%) have generally been assessed at the component level using Rs with historical failure rate data and/or have sufficient redundancy that the Bayesian analysis does not preclude a high success rate. In contrast, systems indicated by light green generally do not have failure rate data readily available and have therefore been assessed by Bayesian testing using mission success/failure events. Furthermore, these systems often require serial operation rather than parallel, and thus represent higher-risk scenarios. These serial/Bayesian systems represent the lowest reliability levels of our design and therefore warrant closer scrutiny at the component level. Additionally, a system failure during the departure stage does not necessarily entail loss of the crew, as abort options to Earth return are still available until hyperbolic velocities are reached. A deeper study incorporating low level component reliability analysis and opinions from subject matter experts is required for a more thorough study.
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Figure 9 Subsystem reliability figures.
The reliability associated with the Launch System assesses the single launch failure as a result of the failure of individual sub-components such as the propellant feed system, avionics, electrical, ADCS, software, structures, engines and thermal system. These values have been acquired from the Exploration Systems Architecture Study (ESAS) undertaken by Flaherty et al.8 The Launch Logistics reliability numbers are calculated based on a go/no-go situation due to delays/interruptions from undesirable weather or logistics delays for all three required launches for the duration of the launch window. The Propulsion Module reliability is determined using the same parameters from the ESAS study referenced for the Launch System. Referencing successful dockings assesses the docking success rate and departures from the ISS, while the ERV and Aerocapture numbers rely on Earth returns from the Dragon capsule as well as the similarly shaped Mercury, Gemini, and Apollo capsules. Habitat and SM reliability is assessed from the successful history of the many pressurized habitats composing the ISS and associated cargo compartment. The reliability of the EPS and TCS systems is based upon the projected 15year lifetime of the individual components9-11 as well as the successfully deployed solar panels and radiators on the ISS. The C&DH, Communications, and ADC systems were designed with a high level of redundancy and therefore afford a high probability of success, even though relatively high failure rates for individual components were assumed.12,13 Finally, the failure rates of the ECLSS components were assessed from predicted values of the ISS ECLSS. Sufficient redundancy on the crew life support is ensured by triple redundancy as well as presence of a storage back-up of fresh water and compressed oxygen. Component failure rates as well as fault-tree analysis for each system are detailed in the Appendix. 2.4.2 Risk Management The number of objectives a mission is able to achieve measures its success. The stakes are even higher when sustaining human life for the mission duration is the most crucial driver. There are bound to be risks associated with such ventures. Therefore, it is imperative to assess and learn from past failures as well as have the foresight to implement risk mitigation and management strategies where possible. As such, we consider its two biggest goals as: a) The ability to preserve human life with minimal psychological effects during the entire mission duration, and b) The ability to perform a successful flyby of the Marian planet. Note that failure of goal b) has serious implications on goal a). With these primary goals at the forefront, the
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subsystem and programmatic risks that have the potential to endanger them are identified and assessed in detail. These risks are quantified as elements of a risk matrix based on their likelihood of occurrence and eventuating consequence(s). This process has been adopted from NASA’s guidelines for risk management.14 Furthermore, steps have been considered to reduce the likelihood and/or adverse consequences, which then leads to a 2nd risk matrix that displays the status of the risk elements ‘after’ applying mitigation strategies. As such, these mitigation strategies have directly influenced many aspects of the design decisions in this study. Note that the lack of expert judgment to assess the likelihood and consequences of risk elements in this competition scenario has been overcome by investigating failure scenarios on previous missions.15 A detailed risk investigation by subsystem, spanning from minute details such as malfunctioning cables/circuits, to cross-wiring, to missed launch windows and desired orbit insertion mishaps, to human state of mind and health attributes and the proposed mitigation strategies are presented in the Appendix. Figs. [10-11] represent the seriousness of all the considered risks both prior to and after the mitigation strategies have been considered. In these figures, the higher the likelihood number, the greater the probability of the event occurring, and the higher the consequence number, the worse the implications of the occurrence of the event. Therefore, even though a risk item might be unlikely to occur, its severity can still rank high if the consequence of it occurring is detrimental to the objectives of the mission. The priority number (severity) of the risk is easily identified by its associated color – low (green), medium (yellow) and high (red) in Figs. [10-11].
Figure 10 Risk Assessment Prior to Implementation of Mitigation Strategies
One of the biggest risks identified in the red cells prior to mitigation is attributed to a higher likelihood and consequence of failure of the aero-capture mechanism during re-entry (7), as this approach has not been tested to warrant a high level of confidence. The other items of concern here are the physical and mental well being of the astronauts on the lengthy and isolated journey through the interplanetary environment (46-47). The medium level risks include aspects such as accommodating scaled-up versions of the Ultra-Flex solar panels (48) and malfunctioning parts such as reaction wheels in the ADCS (8), components in the power system (19, 22) and ECLSS subsystems (40-44), unexpected delays or failures during various mission phases such as launch and fly-by (1, 2, 4, 5), and environment related effects including radiation and impacts from debris and micrometeorites (36-38). The lower level risks in green include items such as the malfunctioning of ADCS thrusters, battery depletion, cable malfunctions and solar cell degradation to name a few. Although some of these items have high consequences, they are deemed to be lower risk items because the likelihood of such events occurring are very low based on historical failure events. Fig. [10] addresses all the identified concerns by suggesting mitigation strategies to lower the priority number associated with a certain risk. Although the ideal situation would be to reduce 9
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both the likelihood and consequence associated with a risk item, this is not always possible. Notice that the high-risk items transform to medium priority risks when, for example, thorough testing is introduced to deal with the aerocapture issue. The other high priority risks associated with the well-being of the astronauts are addressed when amends are made to keep them entertained and active measures are taken to implement routine exercise regimes and monitor their health regularly. Other suggested mitigation strategies include adding redundant systems, as well as providing adequate training to familiarize crew with all the subsystems and how to conduct repairs if required. More details on risk items and suggested mitigation strategies are found in the Appendix.
Figure 11 Risk Assessment after Implementation of Mitigation Strategies
2.5 Mission Schedule In addition to the vehicle design and proposed budget, a defined development and operation schedule must be set. This need is especially critical in order to meet the aggressive timeline required for exploitation of the 2018 window of opportunity. Accordingly, the team utilized GANTTProject software in order to estimate the schedule for this mission. The resulting Gantt chart is shown in Fig. 12.
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Figure 12 Gantt chart for mission.
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.
3 Crew Factors The most critical element of Inspiration Mars mission is the crew, where the goal is to demonstrate that humans can survive, even thrive, on long-duration, interplanetary voyages. As such, much provision must be made for the crew’s physical and mental health. 3.1 Inputs from Professionals We were fortunate to have the opportunity to interact with professionals involved in human spaceflight programs and discuss their experiences with spaceflight. The suggestions received from them are shown in Table 2. These comments further motivate the investigation of the psychological well being of the crewmembers. Table 2 Suggestions by professionals on this mission.
Name
Profession
Suggestion
Eugene Cernan Astronaut
Astronauts are human beings and they want to feel at home even when they are inside spacecraft. Small things like daily use products being out of place make them feel stressed out.
Loren Shriver
Astronauts can adjust to lower habitable volumes. The only feature they can’t compromise with is the window of the spacecraft.
Astronaut
Charles Walker Astronaut
Bring along plants and living organisms. Studying the effects of radiation and microgravity will be crucial for future interplanetary missions. Plus, it will give the crew something interesting to do.
3.2 Kanau Spacecraft’s Interior Design The environment within the vehicle must as best as possible satisfy the human need for natural rhythms. Furthermore, the setup of interior spaces should provide areas for privacy and convey subtle reassurance of a natural environment to the crew. 3.2.1 Baseline Architecture We propose an interior design incorporating six components as shown in Fig. 13.16,17
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Figure 13 Aspects of interior design.
3.2.2 Interior Design of Spacecraft All potential dangers should be removed to ensure: (1) Safety during the mission, (2) Visibility when crew participates in some activities, (3) Flexibility to adapt to various situations, (4) Variability because monotony of visual stimuli leads to strong discomfort, which is certified by past space missions like Skylab Space Station 4,18 and (5) Intuitive and friendly design. Sleeping Bag Customization of the spacecraft is needed because interviews of astronauts suggest privacy is an important factor for crew.19 The crewmembers of Inspiration Mars mission will be a longtime married couple and hence a sleeping bag for the two people should suffice. Physical intimacy is one important factor to maintain the healthy relationship of a married couple. Therefore, such a sleeping bag and placement of any cameras will be important considerations. Lighting Changes and Visual Distance Visual stimuli produced by interior light have the potential to affect crewmembers’ psychological condition. For example, a new lighting system that produces artificial rainbow lights can be created by slowly rotating a transparent ring and a spot light as shown in Fig. 14.17 Furthermore, indirect lighting supplied by LEDs recreates natural color variations during the course of an entire day mimicking biological sunlight effects.18 Additionally, strategic placement of mirrors or images projected on large screens within the habitat can provide a greater sense of physical space.
