Mission Design of a Two-Person Mars Flyby by 2018 International Student Design Competition Team Mars18 - www.mars18.de Margret Barkmeyer Nils Hoffrogge Ferdinand Leinbach Mirjam Schmidt Winfried Burger Heiko Joos Victor Mosmann Rolf Stierle Felix D¨uver Peter J¨ustel Fabian M¨uller Lukas Teichmann Eduardo Finkenwerder Jochen Keppler Paul Nizenkov Tobias Torgau Dan Fries* Ronja Keuper Duncan Ohno Daniel Wischert Stefan Fuggmann Alexander Kunze Adrian Pfeifle S¨orenHeizmann Jonas Lay Minas Salib Christina Herr Hong Anh Le Marcel Scherrmann University of Stuttgart

Table of Contents Page

1 Introduction1

2 Executive Summary2

3 Mission Architecture3 3.1 Trajectory...... 3 3.2 Launcher Selection & Manifest...... 4 3.3 Reentry...... 7

4 Human Factors 11 4.1 Astronaut Selection...... 11 4.2 Crew Health...... 12

5 Spacecraft Design 15 5.1 Configuration & Structure...... 15 5.2 Subsystems Design...... 17 5.3 Scientific Payload...... 35 5.4 Systems Engineering & Budgets...... 36

6 Programmatic Issues 38 6.1 Cost...... 38 6.2 Roadmap & Schedule...... 41 6.3 Risk Management...... 43

7 Conclusion 46

Bibliography 46

*Point of Contact: Dan Fries, [email protected] 1 Introduction

To increase the pace in manned Mars exploration, the Mars Society in collaboration with Dennis Tito’s Inspiration Mars Initiative called for students around the world to develop a complete mission concept for a manned Mars flyby in 2018. The ultimate goal is not only to complete this very ambitious mission but to spark more interest in comparable missions around the world and further a technological competition to put a human being on Mars on a peaceful level for the greater benefit of mankind. Previous efforts are limited to robotic missions and so far not even a sample return mission has been accomplished. While robotic missions are certainly sufficient to gather simple, pre-determined scientific data, they are not fit to actually expand humanity’s sphere of influence. In-situ research and human settlements, however, enable access to resources and hold future potential on a completely different scale. Future problems, like the overpopulation of Earth and lack of essential resources like water, could be tackled right now. Even in the present day, solar system exploration efforts would immediately result in a myriad of scientific and technological developments, employment opportunities and eventually direct financial gain. Furthermore, private space companies and NASA are already working towards heavy-lift launch vehicles, but so far without a clear application. All of the factors mentioned above contribute to the overall result that manned space exploration is not only desirable but also achievable. Mars18 is the student-led team at the University of Stuttgart, Germany, that has taken on the challenge proposed by the Mars Society and Inspiration Mars. The team’s goals are the meaningful contribution to the worldwide efforts in space exploration, education of high- potential students in a hands-on project, engagement of public interest in space exploration through cooperation with local media and achieving a high ranking in an internationally acclaimed competition. As a mission like this has never been attempted before, the design presents a special challenge that also sparks creativity in every person involved. The team consists of about 29 students, most of whom are aerospace engineers. But it is obvious that such a project cannot be successful solely through aerospace technology. From the beginning, the team attempted to achieve a multi-disciplinary composition of motivated participants. Mars18 is comprised of students from medical sciences, social sciences, electrical engineering and economics. Moreover, the team managed to gain professional support from the Institute for Space Systems (University of Stuttgart), Astos Solutions, Constellation (distributed platform for aerospace research) and Airbus Defense & Space. The presented work attempts to show how a manned mission to Mars could be executed realistically by 2018. In general, conservative assumptions were preferred over optimistic ones, in both technological and cost issues. Key technologies that would further access to space in general and for this specific mission were identified and a time schedule developed that would allow for their implementation. Although a certain amount of technologies is employed that have to be qualified, this is only done in absolutely necessary cases or because it presents a considerable advantage. Human factors were evaluated and accounted for. Finally a complete cost estimate was conducted.

1 2 Executive Summary

During the entire development, the Mars18 team followed four principles: simplicity, safety, low cost and feasibility. To evaluate the amount of development still required an estimated Technology Readiness Level (TRL) is used. Through rigorous optimization and evaluation of available systems it is possible to lower the total mass below 15 t and the low Earth orbit (LEO) mass amounts to ∼ 63 t. A concept is devised that allows to launch the entire mass with only two starts of currently or soon to be available carrier rockets. Trans-Mars injection (TMI) is accomplished via staged propulsion of two modified Delta IV 4-m Second-Stages. The system consists of modified versions of the Enhanced Cygnus (referred to as Cygnus) and the DragonRider (referred to as Dragon). Both modules have already been tested in their basic configuration and are currently under fur- ther development. An important mission like this requires absolute priority among the deep space communication systems on Earth (i.e. Deep Space Network (DSN), ESTRACK). To reduce mass several of the AOCS’ thrusters are resisto- jets, using waste products from the life support system. Additionally, a fuel saving model predic- tive control (MPC) algorithm is employed for attitude control. To account for the dangers of radiation exposure outside of the Earth’s magnetosphere, a protection scheme is devised that works highly synergetic with the equipment of other subsystems and provides a storm shelter for solar particle events (SPE). The life support system is designed from scratch and follows a virtually close-loop approach. It also introduces two devices that have not been used on previous missions but are able to reduce the required initial mass considerably and increase the synergies with other systems. As the flight trajectory, a free-return option is chosen that requires ∼ 4.8 km/s from a 350 km LEO. At return to Earth’s atmosphere, the reentry capsule will have a relative velocity of ∼ 13.8 km/s. The resulting kinetic energy is dissipated during two passes through the atmosphere before descending to the ground. The total mission duration from TMI is ∼ 501 days. The presented concept to deal with human factors handles physical as well as psychological issues in a very isolated and confined space. The well-being is of great importance, since they should perform experiments and document as many things as possible during their journey. Thus, the endeavor will result in the maximum scientific benefit for future missions and possible spin-offs. The total cost of such a mission is estimated using current prices, heritage data and cost models. Thus providing an amount of 4.3 B$ that should be around an upper limit for the presented design. In a simple risk analysis not only the dangers stemming from technological aspects but also programmatic issues are presented and how they might be mitigated.

2 3 Mission Architecture

The approach of choosing the most favorable trajectory is presented in this chapter. Further- more, available launcher systems are compared in detail and finally reentry is discussed.

3.1 Trajectory

The objective of the trajectory optimization is to find a feasible flyby trajectory to Mars. By solving the Lambert problem, options for such a trajectory are investigated. Therefore, astrodynamics and interplanetary spaceflight have to be considered.

3.1.1 Tools & Boundary Conditions POINT is a Lambert-solver by Astos Solutions. Ephemerides provided by Jet Propulsion Laboratory (JPL) are used to solve the Lambert problem, providing an accurate initial estimation. POINT determines and optimizes trajectories based on constraints and boundary conditions. Possible optimization constraints are minimum flight time as well as low departure and arrival C3 energies. In the following, different trajectories are compared according to mission requirements. GMAT (General Mission Analysis Tool) is an open-source space mission analysis tool provided by NASA. It enables the simulation of gravitational forces of all celestial bodies in the solar system. GMAT is used for verification of the final trajectory. The following boundary conditions constrain the trajectory design. Launchers starting from Cape Canaveral will lift the spacecraft into a circular LEO. The assembly orbit is at an altitude of 400 km and likely to decrease due to atmospheric drag. Therefore, the interplanetary trajectory is set to start at a 350 km altitude with an inclination of 28.5° in the equatorial plane. Additional mission requirements are a flyby altitude of 100 km at Mars to prevent aerodynamic drag and to start the mission in 2018.

3.1.2 Trajectory Trade-Off Multiple possible trajectories were found with POINT. The focus is on the ∆v at Earth and Mars, total mission time and excess velocity v∞ for arrival at Earth. With fixed mission costs, higher ∆v leads to a smaller throw mass. Shorter mission duration leads to higher v∞ for return to Earth, which in turn significantly increases the loads and stresses on the thermal protection system. Since all investigated trajectories result in v∞ higher than at every reentry that has been done, trajectories with low v∞ are favored. As all mission concepts have a duration of at least one year but only those shorter than 120 days have a significant impact on the Environmental Control and Life Support System (ECLSS) design, it did not influence the trade-off. Although, it can be concluded that a shorter mission duration decreases the total system mass. Thus, a trade-off is necessary between mission duration, total mission mass, the resulting ∆v as well as v∞ upon arrival at Earth. Three suitable trajectories found with POINT are presented in Table 3.1. The first trajectory with a start date in May 2018 and a duration of 435 days requires a ∆v of 1560 m/s at Mars and has a v∞ of 11 km/s on return to Earth. However, an additional

3 3 Mission Architecture 3.2 Launcher Selection & Manifest

Table 3.1 Considered Earth-Mars trajectories found with POINT

Start Date - Arrival Date ∆v Departure ∆v Flyby v∞ Arrival Duration 1 02/05/2018 - 07/12/2019 3586.2 m/s 1558.6 m/s 11 070.1 m/s 435 d 2 07/15/2018 - 07/14/2019 9342.8 m/s 0 12 835.0 m/s 361 d 3 01/04/2018 - 05/19/2019 4825.5 m/s 0 8786.3 m/s 501 d propulsion system is required to provide a ∆v at Mars, which increases system risk and adds a single point of failure to the mission. To conclude, the decrease in flight duration does not justify the additional propulsion module. The second trajectory offers a shorter flight time, but the required ∆v is not feasi- ble. The third trajectory is a good trade-off between ∆v, duration and return velocity. Launch Earth Orbit The detailed data for this free-return trajec- tory verified with GMAT can be found in Table 3.2. A C3 of 38.7 km2/s2 results in a Sun v∞ of 6220.7 m/s at TMI. For a circular de- parture orbit of 350 km altitude, with a ve-

locity of 7702.0 m/s and a hyperbolic veloc- Venus Orbit ity at the same altitude of 12 543.4 m/s, the Reentry required ∆v for TMI is 4825.5 m/s. This Flyby change in velocity has to be provided by Mars Orbit the propulsion system. No further impul- sive maneuver is required to return to Earth Figure 3.1 Planned free-return trajectory safely after a flight duration of 501 days. A schematic of the selected trajectory is presented in Figure 3.1. Table 3.2 Free-return solution: right ascension (RLA) and declination (DLA) of the outgoing hyperbolic asymptote relative to the Earth/MarsMJ2000Ec-frame of GMAT

2 2 Date v∞ [m/s] DLA RLA vperi [m/s] C3 [km /s ] Departure Earth 04.01.2018 18:59:35.501 6220.7 −89.3◦ 16.6◦ 12543.4 38.7 Mars flyby 20.08.2018 18:31:38.676 5375.0 −1.53◦ −121.9◦ 7258.1 28.9 Arrival Earth 20.05.2019 16:28:42.508 8855.7 3.6◦ −72.9◦ 14186.9 78.4

3.2 Launcher Selection & Manifest

Launchers are responsible for launching the Mars Transfer Vehicle (MTV) and propulsion module (PM) into the defined parking orbit in LEO, where the TMI will be carried out.

3.2.1 Design Process & Requirements In general, a launch concept with the least launches is preferable in order to keep down mission and system complexity while increasing crew and mission safety accordingly. This requirement, however, demands both a man-rated launch vehicle which is capable of launching the MTV as well as a second launch vehicle capable of launching the PM.

4 3 Mission Architecture 3.2 Launcher Selection & Manifest

In order to find the ideal launch concept for the mission, carrier rockets from medium- to heavy-lift of the last 60 years were systematically analyzed and the results can be found in the appendix [36]. Launchers, which fulfill the criteria of high payload mass and high reliability, were taken into further consideration. A smaller subset of launch vehicles will be man-rated in 2018 and most of them are not able to launch the complete MTV by a single launch. The collected information was used to develop different concepts of launcher combinations with the respective payloads. The most promising concepts were further analyzed. Due to the fact that there is no launch vehicle in 2018 which is man-rated, flight-proven and able to launch the PM and the MTV by a single launch, a concept with two launchers is chosen. With only two launches, the risk for the crew is minimized in addition to ensuring low cost and high mission safety. One lighter, man-rated and highly reliable launch vehicle to carry the crew is used and another to carry the PM. This concept is preferred due to its reduction of complexity for rendezvous maneuvers, lower boil-off of propellant and the design lifetime of the electrical power system (EPS) and attitude and orbit control system (AOCS) of the PM.