Figure 14 Artificial rainbow generated by spotlight. "Round Rainbow" by Olafur Eliasson, Hirshhorn Museum, USA. Picure copyright: Ayako Ono 2008.
Sound Environment and Feeling Earthly stimuli are also very important since humans that are confined in a small spacecraft for long durations are placed under great psychological stress. One effective method to avoid mental disorders is to lead life as if it is being lived on Earth. To mimic Earth’s
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environment, soundscape design was considered for the spacecraft’s interior. For example, sounds of nature such as sounds of a stream and a bird’s twittering can be used as a sound therapy. Such sounds of nature are helpful for relaxation.20 Plant Cultivation Box To recall nature on Earth, the spacecraft can be equipped with plant cultivation boxes. These plant boxes invoke positive stimulation. After enjoying the cultivation of the plants, crewmembers can eat them for additional variety in their diet. However, cultivating plants in a spacecraft is a difficult task and therefore imitations of wood, grain or artificial flowers may suffice for the mental relaxation of crewmembers.17 Spacewear Astronauts used to wear one pair of cloth for three consecutive days.21 We advise the use of advanced spacewear like the undergarment tested by a JAXA astronaut on the ISS that can be worn for one month at a stretch.22 The exercise cloth recently developed and under test at the ISS can provide pleasant odor environment in the spacecraft and can be worn for longer periods of time.23 We propose the use of advanced spacewear to resolve the issue of either carrying a large amount of clothes or carrying a washing machine on the spacecraft, which at the present does not exist.21 Renderings of Interior Design While the selected crew must provide input for customization of the interior arrangement, some conceptual images of the interior spaces may inspire the investigation of this aspect of the design problem, even during preliminary planning stages. The following pictures show artist renderings and potential schematic views of the interior design. Note, however, that the crew must retain control over their environment during the course of the interplanetary cruise and therefore interior spaces must be reconfigurable.
Figure 15 Inside the shelter.
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Figure 16 Interior design of the habitat.
3.3 Facilities for Crew Physical Health In addition to the emotional and mental well being of the crew, steps must be taken to maintain their physical health. This system needs to provide measures to meet crew bone, muscle, sensory-motor, and cardiovascular standards defined in NASA-STD-3001, Volume 1. Measures shall maintain in-flight skeletal muscle strength at or above 80 percent of baseline values and bone mass consistent with requirements for a safe return to Earth’s gravity.24 The exercise and muscle data gathered from nine crewmembers while on the ISS for 6 months clearly support the notion that changes to the exercise prescription are necessary to protect skeletal muscle for long-duration space missions.25 Facilities for crew physical health are composed of three components as shown in Fig. 17, including the Cycle Ergometer with Vibration Isolation and Stabilization (CEVIS) system, the Treadmill with Vibration Isolation System (TVIS), and the advanced Resistive Exercise Device (aRED).
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Figure 17 Outline of facilities for crew physical health.
Aerobic and Resistive Exercise Devices26 U.S. crewmembers are required to complete a 2.5-hour bout of combined aerobic and resistance exercise on 6 of 7 days during the mission. Onboard the ISS, approximately 1.5 hours were devoted to resistive exercise on the advanced Resistive Exercise Device (aRED) and 1 hour was devoted to either the Treadmill with Vibration Isolation System (TVIS) or the Cycle Ergometer with Vibration Isolation System (CEVIS) or a combination of the two. We suggest implementation of exercise devices such as those in use on the ISS, whilst also being open to new developments in astronaut exercise regimes. One of new developments is hybrid training. Medicine Medicines can potentially be used to prevent the loss of bone mass and formation of kidney stones during long stays in space with a bisphosphonate formula used in the treatment of osteoporosis. Such medicines have been experimentally used by the ISS crewmembers and can be explored for this Mars mission.27 3.4 Facilities for Radiation Protection The interplanetary cruise will expose the crew to radiation not normally experienced on the ISS, due to the lack of shielding from Earth’s magnetosphere and it should be managed through system design and application of appropriate countermeasures.28 Furthermore, particular events, such as solar flares29 or passages through Earth’s radiation belts30, expose the crew to extended period of high radiation doses. Accordingly, the expected radiation dose is calculated to determine if additional shielding will be required. 3.4.1 Sources of Radiation Exposure The crew will receive radiation doses from three sources during the mission. 1) Trapped solar protons in the Van Allen Radiation belts around Earth, 2) MeV Solar Protons from the sun, 3) GeV Galactic cosmic rays from extragalactic sources. In addition, the radiation dose received from any past spaceflight experience must be accounted for. 3.4.2 Radiation Dose Modelling The crew was assumed to spend the majority of their time in the MPLM, which was assumed to have a pressure hull with an areal density of aluminium of 5 g/cm2. Radiation exposure was modelled via NASA’s On-Line Tool for the Assessment of Radiation In Space (OLTARIS). Trajectories were generated for the TMI and EDL phases of the mission and uploaded to OLTARIS. Interplanetary flight was modelled as free space flight at 1 AU during the previous solar minimum (mimicking the environment the spacecraft will experience during the solar minimum in progress in 2018). Even though the mission is expected to take place during a period of minimal solar activity, a direct hit by a large Solar flare (modelled after the event of October, 1989) was assumed to hit the spacecraft while in free space, with the crew taking
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shelter in the Dragon, placing both the ATV and MPLM between them and the storm, providing an average of 64 g/cm2 of shielding from the entire spacecraft. Effective dose equivalents were then simulated for the averaged value of a computerized anatomical male and female, as well as for a mass of silicon (to represent critical flight electronics). The equivalent dose absorbed by the astronauts and the true dose absorbed by the silicon from all three sources is summarized in Table 3. Table 3 Radiation dosages for Kanau mission.
Source/Phase GCR/TMI GCR/Transit GCR/EDL Trapped Proton/TMI Trapped Proton/EDL SPE/Transit Total
Astronaut Flight Electronics (mSv) (mGy) 1.39 .37 516.10 143.40 .78 .21 31.90 55.10 7.60 13.20 144.00 188.60 702 mSv 401 mGy
These numbers represent the total dose received in flight. Assuming the astronauts had previously spent 1 year on the ISS before being selected for this mission will add an additional ~300 mSv to their total accumulated dose, bringing a conservative estimate to the total dose received by the astronauts at the end of the mission to ~1 Sv. This is well under the maximum limit imposed by NASA standards on astronauts over the age of 45. The number, arrived at even after several extremely conservative assumptions (no shielding provided by spacecraft interior objects, anomalous solar activity, extended pre-flight ISS crew rotation etc.) implies that additional radiation shielding is not required beyond the inherent shielding afforded by the spacecraft’s structure. 3.5 Crew Selection from American Astronauts Since the crewmembers will be a married couple, we examined the list of all possible candidates for crewmembers from among the current American astronauts and then selected the best option for this mission: Shannon Walker and Andrew Thomas. While couples with longer marriages will likely have a higher level of bonding, the statement made by Karen Nyberg indicates that astronaut couples that have children will likely refrain from joining this mission.32 More details on this selection process is contained in Appendix. 3.6 Crew Daily Schedule The daily schedule of crewmembers was developed based on current ISS schedule pattern. Specific requirements of this mission have been incorporated as shown in Table 4. CM-1 denotes female crewmember and CM-2 represents male crewmember. NASA’s Onboard short term plan viewer (OSTPV) can be utilized to design timeline for each day.33 Note that this schedule is flexible.
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Table 4 Nominal daily schedule aboard Kanau spacecraft.
Time
CM-1 Activity
CM-2 Activity
6:00 AM 6:45 AM 7:30 AM 7:55 AM 8:15 AM 9 AM 11 AM 1 PM 2 PM 2:30 PM 4:15 PM 5:15 PM 6:15 PM 7:15 PM 7:30 PM 8:30 PM 9:30 PM
Wake up, clean up. Have breakfast. Personal/Group video log. Medical inspection. Daily housekeeping. Exercise. Maintenance. Have lunch. Private time with CM-2. Research, experimentation. Hobby. Daily housekeeping. Video conference/medical inspection. Private time with CM-2. Have dinner. Entertainment/relaxation. Sleep.
Wake up, clean up. Have breakfast. Personal/Group video log. Medical inspection. Daily housekeeping. Maintenance. Exercise. Have lunch. Private time with CM-1. Research, experimentation. Daily housekeeping. Hobby. Video conference/medical inspection. Private time with CM-1. Have dinner. Entertainment/relaxation. Sleep.
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4 Detailed Mission and Vehicle Description A vehicle structure and suite of component systems sufficient to satisfy the requirements for a flyby of Mars and ensure the safe return of the two-person is a critical component of our mission architecture. The primary goal of the current investigation is the detailed design of a feasible crewed interplanetary vehicle where, due to the short timeline needed to reach the 2018 window, currently existing hardware with short modification and integration times are favoured over more advanced technologies that still require significant development. The factors and constraints driving the overall design are identified and prospective solutions are found. The team developed a morphological chart to list down all possible alternatives for all subsystems required for this mission. 880602513408 solutions were generated but the team was able to narrow down its search to 32 design alternatives. Results of the House of Quality obtained earlier while performing Quality Functional Deployment were used in decision-making regarding these alternatives. The morphological chart is shown in Fig. 18.
Figure 18 Morphological chart of Kanau mission.
4.1 Trajectory Analysis and Launch Vehicle Selection As a baseline mission design, we recreate the free-return trajectories identified by Inspiration Mars (IM) in their mission concept documents. We also seek to explore novel 19
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mission architectures that could potentially reduce the needed number of launches, increase operational safety margins, and further enhance the baseline scenario. The following sections contain our recreation of the IM baseline as well as an assessment of launch and Trans-Mars Injection (TMI) requirements to satisfy the mass/budget specifications of our design. 4.1.1 Interplanetary Ballistic Free-Return Trajectory The potential for a fast free-return flyby of Mars by a spacecraft was first identified by Patel34 in 1998; the original Inspiration Mars concept uses these free-returns to enable what could be the first manned mission to visit Mars.35,36 For our design concept, we select as our baseline trajectory as January 2018 departure free-return with the flyby occurring in August 2018. After a chemical boost stage from LEO to escape the vicinity of Earth, the spacecraft coasts to the Mars close approach, uses Mars for a gravity assist maneuver, and returns to Earth in May 2019 to complete a 501-day journey. The magnitude of the TMI maneuver is 4.86 km/s to depart from the staging LEO altitude of 200 km, where the propulsion system and propellant required for implementing this maneuver must be delivered to LEO in addition to the crewed spacecraft. The critical epochs of the free-return trajectory, constructed using Lambert arcs with the Sun as a point mass and the positions and velocities of Earth and Mars from the Jet Propulsion Laboratory’s (JPL’s) HORIZONS system, are contained in Table 5 while the calculated hyperbolic velocities are reflected in Table 6. Table 5 Critical epochs of the Earth-Mars-Earth free-return trajectory. Leg
DEPART
1
Earth
2
Mars
2018 January 04 07:10:33.6 UT 2018 August 20 07:49:43.7 UT
Mars Earth
ARRIVE 2018 August 20 07:49:43.7 UT 2019 May 20 20:57:41.8 UT Total Duration
Flight Time (Days) 228.0 273.5 501.5
Table 6 Velocities and periapsis altitudes of the departure, fly-by, and return hyperbolic trajectories. Leg 1 2
Vinf (km/s) 6.230 5.389
DEPART Alt Peri V Peri (km) (km/s) 200 km 12.649 239.2 6.501
C3 (km2/s2) 38.811 29.038
Vinf (km/s) 5.389 8.865
ARRIVE Alt Peri V Peri (km) (km/s) 239.2 6.501 -14.201
C3 (km2/s2) 29.038 78.596
The interplanetary cruise, illustrated in Fig. 19, is a relatively well-known design that requires only minimal course corrections to address statistical maneuvers and navigation errors. The outbound arc to Mars is shown in black while the return leg is dashed; the paths of Earth and Mars are shown in blue and red, respectively. The close approach to Mars is illustrated in Fig. 20, where the point of closest approach at Mars is 239.2 km above its surface. These trajectory arcs are computed using a patched conic model; however the Inspiration Mars reports detail similar results generated in a full ephemeris force model. Contingency Interplanetary Trajectory – Solar Electric Propulsion The baseline free-return trajectory requires no deterministic maneuvers after the TMI at Earth departure. Accordingly, the vehicle stack does not possess any propulsive capability apart from that provided by the Attitude Control System (ACS). However, Smet et al. have demonstrated 20
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that the inclusion of a Solar Electric Propulsion (SEP) module affords increased mission capability for a potential Mars flyby mission. Flight times, departure mass from Earth, and Earth re-entry velocities can all be reduced by the strategic use of a SEP system. The most significant gains are associated with mission windows after the nominal 2018 departure. For example, a SEP-enabled mission departing Earth in 2021 and requiring deep space maneuvers (DSMs) highlights a decreased launch mass from Earth as compared to the required mass for a chemical propulsion system. We performed sufficient analysis to confirm specific point solutions associated with the 2018 free return; however, the gains afforded do not outweigh the increased cost and complexity required. Thus, while we do not include SEP in our baseline architecture, we do note the potential contingency for future mission planners.
0 AU
1 AU
Figure 19 Ballistic free-return trajectory with Earth departure in January 2018, Mars flyby in August 2018, and Earth return in May 2019. Color bar indicates spacecraft distance from Earth.
Figure 20 Mars close approach,.
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4.1.2 Earth Launch and Departure While the free-return trajectory enables a low-cost transfer option from Earth to Mars and back again, a large excess velocity is required at Earth departure. Thus, for a vehicle of sufficient size for two crewmembers, a large amount of propellant is required to ensure TMI. However, as no Earth launch vehicle is of sufficient size for a direct TMI injection, the crewed sections and an appropriately sized propulsion module (PM) must be assembled in LEO. This assembly process requires a series of launches from the surface of the Earth as well as docking maneuvers in the staging LEO. These launch considerations are discussed in detail. Launch Site The selection of an appropriate launch site has significant implications for the achievable insertion orbits as well as the launch schedule. For example, the latitude of a launch site determines the minimum inclination that is achievable for the insertion orbit. In Table 7 are the launch sites we have considered in our investigation. Of the four launch sites; Cape Canaveral has the closest latitude (28.5°) to the axial tilt of the Earth with respect to the solar system ecliptic (23.6°). Thus, with proper launch timing, the launch payload can be inserted into an orbit appropriate for Earth departure onto a Mars-bound trajectory. While Cape Canaveral provides needed access to our desired staging and departure orbits, inclement weather is not an infrequent local occurrence, necessitating the scheduling of sufficiently wide launch windows to mitigate potential weather disruptions. However, in a best-case scenario, wide launch windows can provide additional on-orbit time for system checks prior to the desired TMI of January 4, 2018. Table 7 Potential launch sites.
Launch Site
Country
Latitude (deg.)
Cape Canaveral Vandenberg Kourou Baikonur
USA USA French Guiana Kazakhstan
28.5 N 34.8 N 5.2 N 46.0 N
Primary TMI Propulsion Module Architecture – DCSS Due to the short timeline for mission development before the 2018 opportunity, booster hardware with a proven flight heritage must be identified for incorporation within the propulsion module (PM). Therefore, an Earth-departure module enabled by the LOX/LH2 propellant is selected as our primary architecture, where this propellant combination is relatively well understood and offers a significant mass advantage. One readily available LOX/LH2 propulsion system with a history of successful performance is the Delta Cryogenic Second Stage (DCSS) from the United Launch Alliance (ULA). The known hardware configuration of the DCSS, in particular the pre-specified wet mass of 30.7 metric tons (mT) and 0.87 mass fraction, necessitates that the TMI maneuver is split into a series of three boosts from three DCSS modules. Thus, the PM is composed of a series of three DCSS 5-meter stages, with associated payload adapters, as illustrated in Fig. 21. The three boosters fired in sequence then are detached as their propellant is exhausted, with each maneuver delivering successively higher ΔV to the vehicle stack, as displayed in Table 8. The excess velocity is reserved for any needed cleanup maneuvers after the main TMI burn.