3.2.2 Final Launch Concept The PM is launched with a slightly modified Falcon Heavy on December 21, 2017 from Launch Complex 39 Table 3.3 Payload distribution at Kennedy Space Center (KSC). The crewed MTV is Launcher Payload Mass launched on the December 24, 2017 by an Atlas V 441 Falcon Heavy PM 48 027 kg from the planned launch complex for manned Atlas V Atlas V 441 MTV 15 000 kg missions also from KSC. The payload distribution is de- tailed in Table 3.3. All launches from KSC during the past five years in December and January were analyzed and it can be concluded that the weather conditions are not a critical factor for the launch window, whereas technical malfunctions can indeed be a critical variable. Therefore, three launch possibilities for each vehicle with an interval of one day for the Falcon Heavy and a three day interval for the Atlas V 441 are scheduled. The begin of each launch is timed to achieve the right orbit parameters for parking orbit. After the PM has reached the orbit, it transforms into the injection configuration and acquires a gravity-gradient stabilized position to minimize the required fuel for the AOCS. The change of the configuration is further explained in Section 5.2.5. Once the MTV reaches LEO, the crewed Dragon capsule detaches from the Dragon trunk, flips over and docks to the docking port of the Dragon trunk. After orbital alignment is arranged by the MTV within 7 h, it docks to the PM. After assembly completion, there are at least two days to test all systems and to prepare the MTV for TMI. Figure 3.2 illustrates the launches and on-orbit assembly. On January 4, 2018, the first stage of the PM is ignited to achieve a high elliptical orbit, which provides the correct Keplerian elements of the Mars transfer orbit for ignition of the second stage. As burnout of the first stage is complete, the MTV and second stage separate and perform trajectory corrections at the apogee. At perigee, the second stage is ignited and the final TMI is performed. Both ignitions produce a maximum acceleration of 0.5 g, due to the relatively low engine thrust of 110 kN. This acceleration is easily tolerable and was already proven and tested during the Gemini 10 mission. In order to validate the required modifications of the launch vehicles, slight changes in already scheduled missions are planned. Firstly, negotiations have to be conducted with the

5 3 Mission Architecture 3.2 Launcher Selection & Manifest

Figure 3.2 Batchart of launch and assembly U.S. Department of Defense or Intelsat to perform their planned Falcon Heavy launches with the modified Asymmetric Payload Fairing (APLF) in 2015 and 2017, respectively. Therefore, it is intended to pay 75% of the launch costs to the respective customer, which is also considered in the overall mission costs. Secondly, a test rendezvous is planned between the Cygnus module and the Dragon at COTS-6 from SpaceX and COTS-7 from Orbital Science. The Cygnus module stays docked at the ISS until the manned Dragon also docks with the ISS. After undocking of Cygnus and Dragon, they will perform a rendezvous maneuver before their reentry to validate the modified docking port. Moreover, the proposed launch concept takes full advantage of the Commercial Crew Development (CCDev) program, which is initiated by the U.S. government and promotes the upcoming private space industry in the U.S. For the improbable case that it is not possible for SpaceX to modify the Falcon Heavy with the APLF by 2018, an alternative launch concept is provided. This backup intends to replace the Falcon Heavy with an Atlas HLV, which carries a Delta IV 5-m Second-Stage, and a Delta IV Heavy carrying a Delta IV 4-m Second-Stage. These stages are modified to be able to perform a docking maneuver. Further information about this alternative launch concept can be found in the appendix [36].

3.2.3 First Launch (Falcon Heavy) To execute the mission with as few launches as possible, usage of heavy-lift launchers is necessary. An initial evaluation was Table 3.4 Heavy-lift launcher trade-off performed during the design process, after which several launch Payload vehicles were selected for further consideration as shown in Table Launcher 3.4. The most promising launch vehicle is the Space Launch to LEO System (SLS). However, as the first launch is scheduled as late SLS 61–81 t as 2017, this mission would be its maiden flight. Therefore, and Falcon Heavy 53 t because of the high cost, it is not considered for the launch concept. Angara A5 24.5 t Delta IV Heavy 23 t This is why SpaceXs Falcon Heavy is utilized. It will have its Ariane 5 21.5 t first launch in 2014 and, hence, has a sufficient trial period. To Proton 20 t launch the PM with the Falcon Heavy, it is necessary to enlarge

6 3 Mission Architecture 3.3 Reentry

the payload fairing so that it offers enough space for the PM. An aerodynamically optimized payload fairing is proposed, which keeps the aerodynamic loads equal and increases the weight only by 30% compared to the default fairing [44]. The used APLF is 18.5 m tall and 9.5 m wide and is able to host the PM in its launch configuration. In addition, the fairing is equipped with ventilation in order to release the escaping hydrogen of the cryogenic tanks of the PM. Furthermore, the upper stage of the Falcon Heavy is equipped with additional thrusters and more fuel for its AOCS to be capable of controlling the changed center of gravity of the payload. To account for these modifications, such as the increased structure mass and the parking orbit above the 200 km reference orbit given by the launch provider, a 10% margin on the overall payload capacity is included in the calculation.

3.2.4 Second Launch (Atlas V 441) As shown in Table 3.5, only a few launchers are man-rated and capable of launching enough payload to LEO. Soyuz-FG is fully Table 3.5 Man-rated launcher trade-off developed and flight-proven in multiple missions but can only carry Payload light payloads. According to SpaceX, Falcon 9 will be man-rated by Launcher 2016. Additionally, Falcon 9 is cheaper, but Atlas V is able to carry to LEO heavier payloads. The Atlas V family is flight-proven. It is one of SLS 61–81 t the most reliable launch vehicles in the world and every U.S. mission Atlas V 8.9–17.7 t to Mars in the last decade was launched by an Atlas V. In addition, Falcon 9 11 t Soyuz-FG 7.1 t Boeing plans to launch their CST-100 capsule in 2016 with an Atlas V and also the Sierra Nevada Corporation plans to launch their DreamChaser in November 2016 with an Atlas V. Therefore, Atlas V is man-rated and extensively tested with multiple crewed capsules until 2018. The Atlas V family can be combined in different ways and the chosen launcher will be an Atlas V 441. It uses four solid rocket boosters and a modified four meter payload fairing. In addition, the upper stage of the Atlas V 500 series is utilized to account for the heavier payload and its shifted center of gravity. To host the MTV, the fairing is modified to cover the Cygnus only. The upper part of fairing is removed and replaced by the Dragon capsule, so it carries the weight of the Dragon. With this modification, Cygnus is not carrying the weight of the Dragon at launch and SpaceXs launch abort system, which is integrated in the Dragon, can be used. In order to take account of these modifications, a margin of 8.5% on the payload capacity of Atlas V 441 is included in the calculation.

3.3 Reentry

Two main factors impose restrictions on the reentry: Loads must be kept below 8 g and the duration must not exceed 14 h. The time constraint is due to limited battery capacity and heat sink, which is used to dissipate waste heat after detachment from the trunk and the Cygnus module.

Considered Reentry Scenarios The reentry maneuver is challenging and, if not altered, will be the fastest manned reentry. Several efforts and ideas to reduce the speed were considered:

7 3 Mission Architecture 3.3 Reentry

ˆ Slowing down the spacecraft with an additional propulsion system ˆ Capturing the spacecraft and slowing it down with a pre-deployed capture vehicle ˆ Using the Moon for a gravity assist maneuver to slow down the spacecraft ˆ Aerobraking with multiple passes before reentry ˆ Using a spring/tether mechanism to achieve a slowdown of the vehicle (quickly discarded due to high mass/∆v ratios)

The first and second option exploit the same idea: employ propulsion systems to slow down the spacecraft. For the first option, only solid rocket boosters are feasible due to the boil-off of cryogenic propellant and a relatively low Isp of storable propellant. This results in a drastically increased mass of the propulsion module and, thus, in high launch cost. While the second option could utilize a pre-deployed propulsion module with electric propulsion, it increases mission complexity and requires an additional launch as well as the development of an additional module. Also, both options suffer from a single point of failure issue: An engine failure and/or a missed capture/rendezvous would lead to a loss of crew (LOC) as well as loss of mission (LOM). The gravity assist option at the Moon is not possible since its position is on the opposite side of Earth at the time of arrival for the selected trajectory. Therefore, only aerobraking is further investigated. To approximate aerobraking reentry scenarios, the tool ASTOS from Astos Solutions was used. ASTOS is a software package to simulate and optimize launch and reentry trajectories.

Assumptions of the Reentry Trajectory Reentry loads depend on a range of factors, including spacecraft geometry and trajectory. Al- though the latter is known, geometry data is unavailable due to proprietary issues. Important geometric characteristics of a reentry probe are: cD/cL values (and the respective ballistic coefficient) and nose radius. The Apollo capsule [41] serves as a reference for cD/cL values since it is similar in size, weight and heat shield diameter. The cD/cL values are considered as a function of altitude at constant velocity, since for constant altitudes and varying velocities aerodynamic values do not change significantly. To account for deviations from these values, a sensitivity study was conducted. The study demonstrated that variations of cD/cL within a reasonable range only have a minor impact on the reentry trajectory. For every combination of tested cD/cL values within the range of the Apollo values, a viable reentry trajectory could be found. Graphical approximations from drawings in the official DragonLab fact sheet led to estimations of the nose radius. From these, a calotte was calculated with a nose radius of r = 4.7 m. The respective data sheet and sensitivity study for the aerodynamic coefficients can be found in the appendix [36]. The heat flux consists of the convective heat flux and radiative heat flux. For all trajectory calculations the 1976 US Standard Atmospheric Model is used. Convective heat flux is covered by the Chapman heat flux model, which is given by

n m q˙conv = C ρ V , (3.1)

8 3 Mission Architecture 3.3 Reentry where C = 1.705 × 10−4, n = 0.5 and m = 3 are model parameters. ρ and V denote the air density and velocity, respectively. This model yields good estimates for convective heat flux and is within a margin of error verified for a wide range of existing missions. Estimating the radiative heat flux is more challenging. There are two widely known radiative heat flux models, the Tauber-Sutton model [53] and the Detra-Hidalgo model [14]. The former is known to be fairly accurate, but is only applicable for nose radii between 0.3 m < r < 3 m. The determined nose radius is therefore not within the applicable range. A study and verification of both models was performed for the Stardust Return Capsule. The nose radius was varied over a wide range to show steadiness and applicability. The result suggests using the Tauber-Sutton model for the approximation. The heat flux calculated by Tauber-Sutton for the investigated nose radii is larger by one order of magnitude than the results of Detra-Hidalgo. Furthermore, the heat flux during reentry of Stardust computed with the model of Tauber-Sutton compares well with literature values [34]. Finally, it is demonstrated that by using Tauber-Sutton out of range of its applicability, ”the resulting error is still generally within the range of uncertainty found in computing the radiative heating with a more computationally intense method” [13].

Results of the Reentry Simulation Based on the final reentry mass of 3900 kg and self-imposed constraints (load factors less than 8 g and total reentry time less than 14 h), a path analysis was performed. The altitude and velocity as well as the load factor over time are shown in Figure 3.3. 12500 15 8

10000 6 10 7500 Altitude 4 5000 Velocity

Altitude [km] 5 Load factor [g] Velocity [km/s] 2 2500

0 0 0 0 1 2 3 4 5 0 1 2 3 4 5 Time [h] Time [h] (a) Velocity and altitude (b) Load factor Figure 3.3 Reentry at perigee altitude and bank angle of: 63.5 km and 98.5◦ (first pass), 65 km and 65◦ (second pass) Figure 3.4 presents the total heat load integrated over time and the heat flux. A worst-case scenario at the lower reentry window at 60 km is assumed here, which differs from the nominal trajectory with three atmospheric passes during aerobraking. Moreover, it was found that a precise control of the bank angle µ is crucial. Hence, keeping the bank angle within a certain margin during reentry is an important requirement for the AOCS. Generally, it can be said that higher perigee altitudes require higher bank angles

9 3 Mission Architecture 3.3 Reentry

1000 103 ] ] 2 2 800 101 Heat Flux 600 Heat Load 10−1 400 Heat Flux [W/cm 200 Heat Load [MJ/m 10−3 0 0 1 2 3 4 5 6 7 8 9 10 11 12 Time [h] Figure 3.4 Thermal stresses for perigee altitude of 60 km and bank angle of 79.5◦ (worst-case) (thus shifting the lift vector downwards) and result in lower load factors and heat flux. At the upper and lower ends of the reentry window, the flight path is more sensitive to the bank angle. At a perigee altitude of 63.5 km, the bank angle leading to a reentry within the constraints is in the range of ∆µ = 8.5◦. This is the favored trade-off since it is a local maximum of the bank angle range at these perigee altitudes. At the upper limit of 71.45 km with a bank angle of 180◦, it is possible to shift to left and right and thus double the range of the bank angle. However, the gradient of the bank angle range is very steep. At 71 km, the range drops to ∆µ = 2◦. Therefore, it is not recommended to enter at this altitude. Fluctuations of the atmosphere or inaccuracies of the predicted trajectory would lead to a reentry outside of the constraints. Finally, with bank angle adjustments during the second pass, a water landing can be guaranteed. Table 3.6 shows at which perigee altitude the bank angle has the largest margin. Results of the analysis can be found in the appendix [36]. Table 3.6 Reentry window (bank angle for first pass, 0◦ for following passes) Perigee Altitude Bank Angle Load Factors Duration Upper window 71.45 km 174.5◦ – 185.5◦ 4.3 – 4.3 g 9.7 – 14 h Selected trajectory 63.5 km 94◦ – 102.5◦ 6.9 – 7.9 g 1.3 – 14 h Lower window 60 km 79.5◦ – 80◦ 8.0 g 12.1 – 13 h

10 4 Human Factors

The primary focus of this section is on the challenges the human health system will face in a microgravity environment. Due to effects of microgravity on the physiological system, solutions are needed to ensure physical health during the whole flight. This includes solutions for the cardiovascular system and bone loss problem in space. To ensure medical care for all potential situations, suitable preventative measures, medication, a systematic training, nutrition protocol and electronic monitoring have been developed. Besides physical health, the focus is also on the mental fitness of the astronauts. Mental health is as necessary and essential for a successful mission as medical care. The concept is outlined in Figure 4.1. 2 Ensure physical health The crew is prepared for all medical risks and is provided with medical treatment possibilities. 3 E-Health 24/7 monitoring and documentation of medical parameters through a health vest and Microflow.

1 Preselecting & Preparation Criteria for the preselection (age, experience, profession,..) are set up. Moreover, the astronauts have to be prepared mentally and physically. 4 Training & Food To prevent muscle degradation due to microgravity, training equipment and a suitable nutritional protocol is provided.

5 Ensure mental health Mental health is ensured by using audio-visual stimulation, a motivation and entertainment kit as well as a daily schedule. Figure 4.1 Health care concept

4.1 Astronaut Selection

The crew consists of a male and a female astronaut, which have both successfully completed their astronaut training. In January 2016, the crew selection process will start to pick one primary and one backup crew for the mission. Preparations and training procedures, which are especially developed for this mission, should start 18 months before the launch. At this time, the crew selection process will be finished and the astronauts will start to specialize their training routine. Mental disorders such as anxiety, post-traumatic stress, insomnia or depression can develop unexpectedly. A study has indicated that the average age of onset depression for healthy people is 41 years [55], so the age of selected astronauts will range from 26 to 37 years. Some crucial factors for astronaut selection are listed below.