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Table 8 Propulsion module performance and departure staging orbits.
Stage
ΔV delivered, km/s
New orbital period, hours
1st Boost
1.064
2.48
2 Boost
1.469
12.29
TMI
2.381
Escape
Excess
0.049
--
nd
Figure 21 Kanau vehicle stack with exploded view of propulsion module. Grey cylinders represent adapters between the booster stages and the vehicle.
Splitting the TMI burn into three separate maneuvers necessitates the use of a series of progressively larger and more elliptical staging orbits about the Earth prior to TMI. One maneuver, corresponding to the use of one DCSS module, is performed per revolution, where this burn is timed to occur at perigee such that the maximum available thrust is directed into the escape from Earth. The LEO staging orbit, departure staging ellipses, and the Earth escape are sketched in Fig. 22, with the corresponding periods of the post-maneuver ellipses provided in Table 8. In order to achieve the Earth departure epoch of early morning January 4, the staging process must start approximately 18 hours prior, that is, shortly after noon on January 3. The crew experiences a maximum acceleration of approximately 3g’s, well within limits for healthy adults.
Figure 22 Schematic of the Earth departure stages.
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Contingency TMI Propulsion Module Architecture – ACES The DCSS has a proven flight heritage and many successful operations. On the other hand, the limited capability of the DCSS entails the need for more complex operations (e.g., number of burns, dockings, stage separations, etc.). Furthermore the cryogenic propellants of the stage boil-off during long loiter times. To address these concerns, ULA is currently developing the Advanced Cryogenic Evolved Stage (ACES) booster for employment as a high-capability module for upper stages of launch vehicles as well as long-term, in-space propulsion. ACES is being evolved using heritage designs from the DCSS and Centaur upper stages, where the new stage offers improved flexibility in staging options as well as reduced propellant boil-off. As demonstrated in the Appendix, this prospective PM reduces the number of required boost maneuvers and the required mass delivered to LEO. Furthermore, whereas the LOX/LH2 in a DCSS would rapidly boil-off during the interplanetary cruise, the novel active and passive thermal control elements of ACES could potentially retain any excess propellant post TMI. This remaining ΔV capability could enable further station keeping as well as potential deterministic maneuvers during the interplanetary transit. However, given the current development status of this PM, we do not expect the ACES system to be flight proven by the 2018 opportunity. Therefore, we simply note the potential option, perhaps available during a 2021 window, while retaining DCSS as the primary configuration for our mission. Launch Architecture and Timeline Of the launch scenarios considered, we select as our primary option the Falcon Heavy launch to the staging LEO orbit. We choose this scenario based upon expected launch cost as well as estimated technology readiness level. The Falcon Heavy is expected to first launch in 2015, and the remaining years leading up to the 2018 mission opportunity can be used to finalize technical and organizational details. While the Falcon Heavy will bring 53 mT to LEO, the mass of our vehicle stack (including the PM) is 126 mT, requiring a total of 3 launches, however only the final launch carrying the crew must occur within a short time frame before the nominal TMI. On the other hand, the cryogenic nature of the LOX / LH2 propellant that the DCSS stage uses means that boil-off is a concern while the propellant is in the staging orbit.37 Therefore, even though on-orbit cryogenic storage may be used for the fuel and oxidizer, launches of the storage tanks should not occur too early before the launch of the crewed capsule. We therefore propose the launch timeline in Table 9, where scheduling a two-week window between each propellant launch also accommodates for weather events at the launch site that could unduly affect the mission timeline. A shorter launch window is used for the crew in order to reduce the crew time in the staging orbit prior to TMI. The last launch opportunity for the crewed launch is constrained to terminate one day prior to the nominal beginning of TMI operations so that onorbit systems checks and docking maneuvers may be performed prior to TMI. After the second launch, the previously launched DCSS docks at the end of the DCSS + SM stack. After the third launch, the DCSS splits from its launch configuration with the habitat moved aft to the end of the vehicle stack whereas the ERV undocks and performs a 180° turn to re-dock at the other end of the habitat. Thus, a total of two separations plus three rendezvous / dockings processes are required for on-orbit assembly of the vehicle stack.
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Table 9 Launch timeline and manifest.
Launch 1 2 3
Payload to LEO DCSS + SM DCSS + Habitat + PMA DCSS + ERV + Crew
Margin (mT) 4.87 8.39 17.10
Window Nov. 25 - Dec. 9, 2017 Dec. 9 – Dec. 23, 2017 Dec. 23, 2017 – Jan. 2, 2018
4.1.3 Earth Return and Aerocapture Upon return to the Earth, the initial re-entry velocity is 14.2 km/s, an entry velocity most likely beyond the capabilities of current heat shield technology. Therefore, the crewed ERV will perform an aerocapture maneuver after separation from the main vehicle stack. Our preliminary investigation of the aerocapture maneuver indicates that, by exploiting the lift of the ERV (L/D=0.18 for the Dragon capsule)38 and optimizing the flight profile through the atmosphere, an effective ΔV of 3.7 km/s can be achieved. This change in velocity places the ERV on a highly elliptical orbit that returns the crewed capsule to the Earth after 15.9 hours; the tracks of the return hyperbola and capture ellipse are illustrated in Fig. 23.
Figure 23 Earth return hyperbola and aerocapture ellipse.
Aerocapture using Lift Modulation Technique Lifting aerocapture involves the re-entry capsule flying at a higher altitude over longer timeline than a comparable drag-enabled aerocapture. Stagnation heat rate and g-loads are minimized as the vehicle stays away from denser atmosphere that is primarily responsible for these two issues. An optimization problem was formulated with an objective to maximize the time of flight of the re-entry capsule with the conditions shown in Table 10. Table 10 Boundary value conditions for lifting re-entry.
Parameter Radial magnitude (km) Downrange (km) Velocity (km/s) Flight path angle (deg) Bank angle (deg)
Entry Value 120 0 14.2 -5.4511 180
25
Exit Value 120 1300 10.5 4.1213 180
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Both convective39 and radiative40 heating effects were taken into consideration and a total heat rate on the ERV was computed. The Thermal Protection System (TPS) on the SpaceX Dragon is made using PICA-X, a material derived from NASA’s PICA (Phenolic Impregnated Carbon Ablator) material. It is capable of handling heat-rates of up to 2000 W/cm2.41 The capsule was allowed to fly either with full lift down (bank angle is 180 degrees) or full lift up (bank angle is 0 degrees). Comparison was made between ballistic and lifting re-entries as shown in Table 11. Optimization was carried out using MATLAB’s boundary value problem solver, bvp4c, which confirms that the vehicle flies with full lift down throughout (bank angle of 180 degrees) to counteract the centrifugal force on it due to the orbital motion and spends more time at a higher altitude. The results are shown in Fig. 24. Human tolerance to g-loading varies depending on the level of training and the duration of exposure.42 Persistent g-loads above 5g’s can cause loss of consciousness. Astronauts weakened after a long-term zero-G cruise will be even more susceptible to these effects. Table 11 Comparison between different aero assist re-entries of ERV.
Parameter Perigee altitude (km) Peak stagnation heat rate (W/cm2) Peak g-load (Earth gs) Time of flight (sec) Control Total stagnation heat load (kJ/cm2)
Ballistic 69.6455 2642.4947 7.1123 79.1273 No control 125.3990
Lifting 72.5703 1985.3517 5.2331 100.8044 Full lift down 126.8961
The benefit of having a higher L/D ratio of the re-entry capsule was also studied and its results can be found in the Appendix.
Figure 24 Various plots for lifting and ballistic re-entries.
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4.2 Attitude Determination and Control System (ADCS) The Attitude Determination and Control System (ADCS) assure the proper orientation of the vehicle stack during the interplanetary cruise and during mission critical events. Attitude determination is supplied at all times by a suite of Star Trackers, Sun Sensors, and Inertial Measurement Units (IMUs). Attitude control prior to the TMI maneuver is assumed to be provided by the PM. After TMI, the spacecraft remains oriented with the Sun astern for the duration of the interplanetary cruise. Furthermore, during the Mars close approach, larger orientation alterations are required to enable the crew to view the surface of Mars. Finally, upon Earth return, the proper attitude for ERV separation and re-entry must be ensured. Fortunately, there are many commercially available sensors and control units with strong flight heritages, in addition to built-in attitude control capabilities for our selected crew modules. Accordingly, the components of the vehicle ADCS are detailed in Table 12, with specified vendors, hardware model and redundancy. Table 12 Specifications for attitude control system.
System Sensor Suite Star tracker Sun sensor IMU Attitude Control Large AC Thrusters Small AC Thrusters Reaction Wheels
Model
Vendor
Redundancy
FSC-701 2-axis fine sun sensor LN-200S
Ball Aerospace Adcole Northrup-Grumman
3 12 3
Draco Low Power Resistojet MWI 100-100/100
SpaceX Surrey Space Systems Rockwell Collins
16 28 4
The sensor suite is developed based upon the attitude determination system of the Mars Reconnaissance Orbiter (2 star trackers, 16 sun sensors, and 2 IMUs). Additional redundancy is required for the star tracker and IMU, while; on the other hand, we have reduced number of sun sensors due to the nominal sun-pointing attitude of our spacecraft. Four sun sensors are placed on the aft, sun-facing side of the SM while two units are placed on each of the four remaining sides of the SM. This configuration ensures that the spacecraft orientation with respect to the sun can be determined for both nominal and contingency operations. However, the three star trackers are oriented such that they point largely anti-sunward, that is, from the habitat to the ERV while still retaining sufficient field of view of the stars. The IMUs are placed on the SM close to the connection between the SM and the habitat, close to the vehicle center of inertia. The attitude control suite is composed of the Draco thrusters provided on the Dragon re-entry capsule as well as reaction wheels and small resistojets placed in the modified ATV service module. The Draco thrusters, possessing a baseline reserve of 1000 kg of NTO/MMH hypergolic propellant, are reserved for contingencies and mission critical events, i.e., Mars flyby and Earth re-entry. On the other hand, the reaction wheels serve to reject disturbance torques during interplanetary flight and ensure sun-pointing. Indeed, the reaction wheels provide an estimated turn rate of 1.48e-4 rad/sec when placed near the vehicle center of inertia (that is, roughly between the SM and habitat), well above the maximum turn rate of 3.5e-7 rad/sec needed to ensure the SM remains sun-facing. The resistojets, using waste water from the crew support system, are used to periodically de-saturate the momentum from the reaction wheels as well as for faster turn rates when use of the Draco thrusters is not desired.
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4.3 Environment Control and Life Support System (ECLSS) The environmental control and life support system (ECLSS) is critical to ensuring the safe return of the crew to Earth. As such, a highly reliable system with high redundancy is required. However, the ECLSS must be relatively low mass and not take up excessive volume in the service module and habitat. 4.3.1 Essential Subsystems of ECLSS and System Parameters We conducted the feasibility study for designing the Environment Control and Life Support System (ECLSS) with data from currently available subsystems. Since super high reliability for safety of manned mission is required, technologies TRL with more than 8 are crucial for building ECLSS. Within various fields of ECLSS elements, four essential subsystems: Atmosphere Management, Water Management, Waste Management, and Food Supply should be of utmost concern because these directly influence daily human life. Each subsystem has the functions as shown in Fig. 25. We classify the key functions, which must be handled for the Mars manned mission as red boxes. Tables [13-14] show the consumption and production parameters from human metabolism and living related to the essential functions.
Figure 25 Functions for each ECLSS management subsystem.
Table 13 Material consumption of human crew (kg/Crew Member-day). Minimum
Nominal
Maximum
Comments
Water Drinking Food supply Hygiene (oral, hand, face) Shower Laundry and Urinal flush Food Food Packaging Air Oxygen
2.0043 0.50 4.4543 2.7243 0.0
uses gray water
0.5444 0.0844
0.61744 0.0944
0.6644 0.1044
0.38544
0.83544
1.85244
28
15% of food
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Table 14 Material production of human crew (kg/CM-day). Minimum
Nominal
Maximum
Comments
Water Urine water Fecal water Respiration water Perspiration water Gray water Solid Waste (dry basis) Fecal waste Perspiration waste Air Carbon dioxide
0.80344 0.03644
1.88644 0.09144 0.88544 0.69944 7.17
0.97544 1.97344 sum of utilized water
0.03244 0.01844 0.46644
0.99844
2.24144
4.3.2 Selection of Subsystems and Reliability The most reliable ECLSS subsystems currently available are technologies used in the ISS. Onboard the ISS, carbon dioxide exhaled from the crew is collected by the Carbon Dioxide Removal Assembly (CDRA) and resolved into water using a Sabatier system (Eq.1), and oxygen is generated by electrolysis (Eq. 2).45 2H2O → 2H2 + O2 CO2 + 4H2 → CH4 + 2H2O
(1) (2)
In order to produce 0.835 kg of oxygen per person per day, theoretically, 1.88 kg of water is needed. If the hydrogen required to be combined with carbon dioxide in a Sabatier reaction is supplied from water electrolysis, only 1.15 kg of carbon dioxide can be reduced, while two people create about 2.0 kg of carbon dioxide per day. The rest of the carbon dioxide must be dumped out. Another concern is that water electrolysis consumes about 0.94 kg of water per day over the amount of water produced by the Sabatier reaction. Thus this closed loop system requires 471 kg of water to be carried in the spacecraft to maintain oxygen for 501 days. On the other hand, NASA recycles urine and drain water by Urine Processing Assembly (UPA) and Water Processing Assembly (WPA) to create fresh water. According to the parameter in Tables [1314], a human consumes 2.5 kg of water for drinking and food, and turns it to 3.561 kg of wastewater as urine, fecal water, respiration water and perspiration water since a human ingests some water from the food itself. Therefore, while the UPA has only an efficiency of 85% at recycling water, an additional 0.8 kg of water is produced per day. 4.3.3 Food Management Growing many crops or vegetables would be difficult within a small habitable area. Also, it would be an inefficient use of space for a crew size of two for the mission duration, so food supplies should be precooked foods that it is ready to eat, requiring only reheating or rehydration. However, we suggest cultivation of some vegetables like lettuce that are easy to cultivate and require a small amount of space. This will not only provide a limited amount of
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fresh food but also assist in the psychological health of the crewmembers. Many types of home cultivation kits are available as shown in Fig. 26.
Figure 26 Lettuce home cultivation kit46.
4.3.4 Simulation of ECLSS using our own simulator, SICLE We have developed a new program for ECLSS simulation, named the SImulator for Closed Life and Ecology (SICLE).47 Two main advantages of SICLE are its user-friendliness and its ability to apply new models and functionalities. Users can easily design and follow their own system designs graphically by utilizing SICLE’s GUI as shown in Fig. 27. Furthermore, it is able to analyze both closed and open loop systems. SICLE would be made available to the public soon. We conducted our ECLSS simulation using SICLE to find out problems and to overview material transfer. In the simulation, crewmembers eat three meals and perform 2 hours of exercise in the morning every day. Fig. 28 is the first one-month simulation result of the quantity of each material in tankage of recycling systems. From the result of the simulation, we estimate the necessary size of tankage for urine, grey water, oxygen, and carbon dioxide within the recycling system. In each tankage, the maximum amount reached is 1.95 kg, 5.55 kg, 0.54 kg and 0.64 kg respectively.
Figure 27 Material flow through ECLSS SImulator for Closed Life, SICLE.
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Figure 28 Material amount in tankages by SICLE simulation.
4.4 Command & Data Handling Subsystem48 (C&DH) The Command and Data Handling (C&DH) subsystem, essentially the brain of the spacecraft, comprising of space flight computer, flight software and solid state recorder, performs the following: ● ● ● ● ● ● ● ● ● ● ●
Manages all forms of data on the spacecraft Carries out commands sent from Earth Prepares data for transmission to Earth Manages collection of solar power and charging of the batteries Collects and processes information about all subsystems and payloads Keeps and distributes the spacecraft time Carries out commanded maneuvers Autonomously monitors and responds to a range of onboard problems that might occur.
4.4.1 Requirements48 Perform all functions requested of a command and control module in a complex spacecraft. Provide continuous audio and occasionally video link with the ground station. Provide internet-style connection.
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4.4.2 Baseline Design48 Several distributed units called Control and Data Management Units (CDMU) are used to implement DCDMS (Distributed Control and Data Management Systems). The baseline concept consists of modular functions listed in Table 15. Table 15 Modular functions and their configuration for DCDMS.
Modular function Processor module Telemetry transfer frame generator Reconfiguration module Distributed memory module
Configuration Includes digital interfaces. Directly interfaced with transponders. Two modules always powered, one of which is a master clock and the other acts as a backup processor or spare. Contains VRAM and NVRAM modules. NVRAM acts as a safeguard memory.