ˆ Age ≤ 37 years ˆ University degree (or equivalent) ˆ Free from any disease, any dependency ˆ No intensified radiation exposure until on drugs, alcohol or tobacco launch

11 4 Human Factors 4.2 Crew Health

ˆ Normal range of motion and functional- ˆ Curiosity ity in all joints ˆ Ability to trust ˆ Visual acuity in both eyes ˆ Creativity / resourcefulness ˆ Free from any psychiatric disorders ˆ Resiliency ˆ Blood pressure below 140/90 ˆ Adaptability ˆ Standing height 1.57 m ≤ h ≤ 1.91 m

4.2 Crew Health

4.2.1 Physiological Challenges of Microgravity Human physiology adapts to microgravity, which sets off changes to the human body. When considering physiological risks during the mission, some of the major problems occur because of altered blood composition. These pathological changes are well-known and treatable in their appearance on Earth [58]. According to data from previous missions (e.g. Gemini or Skylab), the erythrocyte and blood plasma ratio decreases daily by up to 1% and blood pressure drops by 10 mmHg after three to six months in microgravity. These effects are due to decreasing vascular resistance. On the other hand, the autonomic nervous system reacts to microgravity conditions and performs counter-regulation, which also leads to increased cardiac output [58]. Consequences of microgravity can lead to various diseases. Due to a loss of blood plasma, there are comparatively too many thrombocytes in the blood, which can lead to pulmonary embolisms. With decrease of hematocrit, total blood volume drops and can thereby carry less oxygen. On account of these effects, blood viscosity can fluctuate in both directions. With too few erythrocytes, which carry oxygen through the body, an under-supply of the brain occurs, which inflicts irreversible damage after two minutes, fainting and eventual death. In addition to diseases due to the cardiovascular system, there are numerous others. Some of them are temporary, like the space adaptation syndrome, which are not further discussed. For the problems mentioned above, several mea- sures are taken to control the cardiovascular system Table 4.1 Human factors summary and replace missing blood components. Measures Mass Volume consist of a combination of diet, physical exercise Astronauts 189 kg 0.2 m3 and medication. A specific diet plan with dietary Launch entry suits 40 kg 0.5 m3 supplements controls the iron and vitamin deficit. Personal items & drugs 12 kg 0.3 m3 3 A positive effect on blood pressure and blood circu- Medical & biofeedback 24 kg 1.1 m Training devices 158 kg 1.2 m3 lation can be achieved through a low-sodium diet. 3 Only in the case of acute changes or pathological Total 423 kg 3.3 m concerns should the latter dietary supplements be taken. This is detailed in Section 5.2.4. With a purpose-built training schedule and modified exercises, astronauts counteract the inevitable reduction of bone and muscle mass. Positive side effects of these workouts are maintenance of the cardiovascular system and physical fitness of the crew. Furthermore, it supports the muscle-vein-pump, which helps to ensure oxygen supply. To get used to these exercises, the crew should start purpose-built training at least six months prior to launch. If abnormal changes or problems occur during one of

12 4 Human Factors 4.2 Crew Health

the monitoring sessions, drug treatment to ensure the health of the astronauts should be a last resort. In this case, in addition to dietary supplements, the crew has access to drugs for hypertension, osteoarthritis, arteriosclerosis or anticoagulants to prevent thrombosis (e.g. iron, calcium, magnesium, etc.). Drugs are administered orally in order to avoid wounds. Through this combination, the crew is prepared for all foreseeable eventualities.

4.2.2 Physical Training The musculoskeletal system is mostly unencum- 100 bered, but it disintegrates and loses tone and mass during a space mission. Because of loss of minerals, bones become weaker and trunk and leg strength 90 decreases by 10% to 20% [31]. Figure 4.2 shows an age-related bone mineral density (BMD) loss rate 80 combining men and women. It is based on a simu- lation, which is strongly correlated to the combined 70 osteoporotic bone loss in BMD. Bone Mass Density[%] Therefore, a specially-designed workout concept 60 is used to keep bone and muscle loss within a reason- 10 20 30 40 able limit. After completing cardiovascular training Time in Space [m] with an ergometer, the crew has to practice with a Figure 4.2 Predicted bone loss [48] combination of barbells and bungee cords. By using different resistance, intensity of the workouts can be varied.

4.2.3 Monitoring and E-Health System To monitor crew health and physical condition, important vital parameters are checked several times a day. Because a of lack of erythrocytes and other aforementioned problems, viral infections can be reactivated more easily than on Earth. Without any medical staff on board, the crew has to be able to make medical decisions independently and provide mission control with health data. Therefore, two devices for data collection are planned. One is the LTMS-3 vest by CSEM in coorporation with ESA. The vest takes pulse, measures electrocardiography, blood pressure, respiration rate and body core temperature without affecting the astronauts in a negative way. The crew wears their vests at least three times a day for one hour to guarantee sufficient monitoring. These vests are currently used in clinical tests at Concordia Base in Antarctica and will have at least TRL 8 by 2018 [16]. Additionally, Microflow [11] is employed. This technology analyzes blood in real time and detects infections. It is also able of rapid viral or bacterial identification. Furthermore, Microflow is able to analyze radiation exposure and stress level of the crew by evaluating blood consistency. These measurements are taken once a week. Monitoring data from the LTMS-3 vest and Microflow is sent to Earth frequently. Hence, the medical team is able to notice conspicuous variations sufficiently early.

13 4 Human Factors 4.2 Crew Health

4.2.4 Mental Health - A Risk Reducing Strategy The success of human space flight depends on astro- nauts remaining mentally healthy in order to survive Table 4.2 Daily schedule social isolation and extreme physical environments. Activity Time It is crucial for the crew to remain alert and vigi- Getting up/hygiene 8:00 am lant while operating complex equipment. Therefore, 1. Meal (0.5 h) 8:30 am getting enough sleep is an important factor. Mi- Mental activity (2h) 9:00 am crogravity, noise, vibrations and loss of the natural Medical measurement (0.5 h) 11:00 am Sporting activity (1.5h) 11:30 am day-night cycle make sleeping difficult in space [35]. Furthermore, studies have shown that irregular work 2. Meal (0.5h) 1:00 pm and day schedules, high workloads and varying envi- Medical measurement (0.5 h) 1:30 pm Brainwave entertainment (1h) 2:00 pm ronmental factors have negative effects on sleep and Entertainment (2h) 3:00 pm crew performance [35]. Moreover, carbon dioxide and Sporting activity (1.5h) 5:00 pm radiation are other factors which negatively impact Medical measurement (0.5 h) 6:30 pm neurobehavior and performance. To avoid these neg- 3. Meal (0.5h) 7:00 pm ative effects as much as possible, a well-structured Privacy time (0.5h) 7:30 pm daily schedule is established and depicted in Table 4.2. Video call, scientific work (2h) 8:00 pm This schedule provides the necessary variety of sleep, Hygiene/sleep 10:00 pm work, physical training and privacy. It will ensure a specific rhythm during this long-term mission to prevent shifting of the sleep-wake cycle. A daily routine also prevents muscle atrophy and bone loss. However, after some weeks in space, when everything has settled into a routine, issues of boredom and monotony may appear. To prevent such situations, the astronauts must be given challenging, meaningful tasks [25]. Meaningful work is consistently correlated with psychological benefits, as it is associated with overall well-being [20]. Another issue is offering as much privacy as possible within limited space. The astronauts are allowed to take < 5 kg personal items (e.g. e-book reader, symbolic items) with them. However, privacy has to be provided on a physical as well as on a mental level [7]. To create an environment where they can retreat and be on their own, the astronauts use an audio-visual stimulation kit. Audio-visual stimulation is a kind of brain entertainment, which adapts the brain waves with external impulses. Due to the fact that frequencies of brain waves correlate with specific mind states, a harmonization of brain waves can induce the related mind state. By using audio-visual stimulation, brain wave frequencies from the highest state of consciousness state (beta and gamma waves) up to the state of non-rapid eye movement sleep (delta waves) can be evoked. Therefore, this kit provides a supportive method for inducing sleep, privacy or a higher mental active state [1, 30]. Despite all stressors, the astronauts have to cope with during the mission, given solutions should minimize risk for the occurrence of mental disorders. Nevertheless, human factors, such as physical closeness, communication and family support, also have to be taken into consideration. As it is not possible to establish a live connection due to signal delay, astronauts will be given the possibility to keep a blog and record or receive videos from Earth. The goal is to maintain motivation and significance of the mission [7].

14 5 Spacecraft Design

Obtaining detailed information about the general properties of a spacecraft and its subsystems from manufacturers proved to be a challenge. Manufacturers’ information was largely used for the preliminary decision process and for the estimation of available volume. Finally, all subsystems were designed from scratch for this demanding mission. Although this seems like a large amount of modifications, special care was taken to keep cost, complexity and need for extensive research and development at a minimum.

5.1 Configuration & Structure

The design of a spacecraft used for a mission largely depends on the specific mission re- quirements. In any case, all subsystems have to be integrated into the structure in a way, that allows it to be launched by available rockets and assembled easily in orbit, if required. For a manned mission, the spacecraft also serves as a habitat and has to fulfill stringent requirements. Furthermore, as many subsystems as possible have to be accessible for repairs without extra-vehicular-activities (EVA). Finally, the overall imperative of a safe, simple and cheap solution imposes further constraints on the available choices of crew habitats. During the decision process, key aspects for finding a suitable solution were cost, availability, pressurized and unpressurized volume and mass. After evaluating multiple concepts, more advanced concepts like inflatable habitats with artificial gravity systems (i.e. NASA Nautilus- X) were discarded. Possible advantages of large habitable volume and superior shielding do not outweigh present problems of development time and unproven reliability when compared to the traditional aluminum can design. Remaining spacecraft are compared in Table 5.1 to find a combination of systems that best suit the needs of this mission. Due to availability and promising development, the Dragon by SpaceX is chosen as the command module and combined with different modules. To augment the available volume for the different subsystems and as habitable space, a second element is needed. Options for that include the European Automated Transfer Vehicle (ATV), Orbital Sciences’ Cygnus and an advanced inflatable concept by Bigelow Aerospace. The TRL has been estimated according to the presented mission requirements. The comparison is also based on constraints put upon the structure by other subsystems and mission aspects. Table 5.1 Structural trade-off Dragon + ATV + Cygnus Enh. + Dragon + BA330 Habitable volume 58 m3 37 m3 20 m3 340 m3 Mass (no PL) 33.8 t 19.3 t 26 t 34 t Total cost 440 M$ 378 M$ 280 M$ 140 M$ + n.a. TRL 6 + 9 6 + 7 6 + 6 6 + 4 Oversized, Fulfills require- Very small Large volume, little infor- Results expensive ments best mation, lowest TRL A trade-off shows that the Cygnus represents the best compromise between extra volume, launch size, mass and cost. Relevant data sheets can be found in the appendix [36].

15 5 Spacecraft Design 5.1 Configuration & Structure

5.1.1 Final Configuration Both Dragon and Cygnus have to be modified to accommodate the systems required for a long-term, manned mission. To make optimal use of the trunk normally attached to the Dragon capsule, it was decided to connect trunk and Cygnus structure via a bulkhead and to perform an Apollo-style docking maneuver in orbit. A launch and deployed configuration is shown in Figure 5.1. Furthermore, the trunk should be at least partially pressurized to allow the crew to move freely between Dragon and Cygnus. It will also support radiators, two of the solar panels and one of the high gain antennas. This requires extensive modification of the trunk structure but benefits in volume and functionality are well worth it and allow for further use of the trunk structure in future deep space missions.

(a) Configuration of spacecraft during launch (b) Deployed spacecraft after TMI Figure 5.1 Structural assembly of Dragon capsule, trunk and Cygnus

5.1.2 System-Layout & Structure The structural mass of the spacecraft parts was determined through simplified calculations as- Table 5.2 Structure system summary suming several worst-case load scenarios (reen- Mass Volume Power try, launch). Thus, a minimum wall thickness Dragon capsule 968 kg - - and mass for several frequently used spacecraft Dragon trunk 414 kg - - materials was determined and the lightest one Cygnus enh. 776 kg - - 3 chosen (aluminum (7075-T73), aluminum hon- Docking adapters 494 kg 0.4 m - 3D-Printer 165 kg 0.3 m3 144 W eycomb). Moreover, the structure subsystem ISPL-Racks 509 kg 0.6 m3 - includes International Standard Payload Racks Tools & accessories 276 kg 1.3 m3 - (ISPL) for storage of consumables. A 3D-Printer Total 3500 kg 2.6 m3 144 W is utilized for the production of spare parts. Up

16 5 Spacecraft Design 5.2 Subsystems Design

to 0.25 m3 of parts can be produced with 127 kg of raw material with 35% fill. Moreover, tools and accessories such as test equipment, fixtures and restraints [31] are included in the budget in Table 5.2. The distribution of the subsystems across the spacecraft accounts for several critical aspects: storage of H2, dinitrogen tetroxide (NTO) and monomethylhydrazine (MMH) in non-pressurized compartments, spatial distribution of critical systems, fairing restrictions, low mass of reentry vehicle, little necessity for outside wiring or piping and current lack of a docking system with an umbilical connection. Furthermore, it acts as a proof of concept that enough living space remains for the astronauts.

Radiators Solar Arrays

ECLSS TCS AOCS Science Human

Rad. AOCS Factors Human ECLSS Factors

Heat Shield Heat

Science

Docking Adapter Docking

ISPL Adapter Docking COMM

TCS Rad. TPS TCS EPS EPS COMM

Solar Arrays Radiators Figure 5.2 Schematic system layout

5.2 Subsystems Design

This section details the design of the subsystems and the assumptions, trade-offs and design choices are presented. Each section is concluded with an overview table of the relevant parameters of the elements. Finally, mass, volume and power budgets as well as the systems engineering approach are given in Section 5.4.