● The crewmembers should have available mass memory of the order of several terabytes, including dedicated memory for personal use. We suggest use of RAD5500 PowerPC, a new processor that is 10 times faster than RAD750 processor.49 4.4.3 Budgets48 We estimate the need for approximately 450 terabytes of data including a 50% margin for scientific and housekeeping data. The expected C&DH mass and power budget as a result have been evaluated as shown in Table 16. Highly reliable triple modular redundant avionics systems already existing on Dragon and ATV are used.50,51 Table 16 C&DH budgets.
Property Solid State Recorder52
Units 3
Unit Mass (kg) Unit Volume (m3) Unit Power (W) 25 0.03 0.1
4.5 Communications (COMMS) The overall communications must support the full scope of Mars flyby mission, including launch, Earth orbital operations, trans-Mars injection (TMI), Earth-Mars cruise, Earth return, and Earth arrival. Meeting these mission phases would require the combined capabilities of the Space Network for initial near-Earth support, the Deep Space Network (DSN), and dedicated Mars network assets as shown in Table 17. Table 17 Communication for a Mars flyby mission. Mission phase Launch through TMI
Network NASA Space Network
Earth-Mars-Earth cruise
NASA DSN
Services Tracking and Data Relay Satellite System (TDRSS)
Bands utilized S-band and Ka-band X-band for basic telemetry link and Ka-band for highrate link
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4.5.1 Band and Frequency Design48 NASA has made the use of Ka-band compulsory for all deep space missions beyond 2016.53 Ka-band data rates are higher for uplink than for downlink, mainly because of the higher transmitted power by the ground station (G/S). For contingencies, X-band has been used because it has less weather dependence than Ka-band and hence higher availability. For the TV-relay satellite link, X-band has been chosen, since the pointing requirement is lower than Ka-band and it has a high enough data rate. The bands and frequencies used are consistent with the Space Frequency Coordination Group (SFCG).54 Table 18 summarizes this design. Table 18 Communication antennas in a nutshell. Kind of antenna Dish antenna MGA Patch Dish antenna
Quantity
Band
1 2 1
Ka- Band X- band X- band
Gain (dBi) 59.1 18 30
Size 3m 8 cm 45 cm
Data rate Uplink 1.8 Mbps 22 Kbps 30 Mbps
Data rate Downlink 1.5 Mbps 460 bps 30 Mbps
Steering mechanism (hemispherical) 180° 180° 180°
The communications subsystem of the spacecraft is summarized as shown in Fig. 29.
Figure 29 Kanau spacecraft communication layout.
4.6 Electrical Power Systems (EPS) The electrical power system (EPS) supplies the required electricity for all other vehicle subsystems and is, therefore, critical for both the success of the Mars flyby as well as the survival of the crew. Accordingly, we develop an electrical power generation, storage, and distribution system capable of satisfying all mission objectives.
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4.6.1 Source Selection Currently available technology capabilities for long duration spaceflight currently present only two options for power supplies: photovoltaic and nuclear. While nuclear electric systems offer several advantages in terms of power density and reliability, no nuclear system has yet been developed for extended human spaceflight. In addition, photovoltaics have demonstrated flight heritage on a multitude of manned and unmanned missions, including the ISS and ATV, providing lifetime and power levels in excess of those required for the IM mission.55 Based on these considerations, and given that the abbreviated schedule allows for no time to develop novel nuclear electrical systems suitable for spaceflight, we determine that photovoltaics are the superior choice for powering the IM spacecraft systems. 4.6.2 Power System Design Factors The requirements to support a crew for mission duration >1 year suggests a useful benchmark of comparison for designing the IM vehicle power systems is the ISS. Design of the power system shall thus be based on the established heritage of the ISS whenever possible, and constrained by the existing and flight proven power systems of the ATV. However, there are five critical considerations for the design of the IM spacecraft that distinguish it from the ISS as shown in Table 19: Table 19 Critical considerations for power system.
ISS
ATV
IM
Spacecraft size
900 m3
48 m3
82 m3
Crew Size
3-6
0
2
Power Available
84 kW
4.8 kW
-
Eclipse Duration
45 min
45 min
50 min
Eclipse Frequency
90 min
90 min
Singular
Insolation
~1344 W/m2
~1344 W/m2
525-2600 W/m2
Spacecraft Attitude
Fixed Geocentric
Fixed Geocentric
Free
Required lifetime
>7 years
~1 year
1.4 years
Environment
LEO
LEO
Interplanetary Space
In addition, the power system must be designed to accommodate the variable insolation load along the interplanetary trajectory, as shown in Fig. 30.
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Figure 30 The estimated insolation levels on the spacecraft for the interplanetary trajectory.
4.6.3 Power Budget Consistent with the systems approach to the design of our vehicle, the power consumption of each component was estimated, and its role as an emergency critical component was evaluated. If it was judged that the component could not be powered down for a period of 3 hours without endangering the crew, it was designated an emergency critical component. Power budgets were then assembled for a nominal and emergency state for each vehicle system. The combined power budget for the entire vehicle is given in Table 20. Table 20 The vehicle power budget.
System ELCSS ADCS Crew Systems TCS C&DH COMM Total
Electrical Load (kW) Nominal Emergency 5.84 4.61 1.72 0.08 0.73 0.37 0.30 0.30 0.30 0.30 0.26 0.10 9.1 5.4
4.6.4 Power System Design Decisions Having developed the electrical power system requirements, we now detail our selected power generation, storage, and distribution architecture. PV-type: Based on a tradeoff analysis, emphasizing flight heritage and reliability, the best photovoltaic type was identified as the conventional silicon cell, which has demonstrated long duration 35
Team Kanau
success on board the ISS.55,56 However, the existing ATV photovoltaics do not provide sufficient power and will need to be upgraded. Accordingly, the newer ATK UltraFlex circular solar panels were selected to replace the existing ATV panels57. These panels are currently evaluated at TRL 6, and planned for deployment on the NASA CRV. Scaling the power system to operate with a sufficient margin in the worst possible isolation conditions (707 W/m^2), an array area of 65 m^2 was found to be sufficient. Such an array would allow for a 40% power margin under nominal conditions, and a 130% power margin under emergency conditions (again, assuming the worst insolation conditions experienced during the mission). However, at maximum insolation, this system is over designed, placing additional loads on the shunt regulator and thermal control systems. This problem can be avoided by distributing the PV area among the ATV’s existing four arrays locations, and stowing certain arrays when excess insolation is available. Distributing the 64 m2 among four 16 m2 arrays will allow sufficient power to be generated with a single array at perihelion, and two arrays at Earth departure and return. This has the added benefit of increased system redundancy in the later phases of the mission, when breakdowns and array degradation are most likely to occur. Battery Selection: The selection of a NiMH battery chemistry for the ISS is driven by their high energy density and charge cycle lifetime.55,56 However, lithium-ion batteries offer superior energy density, and are superior to NiMH in virtually every respect except for charge cycle lifetime. The lack of regular eclipses in the IM mission means the batteries are unlikely to be extensively cycled, and superior energy density becomes a decisive advantage. Li-Ion or Li-Pol chemistry batteries are identified to be the best alternative battery chemistry for powering the IM spacecraft systems. Based on the required power levels, and the duration of an expected emergency and eclipse, four 8-cell 28V Li-Ion batteries are identified as sufficient to provide the current, power and energy requirements for the IM spacecraft. Power System Architecture: With the addition of larger PV arrays and additional battery capacity, the existing ATV power systems architecture provides sufficient capacity for regulation and distribution. The standard 28 V DC dual redundant bus provides adequate backup and peak current capacity. However, modifications will need to be conducted to allow the ATV to transfer power to external systems through the PMA into the habitat. 4.7 Thermal Control Systems (TCS) Due to the relatively constant thermal load generated by the spacecraft, but the highly variable solar insolation, completely passive thermal management is not possible. Thus, the spacecraft will have two methods of thermal management: 4.7.1 Active Thermal Control (ATC) This is normal mode of operation. The spacecraft maintains a “sun astern” orientation, with the stern of the ATV pointed towards the sun, ensuring maximum power generation with minimal PV slewing, and presenting minimum illuminated surface for minimum thermal load. As on the ISS and ATV, heat is transferred from all systems, including the solar arrays via ammonia and water based coolant loops, where it is rejected through heat exchanger unit(s) and
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radiators.58 Precision modeling of this system is not possible at this point in the design, as it is extremely sensitive to the size of the spacecraft, and the insulation of the pressure hull. However, a rough first order model of the heat absorbed by the solar arrays and generated by the avionics, suggest the heat generated by each module could be rejected by exiting ATV systems, provided two deployable ISS radiator with an area of 5 m2 and an operating temperature of 100 °C was installed on the ATV. The installation of two radiators provides a degree of redundancy and variable capacity similar to the one provided by multiple solar arrays, as the insolation level changes throughout the mission. 4.7.2 Passive Thermal Control (PTC) Passive thermal control can be used for short periods, during times when the spacecraft cannot hold attitude with respect to the sun (e.g., course corrections maneuver, Mars flyby, etc.) or when insufficient power is available for ATC. Photovoltaics are stowed and power is supplied from the batteries to minimize thermal load on the spacecraft. A slow rotation is adopted to distribute the absorbed heat evenly across the exterior of the spacecraft. Heat rejection in PTC occurs principally through radiation from the shaded side of spacecraft. Detailed analysis of this technique is impossible without detailed understanding of the vehicle size and systems. Future work will focus on the design of the pressure hull and the cabin insulation to ensure that a safe interior temperature can be maintained for at least two hours in the event of a complete TCS failure.