5.2.1 Attitude and Orbit Control System (AOCS) The attitude and orbit control system controls orientation and orbit of the spacecraft by disturbance rejection. This is done cooperatively by the sensor and actuator suite as well as the utilized control algorithms. Identified control modes, disturbance environments and the relevant actuator suite are presented in Table 5.3. Sensor and actuator suites are selected and distributed such that all autonomous spacecraft meet the requirements for their particular control modes. An overview is given in Table 5.4. While attitude determination is conducted solely with on-board sensors, orbit determination depends on support from Mission Control Center. Doppler orbitography and radiopositioning

17 5 Spacecraft Design 5.2 Subsystems Design

Table 5.3 Control modes and disturbance environments Control Mode Selected Type of AOCS Major Disturbance Orbit insertion Launch vehicle controls Stabilizing the PM Gravity-gradient stabilization, thrusters Gravity-gradient Transposition & docking Thrusters Dragon Gravity-gradient Orbit adjustment MTV Thrusters Cygnus Gravity-gradient Docking PM & MTV Momentum wheels, thrusters Gravity-gradient Stabilizing in Earth orbit Gravity-gradient stabilization, momentum Gravity-gradient wheels, resistojets Deep space flight Momentum wheels, resistojets Solar, internal Mars orbit control Momentum wheels, resistojets Gravity-gradient Orbit & reentry control Thrusters Dragon Aerodynamic drag with DSN and Sun sensors is conducted as well as GPS positioning in Earth orbit. The attitude of the spacecraft is obtained by a combination of inertial measurement units aligned by star trackers and GPS. For assembly in LEO, thrusters and resistojets are utilized, while after the rendezvous maneuver momentum wheels are added to the actuators. Resistojets operate with backup gas for low thrust operations. However, after TMI, attitude is solely controlled by momentum wheels desaturated with waste gas from the ECLSS (0.86 kg of CO2 and 0.43 kg of CH4) by using resistojets. From 4450 N m s up to 11 120 N m s of angular momentum, depending on the axis, can be desaturated daily. To save electrical energy, resistojets can operate as cold gas thrusters as well. By using resistojets in combination with momentum wheels, small thrust impulses are feasible while the dimensions of the momentum wheels enable the spacecraft to rotate fast enough in case of a SPE. For synchronization in Earth orbit, rendezvous maneuver and orbit control, 400 N m bipropellant thrusters are provided using MMH and NTO. On the basis of estimated ∆v needed for synchronization, rendezvous, or- Table 5.4 Attitude & orbit control system summary Power bit control and reentry the propellant Mass Volume type is chosen regarding its required Average Peak mass. Bipropellant MMH/NTO is Actuator suite 325 kg 0.27 m3 420 W 2295 W used due to mass savings compared Sensor suite 43 kg 0.08 m3 76 W 76 W 3 with a monopropellant system. For or- Docking system 127 kg 0.28 m 84 W - 126 W Propellant tanks 184 kg 1.93 m3 -- bit synchronization, a trade-off is done Utilities 25 kg 0.96 m3 54 W 90 W between synchronization time and re- Propellant 1166 kg 0.0 m3 -- quired fuel mass due to propellant boil- Total 1871 kg 3.53 m3 634 W 2551 W off of the PM. A sufficient amount of propellant is provided to complete docking in seven hours. For assembling the spacecraft in Earth orbit, additional systems and sensors are necessary. GPS provides the information needed for a rough approach of the MTV towards the PM. Docking maneuvers are conducted with a combination of a laser-reflector system (RVS-3000) with video support. Redundancy is ensured by a radio telemetry system (KURS). For the rendezvous maneuver of the MTV and PM, both units are equipped with the those systems for redundancy reasons.

18 5 Spacecraft Design 5.2 Subsystems Design

Algorithm MPC gained popularity as a method for feedback control due to its ability to optimize closed-loop performance of plants while constraints on inputs, internal states, rates of change and outputs are taken into account. The availability of optimization solvers, i.e. Quadratic Programming solvers [29], and sufficient computational performance, dictated by time step intervals, enables the application of MPC in these fields. MPC explicitly takes future time steps and the impact of the input on the system into account and, therefore, it is superior to the backup PID controllers. As a result of the optimality of the input signal for future time steps, propellant mass and electrical energy can be saved. This can be achieved with the same sensor and actuator suite; only the computer systems and algorithms have to be changed. In fact, improved capability of the control algorithm enables the use of an actuator suite with lower nominal performance, though without losing performance of the overall control system. The input cost saving for a generic example is shown in Figure 5.3. The MPC controller needs only 44.8% of control effort compared with the PID controller without a resulting permanent control deviation. 0.4 0.2 αMPC MPC βMPC PID 0.3 γMPC 0.15 αPID β 0.2 PID 0.1 γPID

Angle [rad] 0.1 0.05 Control effort [N m] 0 0 0 5 10 15 20 25 30 0 5 10 15 20 25 30 Time [s] Time [s] (a) Euler angles step response (b) Accumulated control effort Figure 5.3 Comparison of MPC and PID controller with step response of α = 0.1 rad, β = 0.2 rad, γ = 0.3 rad and similar reaction time As for the Curiosity Rover, reliability of control software and reduction of fatal errors can be achieved by using a development technique presented in [21] which is based on three principles: implementation of risk-based coding rules, using tool-based code review and a logic model-checking tool to formally verify mission-critical code segments. By making the control code accessible from Earth, it can be debugged during the flight which results in higher reliability. A more detailed approach is described in the appendix [36].

5.2.2 Communication System For long duration missions, communication is vital for the psychological well-being of the astronauts. This leads to challenging requirements concerning coverage, reliability and availability of the communication system. The onboard system consists of two redundant high-gain parabolic reflectors that can be aimed individually and two low-gain backup antennas. They are fed by four redundant transponder systems, with 70 W output power

19 5 Spacecraft Design 5.2 Subsystems Design

each. If required, power can be saved by deactivating channels separately. An overview over the key characteristics of the system is shown in Table 5.5 based on link budget calculations [22, 32]. A robust and delay-tolerant transmission technique with forward error correction should also be considered [32, 36]. Table 5.5 Characteristics of X-band nominal operation with high-gain antenna at 1.43 AU [36] Transmit Center Transmit Max. System Bit Error Link Antenna Frequency PowerSpace Loss Data Rate Margin Rate Up ø 15 m 7.200 GHz 2 kW −272.7 dB 54 kbps 3.5 dB 10−5 Down ø 2 m 8.425 GHz 4 × 70 kW −274.1 dB 4 × 18 kbps 3.5 dB 10−5 Due to limitations of the ground stations (see below) X- and S-band are chosen for trans- mission. Frequencies should be allocated early in the development phase as the classification as a Space Research Vehicle is to be decided by ITU and critical for subsequent development [23]. As the use of large antennas is very expensive and limited, facilities with a maximum antenna size of 15 m are chosen for nominal operation and baseline ground segment, providing the required permanent communication. This includes standard communication with mission control as well as high priority emergency communication. 15 m antennas are also much more common than the higher gain 35 m and 70 m classes, which is favorable regarding failure safety. Most of the reviewed systems only support X- and S-band and rarely K-band communication [43]. Even though X- and S-band provide less bandwidth, they are less susceptible to rain attenuation [32]. Depending on space agency participation, the higher availability of 15 m antennas has most likely the biggest impact on ground segment choices. Apart from nominal telemetry, tracking & control (TT&C) com- Table 5.6 Major required data transmissions munication, there are two other ma- Subject Data & Data Rate Timing jor data sources. A summary of the TT&C 10 kpbs Continuous most important data to be trans- Video messages 100 MB at avail. rate Every second day mitted can be found in Table 5.6. Medical monitoring 6 MB at avail. rate Three times a day Every other day, video messages are exchanged. For a video of about 20 minutes length and 360p resolution this takes about 14 minutes at 1 Mbps. Additionally, three times a day the medical data is transmitted to Earth for analysis (see Section 4.2). Figure 5.4 depicts the theoretically available downlink data rates with mostly all channels used. They are modeled using the transformed link budget equation [36]. At the beginning and the end of the mission the data rate is limited at sufficient 1 Mbps to visualize the possible power saving channel adjustments. After the fly-by at Mars, the data rate drops to a minimum. Over a time of roughly 150 days the data can only be transmitted with less than 200 kbps to 15 m ground stations. This can be mitigated by increasing the transmission time, reducing the video quality, or using ground stations with larger antennas for high data rate transmissions as exist in the DSN or ESTRACK for example. In Figure 5.5 the different mission phases are depicted from the communication point of view. At the beginning (phase 1) low delay communication is possible for a few days. In the first part of the mission the delay is between ten seconds and about three minutes (phase 2). One critical point prior to Mars is reached when Sun, Earth and spacecraft are in conjunction (point 3). The uplink will contain increased thermal noise, due to the strong radiance of the Sun.

20 5 Spacecraft Design 5.2 Subsystems Design

1 1 s]

/ 0.8 0.8

0.6 0.6

0.4 Data rate 15 m 0.4 Data rate 34 m Distance [AU] Data rate0 [MBit .2 Distance Data rate 70 m 0.2

0 0 0 50 100 150 200 250 300 350 400 450 500 Time [d] Figure 5.4 Theoretically available downlink data rates with different ground station antenna ø

After that, the most important part of the trajectory Earth is reached, the Mars flyby (point 4). There will be a high demand for communication, but also some Spacecraft obstacles. However, the time delay at that point will Trajectory be at about 3.5 minutes, and there will be a blackout Sun while Mars is between the spacecraft and Earth. By using relay satellites of the Mars Exploration Joint Initiative communication during flyby could be provided. In this case, it would also be possible to transmit higher quality video data to Earth. For this purpose an Electra UHF system is utilized aboard the spacecraft [54]. Another option is to record the mission data and send it to Earth retrospectively. Relay Satellite After the flyby, the distance between the spacecraft Mars (at flyby) Transmission path and Earth will rise to a maximum. At the farthest Figure 5.5 Communication phases point (point 5) the transmission delay is at about eight minutes and the spacecraft is about 1.43 AU away from Earth. The available downlink data rate decreases to about 72 kbps for 15 m antennas. From this point on the conditions improve until low delay communication is possible again before reentry. An element summary is given in Table 5.7. Further details on the communication system, architecture and calculations can be found in the appendix [36]. Table 5.7 Communication system summary Power Mass Volume Average Peak Antenna subsystem 35 kg 2.0 m3 3 W 17 W Electra UHF subsystem 12 kg 0.01 m3 19 W 74 W TT&C subsystem 7 kg 0.1 m3 11 W 148 W Payload & harness 96 kg 0.3 m3 122 W 798 W Total 149 kg 2.4 m3 154 W 1036 W

21 5 Spacecraft Design 5.2 Subsystems Design

5.2.3 Electrical Power System (EPS) The electrical power system is responsible for power generation, storage and distribution. The critical case is identified as the flyby at Mars due to the lowest solar flux and moderate degradation of solar cells. Solar arrays cover average power and charge batteries, which cover daily peaks as well as the Mars flyby and reentry phase. The average power of the different subsystems is summed up while the peak power of certain elements (like oven, waste compactor and high gain antenna) is distributed through out the day as can be seen in Figure 5.6 and is detailed in the appendix [36]. 6000 Capacity [W h] Average Peak 4000

Power [W] 2000

0 0 2 4 6 8 10 12 14 16 18 20 22 24 Time [h] Figure 5.6 Average and peak power distribution

Solar Arrays During sizing of the solar arrays, several factors such as material constraints, array losses and environmental losses Table 5.8 Solar array/cell losses [31] have been considered [31]. The assumptions are summa- rized in Table 5.8. The solar constant is assumed to be Array resistance 0.958 inversely exponentially proportional to the distance to the Packing fraction 0.85 Tracking loss 0.996 Sun. Additionally, a 10% contingency is added to account Radiation damage 0.976 for unforeseen eventualities. UV darkening 0.997 As a result of the analysis, four arrays with Gallium- Micrometeroid damage 0.994 Arsenide triple junction cells (with an efficiency of 26%) are Contamination 0.98 envisioned. Since the solar arrays are sized for operation Resistance losses 0.98 Distribution losses 0.917 at Mars, deployment of two arrays is sufficient to cover the power requirements in Earth orbit. The required diameter Total Loss 0.69 of the disk-shaped arrays for the current mission is 5 m (downscaled from 6 m with 7 kW). The specific mass of the arrays is estimated with 175 W/kg [52]. ATK’s UltraFlex arrays are chosen due to their lightweight structure combined with high strength and stiffness. They were successfully utilized on the Mars Phoenix Lander in 2008. The arrays are scalable up to 10 kW and a model with 5.5 m has been successfully tested [52]. Moreover, development of the arrays is ongoing as part of Orbital’s Cygnus module and NASA’s MegaFlex program.

22 5 Spacecraft Design 5.2 Subsystems Design

Rechargeable Batteries To cover daily peaks of high-power ele- ments and the reentry maneuver, conven- Table 5.9 Battery assumptions [31] tional batteries as well as fuel cells have NiH2 Li-ion NaS been considered. Regenerative fuel cells Specific energy [Wh/kg] 60 130 132 are disregarded due to the relatively low Specific density [kWh/m3] 40 160 165 TRL [8]. The ECLSS offered the possibility Depth of discharge [%] 80 80 80 to convert water to oxygen and hydrogen Efficiency [%] 96 93 85 through electrolysis, thus making the use of a conventional fuel cell possible. Although, the integration into ECLSS results in an intolerable mass penalty. Different types of conventional batteries like NiH2, Li-ion and NaS are compared with properties presented in Table 5.9. Consequently, Li-ion batteries [3] are chosen due to their high specific energy and current application on on-orbit systems as well as the prospective use on the ISS. Depth of discharge (DoD) and efficiency are assumed to be 80% and 93%, respectively. Life time with almost constant capacity is over 1000 cycles at this DoD [3]. Due to the use of a separate rechargeable battery for reentry with a similar required capacity, redundancy is given. Power management and distribution is considered with a constant factor (150 W/kg) depending on the peak power distributed by EPS. Mass, deployed area, (packed) volume and technology readiness level are summarized in Table 5.10. Table 5.10 Electrical power system summary Mass Area/Volume TRL UltraFlex solar arrays 128 kg 37.6 m2/0.7 m3 6-7 Li-ion batteries (cruise) 66 kg 0.05 m3 8 Li-ion batteries (reentry) 60 kg 0.04 m3 8 Power management & distribution 52 kg - 9 Total 306 kg 0.79 m3 -

5.2.4 Environmental Control and Life Support System (ECLSS) ECLSS supplies the crew with necessities such as atmosphere, food and water, but also takes hygiene, clothes and waste management into consideration. In this matter, the crew safety is directly dependent on the reliability of ECLSS, which does not allow for any single point failure of the system. A flowchart is provided in Figure 5.7 for an overview of ECLSS and the synergies with other systems.