4.8 Payload Mission According to ESA information kit, Mars500 Isolation Study, during any future human exploration mission to the Moon and Mars, the psychological resilience of the crew will play a critical role for the maintenance of health and performance and hence the success of the mission. One factor impacting on psychological resilience is the personal values of crewmembers defining their motivational goals and attitudes. Scientific experiments are important for this mission to make it more meaningful. We propose several experiments ideal for this long mission, with the four types of experiments in our proposal as shown in Fig. 31.
Figure 31 Examples of payload mission.
Protein Crystal Growth Experiment The microgravity environment, in which neither thermal convection nor sedimentation occurs, is ideal for growing high-quality protein crystals, as demonstrated on the ISS. Furthermore, there are many kinds of proteins with various functions.59,60,61 High quality protein structural information plays a key role in understanding the biological structure-function relationships and in the development of new pharmaceuticals.62 Thus, the astronaut crew can perform experiments similar to those conducted on the ISS, but in a different environment for longer time.
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Radiation & General Health Monitoring Measuring radiation data on spacecraft for long mission will be good reference data for future crewed missions. The data analysis for human spaceflight has been limited over the decades. However, several research experiments to monitor the radiation dose have been performed at the ISS.63 We suggest carrying PADLES, PS-TEPC and RRMD3 to measure radiological dosage, where this information can be used to plan further manned missions to Mars. This approach would also help warn the crew about increased radiation levels and urge them to seek shelter if required. Furthermore, regular medical testing and reporting back to Earth will ensure the well-being of the astronauts and also provide a reliable database of health effects and trends for a human spaceflight to Mars. A select suite of reliable diagnostic tools already used on the ISS is incorporated into the medical suite considered by Kanau. 3D Printer Since this deep space mission will require the crew and vehicle to be entirely selfsufficient, an on-board method to repair equipment is necessary. A 3D printer will be ideal for making necessary parts and tools on the spacecraft due to the limited quantity of spare parts and tools. NASA is planning to send first 3D printer to space in 2014, thus increasing the TRL of this technology to a level suitable for inclusion in the 2018 flyby mission.64 Observatory During this long mission, the crew will need mental refreshment. Observation outside the windows using telescope will fascinate and inspire the crew in the spacecraft and thus should be included for both scientific observation and health reasons. This would be the first time humans will be able to view the surface of Mars with the naked eye and the images they send back to Earth would inspire humanity to further space exploration and development.
5 Outreach Much like Neil Armstrong’s first step onto the lunar surface, the crewed Martian flyby would define a new epoch in human spaceflight, and indeed human history. Therefore, we present some options that could be worthy of consideration for spreading awareness of the mission and to the share the passion of space travel and exploration. 5.1 Astronaut Fashion Competition Initiate a competition for T-shirt designs in the spirit of Mars exploration, which can be incorporated into astronaut day-to-day wear. This would allow for people to connect with the mission. Also, an International ‘fly-by’ day could be organized where the whole world would wear the winning design when the astronauts entered the Martian fly-by phase. This is a great opportunity for the astronauts to connect with the whole world and vice-versa! If producing the T-shirts is infeasible, then a celebration or gesture of some sort could be coordinated around the world. 5.2 World Cancer Day Support Space travel is a very risky and dangerous business, and space radiation in particular can be harmful and even lethal, depending on the dosage received. World Cancer Day is about spreading the awareness of cancer and combating misconceptions about this condition. The
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astronauts could take part in this effort and maybe shave their heads on this day to connect with cancer patients, survivors, and healthy people who expose themselves to such radiation for various reasons. 5.3 Student Mission Simulator A mission simulation tool can be placed on the Inspiration Mars website wherein students, and even the general public, will create their own potential flyby missions to Mars. In fact, this effort could be coordinated with the producers of such popular games as Kerbal Space Program or similar software packages. Thus, students can learn about the unique opportunities and challenges of human spaceflight while connecting on a mental and emotional level to the experiences of the Inspiration Mars crew. 5.4 Thought Exchange Avenues could be arranged for people to send in letters that the astronauts would be able to open every day of their journey. Video snippets could also be transmitted after the astronauts begin their journey. In particular, shows of support or natural scenes would in particular have a positive effect on the mental and emotional health of the crew. In return, responses and/or thoughts of astronauts sent back from the spacecraft could be aired during talk shows/posted online for public viewing. Re-connecting with events on Earth would provide astronauts with fond memories and, additionally, would provide people with the opportunity to understand more about both the joys and discomforts associated with space travel. 5.5 Further Design Competitions Similar to the current student design competition and the competition concept for the astronaut clothing fashion, specific design competition opportunities can be presented to the public. These design competitions, when targeted at a general audience, can garner large domestic and international support for the Inspiration Mars flyby mission and human spaceflight in general.
6 Conclusion Team Kanau’s architecture for a two-person flyby mission of Mars in the year 2018 presents opportunities for the incorporation of many intriguing technologies and techniques for the 501 day crewed journey. We propose astronauts Shannon Walker and Andrew S. Thomas, a married couple of 9 years, as the crew of the spacecraft. Three Falcon Heavy launches will deliver the crewed capsule, along with the required DCSS propulsion stages for the Trans-Mars Injection, to a LEO staging orbit; this launch program relies only upon current or near-term technologies from SpaceX and United Launch Alliance, and so reduces the risk of delays to the schedule and offers significant cost savings compared to other available launch systems. Aerocapture upon return to the Earth will mitigate high re-entry velocities and enable the use of a SpaceX Dragon capsule for the crew re-entry and descent to the surface of the Earth. We consider several factors to maintain the physical, mental, and emotional well being of the married astronauts. Adequate space and privacy are provided within the capsule, and a robust communications system is designed to minimize the chance of crew isolation from Earth. We propose a regenerative water, oxygen and carbon dioxide recycling system based upon the ISS with a back-up set of storage tanks for water and air circulation. This hybrid approach reduces system mass whilst ensuring a continued supply of potable water and breathable air.
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While most food for the crew will be pre-prepared and packaged, a limited supply of fresh vegetables and seasonings can be grown on-board; in addition to providing variety to the crew diet, the growing and nurturing of the plants will give needed mental stimulation and a psychological connection to Earth for the married couple. Incidental tools and equipment for the crew can be manufactured on-board via the use of 3-D printing, a technology that will soon be demonstrated in micro-gravity environments upon the ISS. Simulations of the ECLSS are performed using SICLE, the Simulator for Closed Life and Ecology. These novel human factors technologies, along with the proposed launch and Earth return scenarios, are key components that enable deep space human missions, for example the proposed crewed flyby of Mars in 2018, in both the near- and far-term.