Assumptions for ECLSS Design

The ECLSS design is based on an atmospheric composition of 79% N2 and 21% O2 at a total pressure of 101 325 Pa and a temperature of 295 K. Additionally, atmospheric humidity is controlled by the atmosphere control system (ACS). Atmosphere for a pressurized volume of 52 m3 is ensured and leakage is considered. Daily input and output for ECLSS is summarized in Table 5.11 and described in detail (kg/d/p = kilogram per day per person). For the entire mission duration, dehydrated

23 5 Spacecraft Design 5.2 Subsystems Design

K K

̇ K K K K K K ̇ K

̇ Figure 5.7 Flowchart of Mars18 ECLSS (including synergies with AOCS in green and radiation shielding in red) food is stored in the spacecraft. A regenerative life support system with implementation of a photobioreactor for algae cultivation was not considered due to the low TRL. Food is rehydrated and heated up by the crew. Food lockers, as used on the ISS, are replaced by polyethylene (PE) lockers for mass reduction and radiation shielding. For the mission duration of 500 days, mass and volume of the provision adds up to 1407 kg and 4.85 m3 (including margin). The employed water bags resemble ISS bags but contain additional electrolytes and minerals essential for the crew. Since volume and energy is limited, showers are not possible. Therefore, crew members use different types Table 5.11 Mass assumptions in kg/d/p [19, 18] of wipes (including wet, dry, detergent and disinfectant wipes) for hygienic purposes. Crew members shave Food supply (solids) 0.064 their hair regularly and wear wigs, if desired. Edible Food supply (water) 0.248 Food supply (packaging) 0.270 tooth paste is used for dental hygiene. A waterless Food supply (PE-lockers) 0.119 toilet, similar to that on the ISS is installed (toilet Req. water (rehydration) 1.225 paper included with the wipes). Clothes are supplied Req. water (total) 7.819 and dispensed in the waste compactor after usage since Wipes 0.206 no laundry is installed due to low TRL. Cotton clothes Clothes 0.343 Feces & urine 2.096 are preferred to an advanced clothing concept for higher General trash 1.060 comfort. General trash as well as urine and feces require processing/storage. Additional 62 kg of miscellaneous supplies (including duct tape, latex gloves etc.) are needed during the mission [18].

24 5 Spacecraft Design 5.2 Subsystems Design

ECLSS Concept and Design For the life support system, different con- 10000 cepts are evaluated and compared. First priority is reliability and readiness of the Open System Closed System (MF+CVD) system by launch in 2018. This reduces 8000 the technology options greatly. Increasing Closed System (VPCAR) safety and reducing equivalent system mass (ESM) must be achieved with flight-proven 6000 technologies. For the ESM calculations, mass, volume, power in- and output as well 4000 as crew time are considered [33]. For an optimal trade-off, synergies with other sub- Equivalent system mass [kg] systems are formed. The two most promis- 2000 ing concepts only differ in the waste water 100 200 300 400 500 (WW) treatment. A concept corresponding Time [d] to the ISS water treatment with Multifil- Figure 5.8 Trade-off between different concepts tration [61] (MF) and Vapor Compression Distillation [61] (VCD) as additional urine treatment is compared to one using Vapor Phase Catalytic Ammonia Removal [61] (VPCAR). Both concepts are compared to an open loop concept in Figure 5.8. The VPCAR concept is chosen due to a mass reduction of 944 kg. VPCAR is able to convert different types of waste water into potable water without the need of pre- or post-treatment. The system works discontinuously 101 minutes per day and it can purify up to 134.4 kg/d [61]. In this concept, a time-averaged processing rate of 9.406 kg/d waste water is required. Downsizing would be the better option for minimizing ESM. Since it has to be developed by June 2016 to TRL 8 to ensure sufficient testing, this risk is not taken and an unscaled VPCAR system is used. Four-Bed Molecular Sieves [15, 39] (4BMS), Solid Polymer Water Electrolysis [17] (SPWE), Condensing Heat Exchanger [18] (CHX) and Trace Contaminant Control System [18] (TCCS) are chosen for the air management as these are of TRL 9 and thus ensure a high level of reliability and safety. Other technology options, such as Electrochemical Depolarized Carbon Dioxide Concentrator [39] (EDC) and Solid Amine Water Desorption [15, 39] (SAWD) are excluded due to low TRL. Furthermore, SAWD has a potential hazard of producing toxic vapor by amine degradation. A Sabatier reactor is employed for chemical reactions of CO2 with H2 to H2O and CH4. From recovered H2O, O2 can be reused again. Monitoring, atmosphere control and leakage compensation is guaranteed by ACS [18] as already employed on the ISS. Many state of the art life support technologies are not designed for a crew of two. Hence, 4BMS and Sabatier are scaled to optimize ESM, with 20% of the mass fixed and thus not scaled. The remaining part is scaled linearly. During manned missions, following waste components are produced [18]: General trash (e.g. packaging material, clothes, hygiene wipes, food waste), feces with toilet paper, urine, waste water and brine. Urine and waste water are recycled by VPCAR. Feces are stored. A waste compactor (PMWC [24]) is used to compress trash and brine for volume reduction and water extraction. Compressed trash tiles are used for radiation shielding and extracted water is stored together with feces to support radiation shielding. Recovered water is not reused since its quality cannot be guaranteed.

25 5 Spacecraft Design 5.2 Subsystems Design

Table 5.12 Mass flux balance. Positive values mark system output, negative values system input. Values are in kg/d, rounded to two decimal figures. H2O Rad is water passed to radiation shielding tank. CO H O 2 N O H CH 2 WW Stored Air2 2 2 4 Cycle Rad Air Food ------0.50 - - - Crew - 2.00 - -1.67 - - -7.82 - 4.55 - 4BMS 2.04 -2.04 ------SFWE - - - 1.72 0.22 - -1.94 - - - Sabatier -1.18 - - - -0.22 0.43 - - - 0.96 CHX ------4.55 4.55 PWMC ------0.40 - - VPCAR - 0.04 0.01 -0.05 - - 9.21 - - -5.52 Leakage - - -0.03 -0.01 ------Day [kg/d] 0.86 0 -0.02 0 0 0.43 -0.05 0.4 0 0 Mission [kg] 430.99 0 -8.52 0 0 214.18 -23.17 199.53 0 0

Average mass fluxes per day between elements are shown in Table 5.12. The system is designed to balance out the usage and production of H2 and O2. Excess CH4 and CO2 are used by the AOCS and not actually stored. Leakage is assumed to be 0.625 g fluid/m3/d, based on empirical data from the ISS [9].

Storage and Safety Issues ECLSS is designed to provide water and oxy- gen needed during mission after Table 5.12 Table 5.12 Mass flux balance [cont.] and to compensate failure of any technology Urine Brine Waste Feces for two weeks. Storage and initial fluid supply Food - - 0.42 - is sufficient to compensate the supplies from Crew 3.89 - 1.46 0.30 any system as well as being able to store waste PWMC - -0.19 -1.88 - water, hydrogen etc. for two weeks in case of VPCAR -3.89 0.19 - - emergency. Enough lithium hydroxide (LiOH) Day [kg/d] 0 0 0 0.3 cans are provided for emergency carbon diox- Mission [kg] 0 0 0 151 ide filtering for two weeks [15]. In that time, the crew has to repair the affected systems. Additional N2 and O2 are provided to replace the atmosphere in the case of a fire. Gas is stored in high-pressure tanks with the mass estimated per stored fluid mass (based on reference technologies and Barlow’s formula [9]). CO2 and CH4 storage is designed to store produced gases for one day each, before the gases have to be vented by AOCS. Waste water and urine is stored together. Brine is processed in the waste compactor and stored for one day. Water output from the waste compactor together with feces is stored in PE tanks, which also act as radiation shielding. Fresh water is also stored in a separate PE tank. Storage volume and initial supply is shown in Table 5.13. Throughout the spacecraft, IR-cameras are installed for automatic fire detection and localization. In case of a fire, the crew wears masks connected to oxygen supply and extinguishes the fire with portable fire extinguishers. The pressurized volume can be divided in two parts at the connection between the trunk and the Cygnus module through a bulkhead.

26 5 Spacecraft Design 5.2 Subsystems Design

Table 5.13 Storage capacity and initial supply WW, Feces, O N H CO CH H O Brine 2 2 2 2 4 2 Urine Water Capacity [kg] 24.11 33.2 3.01 2.04 0.43 137 186 2.63 355 Supply [kg] 24.11 33.2 0 0 0 137 0 0 0 If the fire cannot be controlled, the crew has to evacuate the affected part and the atmosphere of the damaged section is removed to extinguish the fire. Afterwards, the atmosphere is renewed with the designated gas supply. Therefore, 24.7 kg N2 and 7.5 kg O2 is initially on board to replace the Cygnus atmosphere, which is the largest possible volume to be isolated. Additional CO2 filtering is required to ensure crew safety during reentry. For this purpose a LiOH-filter is employed through which crew expiration is led. About 2 kg of LiOH are needed to reduce the 1.17 kg CO2 exhaled in 14 hours (worst-case) [15]. A mass of 5 kg is estimated for the filtering system.

Simulation with ELISSA The concept presented above is simulated time resolved with the tool ELISSA (Environment for Life-Support Systems Simulation and Analysis) provided by the Institute of Space Systems at the University of Stuttgart. Simulation results prove adequate supply and storage and are shown in the appendix [36]. Table 5.14 Environmental control & life support system summary (technologies marked with * are scaled, shown power consumption is not continuous) Technology Mass [kg] Volume [m3] Power [W] TRL

CO2 Filtration 4BMS* 69 0.26 381 9 O2 Production (electrolysis) SPWE 292 1.79 351 9 Thermal control and humidity sabatier* 23 0.04 0 8 Trace contaminant control system CHX 144 0.60 263 9 CO2 Reduction (get H2O and CH4) TCCS 18 0.07 13 9 Waste water treatment VPCAR 494 1.88 2618 6 Atmosphere control system ACS 169 0.05 32 9 PMWC waste compactor PMWC 71 0.48 630 6 Waste collection system + supplies 122 2.85 0 9 Gas storage and tanks (incl. brine) 156 0.43 0 9 Water suppy 144 0.14 0 Food supply and packaging 1407 4.85 0 Conduction oven and rehydration 38 0.10 1008 9 Clothes 360 2.10 0 Personal hygiene 252 0.37 0 Lithiumhydroxyd cans 51 0.26 0 9 Fire detection system from ISS 10 0.04 0 9 Reentry CO2 filtering 5.25 0.01 0 9 Total 3825 16.3 5294

27 5 Spacecraft Design 5.2 Subsystems Design

5.2.5 Propulsion System The PM is responsible to inject the MTV into the determined free-return trajectory on January 4, 2018. As already mentioned in Chapter 3.2, the PM is closely related to the launch concept. This chapter, however, will focus on the PM only.

Design Process & Requirements The PM needs to be capable of injecting roughly 15 000 kg into the Mars transfer trajectory. Therefore, a total ∆v of 4841 m/s is required. To account for losses due to thrust inaccuracies and the Oberth Effect, a margin of 5% is added to the ∆v. However, the g-load must not exceed the maximum load rated for humans. All upper stages and their respective engines as well as solid and electrical engines were evaluated in-depth to get a comprehensive overview. In order to keep the propellant mass of the PM as low as possible, engines with a high specific impulse (Isp) are compulsory. Thus, solid engines were disregarded and the anal- ysis shows that only a handful of engines actually Table 5.15 Engine trade-off support an adequate Isp for such a mission, as shown Engine Isp [m/s] Upper stage in Table 5.15. The respective data sheets can be RL-10 B2 4566 Centaur- & Delta found in the appendix [36]. The arcjet engine TIH- IV Second-Stage TUS developed by the Institute of Space Systems in Vinci 4561 ESC-B Stuttgart seems promising due to its extremely high J2-X 4395 Earth Departure Stage Isp and thrust for an electric engine of 100 N but TIHTUS 10 000 - is disregarded because of its high electrical energy consumption of 500 kW. The J2-X and Vinci as well as their respective upper stages on the other hand will not be sufficiently tested until 2018. In contrast, the RL-10 B2 proved to be a very reliable engine in a multitude of missions. It is one of the most used upper stage engines and provides the highest Isp due to its extendable nozzle. Additionally, many unconventional concepts have been considered like using cross-feeding between different propulsion stages, using alternative propellants like lithium and fluorine to get a higher Isp or launching empty rockets to use the remaining upper stage as the PM. However, only two concepts prove to be satisfying. The chosen concept is described in the next chapter and the alternative is explained in the appendix [36].

Propulsion Module Concept The final PM consists of two Delta IV 4-m Second-Stages with one RL-10 B2 engine per stage [5]. It uses liquid oxygen (LOX) and liquid hydrogen (LH2) for propellant and offers a 5% ∆v margin. In order for the PM to fit into the Asymmetric Payload Fairing (APLF), the two stages are connected over four hinged telescope beams which hold them in their determined position. In launch configuration, the stages are positioned side by side as shown in Figure 5.9. After the PM reached parking orbit, the hinged telescope beams are extended electrically. Afterwards, the second stage moves backwards until final TMI-configuration is attained. It will then be safely locked in its position. Due to this construction concept, another docking maneuver is not necessary and the mission complexity as well as the fuel mass for the AOCS can be reduced enormously.