7 Kanau Prospects for 2021 Mission Opportunity Our design is directed toward the exploitation of the free-return flyby of Mars available in 2018. As such, there is a critical need to incorporate technologies with a proven flight heritage as well as pursue an aggressive design, test, and operations cycle. However, as Inspiration Mars has altered their baseline architecture to exploit the 2021 flyby opportunity, many potential avenues of exploration could prove fruitful within the extended development time afforded by this later departure window. That is, from a programmatic level, more extensive Phases A-D permit further development and refinement of advantageous technologies and the maturation of novel concepts. For example, the ACES module being developed by ULA could reach flight status, affording greater capability during the TMI. Furthermore, the use of deep space propulsion, either with ACES, SEP, or another system, could enable greater mission flexibility. Furthermore, the SLS is slated to reach operational status by 2021, providing greater capability for launch from the surface of Earth as well as, perhaps, enabling direct insertion of the crew into the Mars-bound trajectory. Further reducing the hazard posed by radiation, the development of active radiation shielding, even if just for the ERV, could further reduce the doses received by the crew. In addition to active radiation shielding and propulsion systems, extensive modifications are desirable for the ATV, MPLM, and Ultra-Flex solar arrays and a longer development cycle can improve the quality of these alterations as well as reduce the associated development cost. Furthermore, as the currently closed end of the MPLM will be replaced with a docking mechanism, this modification also enables the placement of larger viewing windows, similar to the Cupola on the ISS. Greater analysis of the aerocapture optimization will increase the reliability of the Earth-return phase and perhaps indicate beneficial modifications to the Dragon capsule, especially its lifting abilities. Further advances in medical knowledge and capability could improve the odds of crew survival and long-term health upon return to the Earth, while bio-regenerative systems for the ECLSS and food supply could increase self-sufficiency and reliability. Novel concepts such as the combined use of a tether as well as the expended TMI propulsion module to simulate Martian gravity could demonstrate critical techniques needed for future deep space human exploration. Finally, selection of the 2021 departure window could enable a close approach with Venus, Earth’s sister planet. While Venus is not a viable prospect for near-term colonization by human beings, a human encounter with this planet could spur additional public interest in human exploration of the solar system.
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8 Kanau Team Workflow, Website and Animation Video Table 21 Major tasks performed by individual team members.
Team Member Ayako Ono Ashwati Das
Daichi Nakajima Eriko Moriyama Jeff Stuart
Koki Tanaka Kshitij Mall
Max Fagin
Nick Gillin Shota Iino
Takuya Ohgi Yuri Aida
Major tasks Kanau spacecraft’s interior design and facilities for radiation protection. Trajectory design and launch vehicle selection, concept of operations, safety analysis, reliability assessment, risk management, radiation protection, and cost assessment, outreach, mission schedule, 2021 prospects. Launch vehicle selection. Environment control, life support system and its cost. Trajectory design and launch vehicle selection, concept of operations, safety analysis, reliability assessment, risk management, radiation protection, outreach, mission schedule and cost assessment, 2021 prospects. Kanau spacecraft’s interior design. Project management, systems design, launch vehicle selection, concept of operations, aerocapture, command and data handling, communications, safety analysis, crew selection, crew schedule, mission schedule, website development and logo design. Systems design, power systems, thermal control system, radiation protection, command and data handling, communications, mission schedule, vehicle stack design, concept of operations and logo design. Mission animation videos. Project management, systems design, payload mission, Kanau spacecraft’s interior design, environment control and life support system, and facilities for crew physical health and radiation protection. Environment control and life support system. Facilities for radiation protection and crew physical health. Table 22 Major roles of team's advisors.
Advisor Dr. H. Miyajima Dr. M. J. Grant
Major roles Mission design, and environment control and life support system. Selection of launch vehicle, command and data handling, communications, and aerocapture.
Team Website Address: https://sites.google.com/site/occupyplanet4/ Kanau Mission Animation Video: https://sites.google.com/site/occupyplanet4/animation Appendix of this report: https://sites.google.com/site/occupyplanet4/final-report
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9 Acknowledgments The team would like to thank the Mars Society and Inspiration Mars for providing a wonderful opportunity to work on a real world project. Special thanks to the following contributors: Mr. Alfredo Caro (Los Alamos National Laboratory, USA) - Radiation Dr. Dan Dumbacher (Professor, Purdue University, USA) – Systems design Dr. David Filmer (Professor, Purdue University, USA) – Communications Dr. James Longuski (Professor, Purdue University, USA) – Aerocapture Mr. Jim D'Entrement (Graduate Student, Purdue University, USA) - Propulsion Dr. Kazuhiro Terasawa (Professor, Keio University, Japan) – Radiation analysis Mr. Kiosuke Murakawa (Japanese Mars Society) - Publicity Ms. Magdalena Serrano de Caro (Los Alamos National Laboratory, USA) - Radiation Mr. Marat Kulakhmetov (Graduate Student, Purdue University, USA) - Propulsion Ms. Nasa Yoshioka (Graduate Student, Keio University, Japan) - ECLSS Mr. Rizwan Qureshi (NASA Goddard Space Flight Center, USA) – Aerocapture analysis Mr. Seiko Shirasaka (Professor, Keio University, Japan) - Safety analysis Dr. Shin Yamada (Lecturer, Kyorin University, Japan) – Medical analysis Mr. Shreyas Subramanian (Graduate Student, Purdue University, USA) - Systems design Dr. Stephen D. Heister (Professor, Purdue University, USA) - Propulsion Mr. Thomas Antony (Graduate Student, Purdue University, USA) - Aerocapture analysis Mr. Yoshiki Annou (Japanese Mars Society) – Publicity
10 References 1
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Shiela Bailey and Ryne Raffaelle, “Space Solar Cells and Arrays”, Ch. 10, Handbook of Photovoltaic Science and Engineering. John Wiley & Sons, 2003. 11 A.R. Jha, “Next-Generation Batteries and Fuel Cells for Commercial, Military, and Space Applications”, CRC Press, 2012. 12 “Military Handbook: Reliability Prediction of Electronic Equipment”, U.S. Department of Defense, MIL-HDBK-217F, Dec. 1991. 13 Reliability Information Analysis Center, Nonelectronic Parts Reliability Data 1995. U.S. Department of Defense, NPRD-95, 1995. 14 “Guidelines for Risk Management”, NASA, S3001, Rev B. URL: http://ims.ivv.nasa.gov/. Accessed 25 Jul. 2014. 15 D. M. Harland and R. D. Lorenz, “Space Systems Failures”, Praxis Publishing, 2005. 16 Ono, A. et al., “Designing Crew Habitats for Long Term Physical and Psychological Health and Radiation Safety,” IAA Space Exploration Conference, Planetary Robotic and Human Spaceflight Exploration, 2014. 17 Ono, A., Irene Lia Schlacht. “Space art: aesthetics design as psychological support,” Personal and Ubiquitous Computing, Vol. 15, No. 5, pp. 513-517. 18 Schlacht, I. L., Birke, H., “Space design visual interface of space habitats” Personal and Ubiquitous Computing, Vol. 15, No. 5, pp. 498-503. 19 Schlacht, I. L., et al. "Human factors in space mission," A. Lichtenstein, C. Stößel, & C. Clemens (a cura di)(Eds.), Der Mensch im Mittelpunkt technischer Systeme 8, 2009, pp. 326-330. 20 Ono, A. et al. “Countermeasures for Psychological Issues: Soundscape Design for Confined Environments,” 60th International Astronautical Congress, Behavior, Performance and Psychosocial Issues in Space. Daejeon, Korea. IAC Paper 09: A1.1.11., 2009. 21 Astronauts’ Dirty Laundry”, NASA’s Aerospace Technology Expertise. URL: http://www.nasa.gov/vision/space/livinginspace/Astronaut_Laundry.html/. Accessed 26 Jul. 2014. 22 Malik, T., “Japanese Space Underwear Keeps Stink Out”, SPACE.COM. URL: http://www.space.com/7077-japanese-space-underwear-stink.html/. Accessed 26 Jul. 2014. 23 “Intravehicular Activity Clothing Study (IVA Clothing Study) - 07.22.14”, NASA. URL: http://www.nasa.gov/mission_pages/station/research/experiments/1084.html/. Accessed 26 Jul. 2014. 24 NASA-STD-3001 NASA SPACE FLIGHT HUMAN SYSTEM STANDARD VOLUME 1: CREW HEALTH, VOLUME 2: HUMAN FACTORS, HABITABILITY, AND ENVIRONMENTAL HEALTH, 2007. 25 Trappe, S. et al. "Exercise in space: human skeletal muscle after 6 months aboard the International Space Station," Journal of Applied Physiology, 106.4: 1159-1168, 12 Jan. 2009. 26 McPhee, J.C. and Charles, J. B., “Human health and performance risks of space exploration missions,” NASA. 27 Matsumoto, T., “Space Medicine Contributes to Medical Advances on Earth,” JAXA, URL: http://www.jaxa.jp/article/special/expedition/matsumoto01_e.html/. Accessed 17 Dec. 2013.
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