28 5 Spacecraft Design 5.2 Subsystems Design

Figure 5.9 Configuration change of PM after launch After the configuration change, the PM turns with an additional mounted magnetorquer into a gravity-gradient stabilized position to reduce the fuel consumption of AOCS until it docks with the MTV. For this docking maneuver, the second stage of the PM is equipped with a modified payload adapter with a reinforced Soyuz probe and drogue docking port, which will connect to the docking port of the Cygnus. The PM, additionally, has an enhanced AOCS with a reflector system, a GPS module and a KURS docking system to be able to perform this rendezvous. Due to the fact that in the worst-case the PM has to stay in the parking orbit for twelve days without the electrical supply of the MTV, it is also equipped with body mounted solar cells that provide the electric energy for the magnetorquer and the avionic. Moreover, as the PM will be in the orbit 14 days before the TMI, the boil-off of the cryogenic propellant of the Table 5.16 Propulsion module PM must also be carefully examined. The propellant 1. Stage 2. Stage tanks are equipped with 50 layers of double aluminized Structure 2850 kg 2850 kg Kapton insulation (MLI) which decreases the boil-off to Propellant 20 410 kg 20 410 kg Add. structure a tolerable level of 1.3% of the total propellant mass for 142 kg 580 kg these 14 days [12]. All these modifications are shown in & EPS AOCS 35 kg 152 kg Table 5.16 and the calculation of the TMI with the above MLI insulation 334 kg 334 kg mentioned worst-case boil-off of 272 kg of each stage is shown in Table 5.17. For more detailed information see Total at TMI 23 771 kg 24 326 kg the appendix [36]. Table 5.17 Propulsion module before TMI (including losses due to boil-off)

∆v Isp Dry Mass Propellant Stage Mass 1. Stage 1774 m/s 4566 m/s 3361 kg 20 138 kg 61 349 kg 2. Stage 3310 m/s 4566 m/s 3916 kg 20 138 kg 37 850 kg Total ∆v 5084 m/s Total Wet Mass 47 553 kg

5.2.6 Radiation Protection Space radiation mainly consists of particle radiation, which is very dangerous for humans if certain dose equivalent rates are exceeded. If primary particle radiation impacts matter, secondary radiation is produced, which can be even worse for humans. There are different sources of radiation in space. Trapped radiation, as it can be found in the Van Allen radiation belts, can be neglected. Also electromagnetic radiation does not cause problems for manned missions. Galactic cosmic rays (GCR) are highly energetic and consist mainly of protons with energies up to a few TeV [59] at low flux densities. GCR are especially critical for manned long-term missions. Because of high energies, GCR cannot be shielded effectively

29 5 Spacecraft Design 5.2 Subsystems Design with available technology. Another radiation danger is caused by solar activity such as flares and coronal mass ejections. SPEs increase the particle flux, mainly protons, significantly at an energy of a few hundred MeV [59]. This kind of radiation must be shielded since it can produce lethal radiation doses in a few hours. Therefore, storm shelters are used. Frequency and strength of SPEs depend on the solar cycle. Since the mission is planned to be during a declining Sun cycle near a solar minimum, only a few SPEs are expected [60]. The strength of SPEs is evaluated based on its spectrum of energy and flux. A strong event can neither be predicted nor guarded against at the present time, but these are rare [50, 51]. For example, during solar cycle 21, no unusually large flare occurred [38]. Nevertheless, the assumption that SPEs do not take place during a solar minimum cannot be made. Alternatives to mass shielding are for example electrostatic shielding, magnetic shielding or biomedical solutions. So far, only concepts and proposals have been found whose power demand and system weight are very high and/or whose TRL are too low to be ready in 2018 without making large investments [45]. For these reasons, the proposed solution for GCR and SPE radiation shielding is the usage of synergistic effects with other systems, smart positioning of actual load, adjustment and replacement of some subsystem material and personal shielding by vests and blankets.

Sensor Systems and Early Warning for SPEs In order to collect data for future missions, as it was done by the Radiation Assessment Detector (RAD) on the Mars Science Laboratory [62], and to record the received doses for aftercare reasons, astronauts are equipped with personal active dosimeters and two radiation sensors are mounted on the spacecraft. This poses a scientific experiment since body dose values are received during a long-term mission which could not be done before. Since passive dosimeters would become saturated during long-term missions, a new gen- eration of active dosimeters, being developed by the German Aerospace Center (DLR) in Cologne, with the goal to fly at the end of 2015 is employed. Furthermore, two Timepix sensors developed at CERN and used on the ISS are mounted on the spacecraft [46]. One is placed on the inside of the Cygnus, which is the preferred habitat for the astronauts. The other one is placed on the outside facing the Sun to detect SPEs. In case of an SPE detected by an increase of X-Rays, the crew is warned and has to move to their storm shelter. For redundancy reasons, the spacecraft also communicates with the satellites GOES and SDO via ground stations, which are capable of announcing SPEs as well.

Proposed Strategies for Dose Equivalent Rate Reduction Dose equivalent rates caused by GCR cannot be shielded Table 5.18 Overview of different effectively without a large increase of system mass. An assembly zones 2 2 increase of the shielding from 10 g/cm to 40 g/cm alu- Area Weight minum reduces the dose equivalent rates for near Earth Heat shield 9.8 g/cm2 interplanetary space only by 2% during a solar minimum 2 Dragon 12.9 g/cm [42]. Therefore, the system is designed to provide adequate Dragon trunk 9 g/cm2 shielding for regular SPEs. To minimize system weight, Cygnus trunk 31.9 g/cm2 only Cygnus is equipped with additional shielding, reduc-

30 5 Spacecraft Design 5.2 Subsystems Design

ing Dragons reentry mass. A mean area weight for the different areas in the assembly is determined, as shown in Figure 5.10 and Table 5.18. Material for the first layer of radiation shielding is assumed to be aluminum.

Figure 5.10 Definition of different shielding zones Since Cygnus becomes the main habitat that is only left for privacy time spent in Dragon (1.5 h/d) and repair work in the Dragon trunk (1 h/d), additional arrangements are made. Cygnus is equipped with two tanks made of PE, which is superior to aluminum for shielding purposes due to its high hydrogen content, lower density and less production of secondary radiation [59, 60, 6]. Each tank covers half of the Cygnus’ outer area. The tank on Cygnus side 1 is empty at the begin of lifetime (BOL) and is filled with feces during the mission and the water recovered by the waste compactor [24]. The tank on Cygnus side 2 contains process water from ECLSS at BOL. In addition, food, wet wipes and disinfectant wipes are arranged on the inside of the capsule to produce a protective curtain whose dose reduction effect was proven in [27, 47]. The waste compactor generates a plastic tile of 0.41 m × 0.41 m per day, which can be used to improve radiation shielding. The development of the shielding area weights are shown in Figure 5.11. 25 20 i i 2 2 20 Structure & systems 15 cm cm / / g g Water from food h h 15 Food & packaging Structure & systems 10 Feces, tiles/water from waste Processed water 10 Total Wet & disinfectant wipes 5 Tiles from waste Area weight Area weight 5 Total

0 0 0 100 200 300 400 500 0 100 200 300 400 500 Time [d] Time [d] (a) Side 1 (b) Side 2 Figure 5.11 Area weight over time of Cygnus module surface area Since SPEs can be predicted neither in occurrence nor in direction, additional measures must be taken, based on the development of the shielding strength. In Figure 5.11 the critical phases can be seen: for Cygnus side 1 it is the end of life (EOL) since the total

31 5 Spacecraft Design 5.2 Subsystems Design

area weight decreases over time and for Cygnus side 2 it is BOL (increasing area weight over time). The thickness of the tank walls, made of polyethylene, is chosen to match the area weight of the TransHab Radiation Shield Water Tank of 5.74 g/cm2 [2]. This results in a wall thickness of 1 mm for Cygnus side 1 and 10 mm for Cygnus side 2. In order to achieve a high shielding/weight ratio, it is effective to implement any shielding as close as possible to the human body. This personal shielding will consist of garments and sleeping covers made from PE, which has proven to be highly effective in absorbing particle radiation. Breathable, sleeveless vests will be made in a shielding quality of 1 g/cm2. This will protect blood-forming organs while not being uncomfortably hot for the astronauts. They should be worn as often as possible, especially while staying in the lesser shielded Dragon capsule. Another measure to increase shielding will be sleeping bags made from PE, again with an area weight of 1 g/cm2. Because of the lack of natural convection, cylindrical bags with a large diameter of 0.5 m will prevent overheating. In both cases the fabric will be made from extruded PE fibers in different weights/diameters depending on application. Suggested fiber weights are 500 dtex for the shell of the sleeping bags and 100 dtex for the fill material as well as the woven fabric for the vests. While these measures do not apply throughout the whole day, it still decreases the statistical radiation dose and should be seen as a bonus on top of the aforementioned spacecraft shielding. When an approaching SPE is detected, the astronauts put on the vests and wrap in their sleeping bags. Until the sensors register the end of the SPE, the astronauts have to stay in the radiation shadow zone between the blue ring in Figure 5.10 and the Cygnus wall. The ring provides radiation protection from less shielded areas than Cygnus (paths through the heat shield, Dragon trunk or Dragon capsule). Due to its chemo-protective effect, amifostine is used for radioprotection at a radiation exposure to achieve a reduction of the hematocrit toxicity. This drug was developed for radiation therapy and may also be helpful to prevent physical damage caused by SPEs. A one time application of 740 mg/m2 body surface seems reasonable [10].

Verification of Shield Design via Simulation To get the equivalent dose rates for the as- tronauts and to prove the radiation shield- Table 5.19 Overview of shielded zones and received dose rates by GCR (mission total) ing concept, simulations with SPENVIS and the Geant4 tool MULASSIS have been Ionizing dose Dose equivalent performed. MULASSIS calculates the re- Cygnus side 1 0.133 Gy 0.532 Sv sulting ionizing doses of a multi-layered Cygnus side 2 0.135 Gy 0.540 Sv shield design. Table 5.19 shows the equiv- Cygnus trunk 0.133 Gy 0.533 Sv alent dose rate for blood-forming organs Dragon trunk 0.154 Gy 0.617 Sv Dragon 0.165 Gy 0.661 Sv behind planar slabs of different composi- Heat shield 0.169 Gy 0.677 Sv tions with a mean quality factor of four as it has been found for a Mars transit for 500 days [62]. The GCR environment is calculated for the mission time frame and doses for Cygnus side 1 and 2 are a mean value for BOL and EOL. Weighted by duration of stay and percentage of slabs on the total surface, this gives a total equivalent dose of 0.562 Sv (detailed calculation can be found in the appendix [36]).

32 5 Spacecraft Design 5.2 Subsystems Design

Maximum dose for a single SPE plus GCR in Table 5.20 Overview of shielded zones and 30 days is 0.25 Sv [60]. Since SPEs can produce received dose rates by a single SPE a high equivalent dose in a very short time, it has to be guaranteed that normal SPEs do not Dose equivalent exceed this dose. A measured SPE from October Cygnus side 1 EOL 0.0254 Sv 24, 1989 (near solar maximum) is simulated with Cygnus side 2 BOL 0.0303 Sv Cygnus trunk 0.0166 Sv the given shielding. Its spectrum can be found Dragon trunk & shelter 0.0328 Sv in the appendix [36]. The doses presented in Table 5.20 are for a SPE duration of one day, the used quality factor is one [62] and values are given for the most critical paths. The design of a mass-based radiation shielding concept for a 500-day, manned Mars mission is exemplified and its effectiveness, at minimal additional mass, is proven by simulations. The dose equivalent rate by GCR for the whole mission is found to be 0.562 Sv and the doses by regular SPEs clearly do not exceed the 30-day equivalent dose rate of 0.25 Sv. A summary of the elements is given in Table 5.21. Table 5.21 Radiation protection summary Mass Volume TRL Radiation sensor (Timepix/Medipix) 0.1 kg 0.044 m3 9 Active dosimeters (DLR K¨oln) 0.6 kg 0.0002 m3 5 PE-blanket & vests (with 1 g/cm2) 138 kg 0.174 m3 - Feces/PMC/processed-water tank 352 kg 0.37 m3 - Shelter-ring 240 kg 0.26 m3 - Total 731 kg 0.9 m3 -

5.2.7 Thermal Control System (TCS) The Thermal Control System (TCS) is responsible for maintaining the cabin and all electronic devices at an appropriate temperature. The sizing is achieved with a simplified energy balance formula [40]. Nevertheless, all heat sources are considered, such as solar radiation Qs, albedo radiation Qa, IR radiation Qir, and dissipation heat Qdiss:

Qs + Qa + Qir + Qdiss − Qout = 0. (5.1)

The emission term Qout is defined by the thermal control system. It is designed to reject the critical heat load amount of 11 430 W encountered in Earth orbit. Other considered cases include the crossing of Venus’ orbit and the dark side of Mars. The TCS consists of two main loops and is depicted in Figure 5.12. The external thermal control system (ETCS) transports heat load from the heat exchanger to the radiators by the means of a water-glycol mixture. The second loop is the internal thermal control system (ITCS), which has a fully-independent fluid loop and uses water as working fluid. Since no EVA is planned, reserve pumps are installed inside the cabin in case of an emergency. Consequently, a non-toxic mixture of water and propylene-glycol [19] is chosen as a working fluid, which has a low freezing point (225 K). Furthermore, MLI of aluminized polytetrafluoroethylene is utilized on areas facing the Sun. White paint is used on the remaining structure to further reduce radiator size.

33 5 Spacecraft Design 5.2 Subsystems Design

Radiator Radiator

ETCS Radiator Radiator

Pumps Pumps (redundant) Heat Exchanger (redundant)

Environmental ITCS Temperature Subsystems Waste Sources Control Electronics Heat

Figure 5.12 Flow chart of the thermal control system In order to increase flexibility and redundancy, four deployable carbon-carbon radiators with an area of each 7.5 m2 have been envisioned. For a cold-case situation, e.g. on the dark side of Mars, only two radiators are active, while two stay redundant on standby. Additionally, the temperature of the working fluid is adapted by the mass flow rate to avoid freezing. Carbon-carbon radiators improve the overall mass by reducing 50% mass towards common aluminum radiators [4]. The increased development cost is included in the budget. During the prolonged reentry due to the aerobraking maneuver, TCS has to be active. Before the separation, approximately 137 kg water from the Cygnus is pumped into the Dragon capsule. The total water quantity of 221 kg is used as a heat sink [31] and all heat generated in Earth orbit and residual heat from reentry is absorbed and released as vapor into the atmosphere. An overview of TCS is given in Table 5.22. Table 5.22 Thermal control system summary Mass Area/Volume TRL Deployable radiators 232 kg 34 m2/3.7 m3 7 MLI insulation 58 kg - 9 Pumps, valves, heat exchanger, etc. 153 kg 0.62 m3 9 Working fluid (water & glycol) 97 kg 0.09 m3 9 Heat sink reentry 88 kg 0.08 m3 9 Total 628 kg 0.79 m3 -

5.2.8 Thermal Protection System (TPS) The worst-case scenario occurs for a reentry with three passes as opposed to the nominal reentry trajectory with two passes presented in Section 3.3. The maximal thermal stresses, as shown in Figure 3.4, occur for a trajectory with a perigee altitude of 60 km. The computed heat 2 2 flux peak is qmax = 1878 W/cm and the total heat load is Qtot = 1165 MJ/m . Furthermore, the TPS is responsible for a soft-landing of the capsule.

Heat Shield To protect the spacecraft during reentry, the improved phenolic impregnated carbon ablator (PICA-X) is chosen, which is a further development of NASAs PICA material. It is designed

34 5 Spacecraft Design 5.3 Scientific Payload

for reuse, making it suitable for multiple passes. Furthermore, it is ten times less expensive in production than the original material. Due to proprietary issues, actual values about PICA-X are not available. A conservative estimate for the vaporization heat is 26.516 MJ/kg which is equal to the vaporization heat of the Apollo heat shield material Avcoat [26]. Since it is a fairly old material, it is assumed to be a conservative worst-case of what a modern ablator like PICA-X is capable of. To estimate the thickness of the heat shield, a virtual ablation analysis was performed [26]. 86% − 88% of the total heat load occur at a temperature above the ablation temperature of the heat shield material [26]. With this assumption and the vaporization heat, the thickness of the ablated layer can be calculated, which ended up being 162 mm. Using the heat shield area and the density of the material, the mass can be determined. The heat shield has to be protected against impacts of micrometeorites, therefore an aluminum-cover with a jettison-mechanism is mounted on top.

Soft-Landing System For a mass assumption of the parachute system, a design table [31] was used with a sink rate and Table 5.23 Thermal protection system sum- mary the total mass of the return capsule. The sink rate is chosen to be 11 m/s; similar to the Apollo Mass Volume TRL missions [57]. For the water landing, the system Heat shield & cover 594 kg - 7 needs to provide a stable floating position. This Parachute system 156 kg 0.4 m3 9 will ensure that the crew can open the hatch Overwater recovery 65 kg 0.2 m3 9 and leave the capsule. The mass of the overwater Total 815 kg 0.6 m3 - recovery system is estimated with an extrapolated assumption [37]. Due to the capability of changing the bank angle after the first pass of the aerobraking maneuver, the landing site can be determined. An overview of the elements of TPS is given in Table 5.23. Further simulation results are presented in the appendix [36].

5.3 Scientific Payload

Additional scientific missions do not only result in a greater benefit for the mission but will also help to keep the astronauts busy. They could serve as something the crew feels in control of and thus stabilize their mental state. Therefore a number of experiments was deviced that can be implemented with little effort but still produce interesting results. It is difficult to actually estimate space, mass and power requirements of all these experiments. However, they were chosen as to not exceed a total of 300 kg (i.e. total habitat mass below 15 t), require a small volume and use little electric power (thus, covered by 20% system margin). Communication: After the succesful test of a lunar laser downlink, communication over a distance of 1.43 AU during this mission would be a good opportunity to qualify the system for deep space communication and allow future missions to send and receive much larger amounts of data. Based on the LADEE spacecraft the payload weighs ∼ 30 kg and requires an additional ∼ 50 − 140 W of electrical power. Since the Laser Communications Relay is scheduled for 2017 it can be expected that the TRL of this technology will be sufficient.

35 5 Spacecraft Design 5.4 Systems Engineering & Budgets

Biological: Investigation of algae growth in deep space could further the use of algae as a regenerative ressource for future missions (production of oxygen and food). The most interesting parameters are algea growth rates and response to radiation as well as signs of mutation. Since algae can be cultivated in completely enclosed tanks there are no interfaces to the actual ECLSS required and measurements can be made directly at the experimental device. More precisely a so called panel reactor which is already being researched could be used. The size of these devices is highly scalable and the only additional electrical power required is for a water pump and artificial lighting. Medical: With the already issued medical equipment a number of experiments can be conducted over the course of the mission. The words in bracktets are the names of similar experiments aboard the ISS. Cardiac activity during exercises (Cardio ODNT) with monitoring equipment, efficiency of drugs in micro gravity (Farma) and changes in the hematokrit (Gematologija) using Microflow and chromosomal aberations in blood lymphocytes back on Earth. Furthermore, the results of active dosimeters in combination with measured blood values indicate the actual dosis the astronauts were personally exposed to. While some of these experiments have been conducted before, this mission presents a completely new frame to study long term effects of the deep space environment on the human body resulting in invaluable information for more effective counter measures. Social: Analyzing the personal diaries of the crew will give insight into mental effects of such long term missions including actual isolation from the rest of mankind. Therefore, the participants have to give an evaluation team the permission to access their diaries after the mission, maybe under certain conditions regarding the publishing of private information. Optical: Although other missions already exensively cartographed Mars the astronauts should have access to a camera and telescope system as this might be a big motivational factor. Thus, pictures of the flyby can be taken as well as during the mission of motives the crew deems interesting. An interesting candidate, unless it proves to be too heavy, together with this mission’s high gain system would be the ISS tested high resolution RAL space camera.

5.4 Systems Engineering & Budgets

During design of the spacecraft, a systems engineering approach is applied [31]. From a mission statement, top-level and subsequently system- and subsystem-level requirements are derived. Trade-offs between different concepts are judged based on mass, cost, complexity and technology readiness. Element margins are applied based on the following: 5% for off-the-shelf items with only minor modifications, 10% for major modifications and 20% for new developments and drastically modified elements. Furthermore, a 20% system margin is applied to mass, power and volume of the MTV. Table 5.24 provides budgets for the MTV, PM and mission total. The volume is divided into pressurized, unpressurized and packed volume. The latter describes elements such as solar arrays, radiators and other elements which are deployed after launch but have to fit the launcher fairing. Power is differentiated based on the average and peak power as well as waste heat. The total mass of the MTV shortly after TMI is roughly 14.5 t. The total wet mass of the system is approximately 63 t.

36 5 Spacecraft Design 5.4 Systems Engineering & Budgets

Table 5.24 Mars18 mission summary by subsystem Volume [m3] Power [W] Mass [kg] Press. Unpress. Pack. Sum Average Peak Waste Structure 3602 2.2 0.4 0.0 1.3 0 144 0 AOCS 1871 1.2 2.3 0.0 3.5 634 2551 144 Communications 149 0.4 0.0 2.0 2.4 154 1036 519 EPS 306 0.0 0.1 0.7 0.8 0 0 398 ECLSS 3825 15.9 0.4 0.0 16.3 1474 5294 1819 Radiation protection 731 0.8 0.0 0.0 0.9 2 2 0 TCS 628 0.8 0.1 3.7 4.6 422 422 156 Human factors 423 3.3 0.0 0.0 4.6 367 378 535 TPS 815 0.0 0.6 0.0 0.4 0 0 0 Scientific payload 300 0.0 0.0 0.0 0.0 0 0 0 MTV Total 11 899 24.9 4.0 6.4 35 2696 7452 3572 MTV Total + Margin 14 988 29.5 4.8 7.6 42 3236 8943 4286 Propulsion module [kg] 48 097 Mission total [kg] 63 085 Moreover, a budget per vehicle is available in Table 5.25 to determine the respective mass and volume distribution. The propulsion module is excluded from this budget, although it can be regarded as an autonomous spacecraft. Moreover, propellant for the docking maneuvers is excluded as well. The living space is evaluated with NASA guidelines for manned missions. Minimal living space at the tolerable limit is approximated with 5 m3 per person [31]. The total available living space at TMI results in 15.7 m3, which increases in correlation to the mission duration and is well above the minimal required 10 m3. Table 5.25 Mars transfer vehicle summary by spacecraft (before TMI) Dragon Cygnus Capsule Trunk Hab Trunk Mass [kg] Vol. [m3] Mass [kg] Vol. [m3] Mass [kg] Vol. [m3] Vol. [m3] Structure 1215 0.2 579 0.3 1534 0.6 0.2 AOCS 664 1.8 0.0 0.0 457 0.6 1.1 EPS 78 0.0 147 0.1 81 0.0 0.3 TPS 736 0.6 - - - - - Communications 114 0.4 35 2.0 0 0.0 0.0 Radiation protection 0 0.0 0.0 0.0 730 0.8 0.0 ECLSS 5 0.0 230 1.0 3590 15.3 0.0 TCS 98 0.09 254 0.4 275 0.4 0.0 Human factors 219 0.6 13 0.2 467 3.8 0.0 Scientific payload 300 0.0 0.0 0.0 0 0.0 0.0 Total 3562 4.0 1259 3.8 7132 21.6 1.6 Total + margin 4274 4.8 1510 4.6 8559 25.9 2.0

37 6 Programmatic Issues

In addition to the technical challenges, a mission of this scope requires thorough cost estimation, planning and scheduling. In the following, these points are discussed. The chapter concludes with a risk assessment including mitigation techniques.

6.1 Cost

Nowadays, spacecraft engineering is mostly driven by the influence of budgets. Back in the 1960s and 1970s with governments providing substantial budgets, cost was not as important as it is today. The paradigm shift came in the early 1990s. A new business model in spaceflight history became prominent: the public-private ownership. Private companies are mostly focused on their return on investment. Governmental organizations, such as NASA and ESA, try to reduce costs by outsourcing parts of their development and production structures. These points make cost estimation very important but also very challenging. Most of the common parametric cost estimating tools are driven by weight. Nevertheless, there are also other factors such as mission difficulty. Despite that, the biggest problem is access to these cost models. Cost databases such as CADRe from NASA are not available to third persons. Others, such as PRICE-H, are software-based and not open source. Notwithstanding, the following cost estimation is based on proven methods that will give a detailed breakdown of the mission cost for Mars18.

6.1.1 Introduction of Applied Cost Estimating Tools Cost estimating tools can be subdivided into two main categories. The first category is called the commercial off-the-shelf (COTS) method. The second is called government off-the-shelf (GOTS) method. Among other, COTS includes [56]:

ˆ TransCost Model ˆ Unmanned Space Vehicle Cost Model (USCM) ˆ Small Satellite Cost Model (SSCM) ˆ PRICE-H ˆ SEER-H.

Among other methods, GOTS estimating tools include:

ˆ NASA/Air Force Cost Model (NAFCOM) ˆ Aerospace Launch Vehicle Cost Model (LVCM) ˆ Advanced Missions Cost Model (AMCM).

The cost approximation for the Mars18 mission focuses on the TransCost Model, USCM, AMCM and Analogy/Build-up.

38 6 Programmatic Issues 6.1 Cost

TransCost Model is designed to be applied in the early mission phases for cost estimation of the development and production of space vehicle systems. TransCost uses a parametric methodology with cost estimating relationships that are deduced from costs for European and US space vehicles of the last 50 years. Moreover, TransCost supports the calculation of operating costs at the ground during the mission. It will be used to approximate the ground and flight operation cost of Mars18. The estimation is based on the fourth and final edition of TransCost [28]. USCM is a parametric handbook and cost model for estimating costs for unmanned space vehicle missions. The model is based on a NASA, military and commercial satellite database. USCM facilitates the calculation of recurring and non-recurring mission costs. The model is mostly driven by weight and was established by the US Air Force because of its possibility to estimate the costs of communication subsystems [56]. Therefore, USCM will be used to calculate, among other things, the communication subsystem. AMCM is developed by the Johnson Space Center. Compared to other cost models such as USCM, it does not solely focus on weight. Moreover, it includes five additional input parameters: quantity of developed and produced units; specification factor for mission level; year of first operation; level of design heritage (block number) and difficulty [31]. The model allows calculation of the costs of the flight hardware for development and production phase. The model is used to determine costs of those subsystems that could not be calculated by the analogical methods. Analogy/Build-up is used to determine costs on the basis of historical data such as work hours and bills of material. Analogy is applied to adjust or extrapolate data. Estimations on Analogy/Build-up are used for Mars18, when information is available and can be adjusted or extrapolated. This estimation method is mainly based on literature analysis.

6.1.2 Overview of Estimated Costs In the following, cost approximation is presented. A more detailed description of how the costs are composed and cost drivers are shown in the appendix. Table 6.1 specifies the development and production costs for all different subsystems. Table 6.1 Development and production costs using applied cost estimation tools Subsystem Costs [$] Applied cost estimation Share [%] Structure 464,342,590 Analogy/build-up 10.8 AOCS 318,128,143 AMCM 7.4 EPS 15,960,936 Analogy/build-up & USCM 0.4 Propulsion 684,244,933 Analogy/build-up 15.9 TPS 5,423,893 Analogy/build-up 0.1 Communication 39,604,088 USCM 0.9 Radiation 3,004,707 Analogy/build-up 0.1 ECLSS 1,652,027,918 AMCM 38.5 TCS 715,027,857 AMCM 16.6 Human factors 36,027,900 Analogy/build-up 0.8 Sum 3,933,792,966 91.6

39 6 Programmatic Issues 6.1 Cost

Cost estimations of the subsystems differ significantly. The ECLSS, for example, causes the highest costs due to high complexity and necessary reliability. Cost of ground and flight operation as shown in Table 6.2 claim only 8.4% of the total costs, which may seem small but results from the knowledge of a definite end. Operation costs for missions with a definite end are less than 20% [31]. Finally, the total costs are given in Table 6.3. Table 6.2 Ground and flight operation costs calculated by TransCost Term Costs [$] Share [%] Prelaunch ground operation cost 21,027,201 0.5 Crewed vehicles mission cost 294,232,708 6.9 Launch, ascent and descent operation cost 16,128,000 0.4 Recovery operation 84,807 0.0 Launch site user fee 347,779 0.0 Commercialization cost assessment 28,800,000 0.7 Sum 360,620,495 8.4

Table 6.3 Total costs of Mars18 mission Term Cost [$] Share [%] Development and production 3,933,792,966 91.6 Ground and flight operation 360,620,495 8.4 Sum 4,294,413,461 100.0

6.1.3 Life Cycle Costs Costs per phase vary from project to project. To portray an overview of cost development throughout the project, a cost diagram is provided in Figure 6.1. Cost trend is displayed by an upward curve which is determined by the accumulated cost throughout the project. Cost distribution of each phase is provided with aid of a bar chart. In this case, costs per phase are calculated on a monthly basis which consists of 21.66 work days.

4000

] 150 $ ]

Monthly $ Total 3000 100 2000

50 Total cost [M Monthly cost [M 1000

0 0 Phase 0/A Phase B Phase C Phase D Phase E Figure 6.1 Life cycle costs of Mars18

40 6 Programmatic Issues 6.2 Roadmap & Schedule

6.2 Roadmap & Schedule

The most critical parts of the mission are technologies which require an increase in TRL and launch preparation and scheduling. A detailed plan to develop all required launch technologies and gain the necessary experience within the given time frame is presented in Table 6.4. The overall imperative for the development of this mission was to use as much flight-rated equipment as possible. Nonetheless, in some cases systems have to be employed that still require a certain amount of research and development to make the whole mission feasible. Care was taken to ensure that only a few of these technologies are currently below an estimated TRL 6, making it realistic that they can be fully developed by the first launch in 2018. The Asymmetric Payload Fairing (APLF), which has a TRL of 4, is used due to its advantages for the mission, which are explained more detailed in the appendix [36]. To ensure that it is verified and tested until 2018, all measures are undertaken and even a test flight is foreseen. In addition, the beginning of development is scheduled as soon as possible, to have an adequate amount of time to solve unexpected problems. One of the biggest problems arises from ECLSS, specifically with the development of VPCAR/PMWC and life-cycle testing. Also, the development and integration of the novel AOCS software will consume a considerable amount of time. Moreover, the heavy modifications on the Dragon trunk and the development of the personal dosimeter through the DLR are crucial elements that have to be completed in time. In the case of VPCAR and PMWC, to achieve the necessary increase in TRL scaling of the technology was foregone. Thus, the resulting machine will be slightly oversized but more likely ready for operation. Usually the complete ECLSS has to be tested with two life-cycles. To stay within the given time frame, it was decided to conduct only one complete life-cycle test. This can be justified with two arguments: first, for the most parts (12 out of 14) proven technologies are used and only slightly modified. Second, should there be modifications necessary during the testing, they will be implemented directly and the results extrapolated. Since MPC has been used extensively in large and complex systems already the adaption to the spacecraft within the given time is possible. Furthermore, as mentioned before, remote access to the code from Earth will allow for in-flight debugging and increased safety should something actually go wrong. Also, the usage of waste products for spacecraft control and a backup PID system can contain an absolute failure of the new software. The overall schedule for the development is presented in 6.2. A more detailed timetable can be found in the appendix [36]. In this copy, only the rough outline and the already mentioned critical parts of the subsystems are presented. Development periods have been estimated based on experience and heritage. It is clear that a launch in 2018 is very ambitious but with proper planning and collaboration between all involved parties it should be possible. The launch manifest in 6.4 has been developed based on current research programs and requirements identified in this report. The entire plan integrates well into existing schedules of agencies and companies. Therefore, crucial scientific/engineering knowledge and flight qualifications with regard to space transportation can be obtained. At the same time, the overall cost is kept low and already arranged plans are not majorly interrupted.

41 6 Programmatic Issues 6.2 Roadmap & Schedule Critical technologies Figure 6.2

42 6 Programmatic Issues 6.3 Risk Management

Table 6.4 Launch system development, integration & manifest 3. Q Contract & order of Asymmetric Fairing development Contract & order for KURS integration

2014 Falcon Heavy testflight 4. Q Search partner for asymmetric fairing launch test Begin of development of asymmetric fairing 1. Q COTS-6 with KURS dock (Dragon) 2. Q COTS-7 with KURS dock (Cygnus) 2015 Rendezvous & docking of Dragon and Cygnus 3. Q Falcon Heavy first launch Cape Canaveral Atlas V551 man-rating - Dreamchaser first launch 4. Q Asymmetric fairing - first test launch 2016 1. Q Atlas V551 man-rating - CST-100 first launch 2. Q Asymmetric fairing - first cargo test with partner

2017 3. Q Integration of payload and launcher Launch PM with Falcon Heavy 4. Q Launch MTV with Atlas V551 Rendezvous & docking 1. Q In-space final system test

2018 Trans-Mars injection

6.3 Risk Management

A manned Mars flyby poses significant challenges and risks. Some of them are referenced by numbers in brackets and can be found in tabelle riskmatrix. One of the most prominent risk in spaceflight is the chance of a single point of failure (SPOF). The risk of a SPOF is present in every technical aspect of the mission. More specifically SPOFs could result in a LOM or LOC of which the latter is unacceptable. Furthermore, a cancellation of the mission due to financial reasons is a possibility (16). This is especially a threat, if total mission costs are gravely underestimated, as it is often the case in such large endeavors. Lack of public attention and/or support could also endanger the entire execution of the mission, i.e. commencement of the mission and continued ground segment operations. Lastly, missions executed by humans are always endangered by human failure.

6.3.1 Technological Risks In general, SPOF can be avoided through redundancies, heritage and testing. This is the most simple approach and has been implemented for all subsystems of the spacecraft. However, there are several facets of the mission that can not be treated that way and some of them will be presented in the following. Table 6.5 shows a more comprehensive summary of risks

43 6 Programmatic Issues 6.3 Risk Management

during the mission. Whereever applicable bracketed numbers in the upcoming paragraphs will reference the content to the table. A complete breakdown of the numbers can be found in the appendix [36]. During TMI a failure of either PM would lead to a LOM and, if not designed carefully, to a LOC during the second propulsive maneuver (7). A LOM is unlikely since the PM is regarded as highly reliable and has been tested in many missions. Before reaching hyperbolic velocities during the second stage firing the crew has ∼ 490 s of time to trigger an emergency shutdown and remain in a highly elliptical orbit around Earth. Thus, a LOC has been avoided and a rescue mission is still possible. Changes in the requirements regarding heat control due to a varying solar constant and albedo radiation have to be compensated for by the TCS. Failure of this system could result in a LOC. As a result, the radiators have been designed to allow for a large operating range regarding excess heat without failure by frozen cooling fluid. Since the spacecraft will return to Earth with a high hyperbolic velocity special precautions have to be taken for reentry. The heat shield has to be in an acceptable condition after being exposed to interplanetary space for the entire mission duration (27). The structure has to be able to withstand the heat and mechanical loads and posses the right aerodynamic properties (18,19). A malfunction in either of those systems would result in a LOC. Therefore, it has been decided to perform an aerocapture and multiple aerobrakes before reaching the ground to reduce the peak heat and mechanical loads. Before performing the Earth-return manoveurs the heat shield is covered by an aluminum shell. To ensure that no LOM or LOC occurs due to failure of minor parts a 3D printing system will be part of the mission (28). Thus, the crew can reproduce and replace small malfunctioning parts. The mission risk and uncertainty is actually very low. In-orbit assembly (8,9), gravity assists (26) and aerobrakes (24,25,28) have been conducted in the past several times without failure. Moreover, the size of the Dragon capsule puts it in the lowest risk class for an aerocapture. Moreover, the risk of failure associated with the total habitat mass is considered to be below 10% based on heritage data [49].

6.3.2 Programmatic risks A big, but often neglected, risk are programmatic issues, such as public support and finances (16,17). While the general opinion is that interest in space exploration has been declining, the experience gathered during this project suggests othwerwise. However, if a focused effort is made, it should be possible to convince large parts of the population of the usefulness (scientific progress, jobs, competition, long term investments) of a manned Mars mission. This assessment stems from the evalutation of the efforts made by Mars18 to gain public attention and support. Without spending any money, the Mars18 website had ≈ 429 visits per month since November 2013. The Facebook page currently has 173 followers and a reach of ≈ 212 per post. While these numbers are certainly not representative they show that people are still fascinated by space exploration and its endless possiblities. To mitigate the danger of a financial overkill in a late project phase a conservative hybrid approach was employed in the cost estimate of this report, utilizing up-to-date information, heritage and cost models. This method ensures that unknown factors are accounted for

44 6 Programmatic Issues 6.3 Risk Management

realistically while not overshooting the known costs. The hope is to keep the actual costs after the program has been completed below the number given in this report. This might be possible through current developments in the commercial space sector. In general, these developments will probably lead to an overall increased launch frequency and decreased costs. Additionally, it is absolutely necessary to make the financial decisions and necessities as transparent to the public as possible to avoid frustration. Thus, it is more likely that the program budget will not be cut severely after the initial planning phase. Furthermore, the financial participation of stakeholders should be considered. For example, selling slots to perform scientific experiments during the mission or broadcasting rights can help to cover some expenses.

6.3.3 The Human Factor Every man-operated system is subject to human failure. This factor can not be neglected in the risk assessment. Several different approaches exist to mitigate this danger: the first and most obvious one is a well-trained and selected team. This is ensured by the selection and training process presented in Section 4.1 (5). A second approach is the implementation of plans that take effect when critical situations emerge, e.g. outbreak of a fire (11,15), short circuits, gas leakage or malfunctions (12,13,22). The risks and their effects have to be rated by timeframe, likelihood, severity (e.g. impact on other system, LOM, LOC) and uncertainty. Furthermore, the ground segment has to have a possiblity of overriding command inputs given to the spacecraft by the astronauts. This could become necessary when unexpected behavior caused by psychological changes or extreme situations make the astronauts choices untrustworthy. Unfortunately there are aspects like very hard SPEs (6), high energetic micro meteorites or space debris impacts (18) that can neither be foreseen nor can their effects be avoided. These events could lead to the death of part or the whole crew. This means that both astronauts have to be aware of the possibility that they might have to deal with a deceased partner during the mission (4). If that was the case, a nontransparent vacuum bag is provided in which the body can be stowed. All mentioned aspects have to be ranked in likelihood and severity to allow for a justifiable decision whether the risk is acceptable. Point (23) represents a case in which the Mars mission will fail. This point had to be included since a realistic view at the endeavor shows that with such a tight schedule failure due to development problems is a possibility. Thus, rigorous project management, more feasability analysis and well worked out schedules are an absolute must for this mission. Table 6.5 Risk matrix very likely likely 16,24 possible 17,26,28 18 23 unlikely 5 14 6,15,27 7,22,25 rare 1,10 4,9,20 2,3,8,11,12,13,19,21 insign. minor moderate major catastrophic

45 7 Conclusion

The foregoing review shows the feasibility of a manned Mars flyby for the year 2018 by presenting a concept that accounts for the following properties: Requirements are identified and implemented. Risks are discussed, rated and minimized. At the same time the total cost is kept as low as possible by using or modifying existing systems and reducing the overall habitat weight. Synergies are employed wherever it was deemed possible to decrease system weight and complexity. Reasonable conservative estimates were assumed throughout the whole mission to account for all avoidable worst case scenarios. The mission has a duration of 501 days and offers space for two astronauts, a man and a woman. A modified Dragon and Cygnus are assembled in LEO with the PM before TMI takes place. The trajectory and several key moments during the mission can also be followed in a video in the appendix on the Mars18 homepage [36]. Returning to Earth, an aerobreak maneuver is performed. The final results of this report also show that a mission to Mars in 2018 is an extremely challenging idea. In order to increase the likelihood of success and decrease the risks without postponing the mission too much a Mars flyby in 2021 could be considered. Thus, there is more time available for testing, development and the qualification of additional heavy-lift launch systems like the SLS.

Acknowledgements It is no small feat to design a manned mission to Mars in such a short time frame even with a large group as ours. With that said, none of this would have been possible, without the support and help from many companies, organizations, societies, institutes and individuals. We would like to thank both Brainlight and Airbus Defense and Space for their professional guidance and support. In addition to contributing their time and knowledge, Astos Solutions also provided software licenses. We also had a great deal of support from the University of Stuttgart community. The importance of the knowledge gained through exchanges with the Institute of Space Systems cannot be overstated. In particular, we would like to sincerely thank Tilman Binder, Emil Nathanson and Christine Hill. We also greatly appreciated the encouragement and interest of our peers from Constellation and the aerospace society DGLR-BG Stuttgart. Finally, it behooves us to recognize and thank a handful of individuals whom critically facilitated our work. We are grateful for the knowledge and assistance from Thomas Berger and Daniel Matthi¨afrom the German Aerospace Centre (DLR) as well as J¨urgenHerholz, Karin Schlottke and Patrick Wang.

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