PART I SPACE IS MORE. Introduction

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3 3 PART I SPACE IS MORE Introduction The purpose of Inspiration Mars is to design and launch Mars flyby mission for an american coulpe. Mentioned nationality is important mainly because of the idea that is behind the whole project. Inspiration Mars Foundation sees space exploration as a catalyst for social development, human prosperity, knowledge, and world globalization. For this reason they are trying to encourage the next space race between the nations. Needless to say that this time it's not about the clash of two superpowers, but the competition based on national pride, spirit of competition and desire for human development. Admittedly, the idea proved to be the most accurate, as evidenced by the number of teams competing in a competition organized by Inspiration Mars Foundation along withthe Mars Society. Before presenting our concept for this competition we have decided to share our point of view on the given subject. It is crucial for understanding why we stressed some issues and why we made certain choices. The first thing we had to determine was what is the true foundation for a successful mission and it bacame clear that it was not Mars flyby. Red Planet has always been the dream for humanity, a romantic idea that continues the cult of pioneers and explorers. This objective serves not only as an inspiration for millions of people who want to witness such a memorable event, but also as a motivation and goal for the pioneers themselves. But it is merly a result.the achievment itself will not contribute to the development of humanity but the way how we reach it, will. The real challange is to keep a man who will live for 500 days in the hostile environment of interplanetary space in a good health. Space is More Our focus on a human needs manifests itself in our team's name Space is More, which is also our motto. It is hard to guess its meaning without putting it in a right context. It is a paraphrase of a saying Less is More. It was said by Mies van der Rohe, who among such names as Le Corbusier, Frank Loyd Wright and Alvar Alto is widely recognised as one of the fathers of modern architecture. Modernism changed our perception of architecture. The mouvment dropped traditional methods of erecting buildings for a favor of open plan spaces, transparency, light construction, and above all it focused on people and their needs. The living space has become more human. Now it's the time for the same thing to happen in the outer space. Present structure of space habitats, such as ISS, is designed for science purposes and the majority of space is designated for science equipment. There is very little space dedicated to humans. Any previous initiative of creating a space dedicated to residential purposes, focused on human needs, for some reasons has not been realized or never gone beyond theory. In our opinion, the mission project Inspiration Mars is the best opportunity to not only face the technological challenges during long space missions but also to learn more about psychological and physical comfort of people spending a long period of time in the state of weightlessness. We think that every designing decision in our report has been investigated from the abovementioned angle. Our report is composed of the starting system project, trajectory, habitat and technologies related to these problems which have been designed with the future crew in mind. Below, there are main dangers awaiting the crew during long interplanetary flights. Human health The so far experiences of people who lived in the state of weightlessness had shown a lot of issues that crews had to face daily to maintain their body and mind in a healthy state. Main factors influencing the crew s condition which appear in most of the sources are problems with accommodation and functioning of the receptors in the state of weightlessness, degeneration of the supporting tissues,

4 4 sleep deprivation, stress, isolation, irregular day and night work plan. Because of technical reasons we have decided to agree to the crew remaining in zero gravity and the main factors for the crew selection were their ability to function well and the ability of their organisms adapt to new conditions. We decided to treat the issue of degeneration of bone and muscle tissue as priority to the crew s health who are going to spend over 500 days in such conditions. Apart from designating some space in the habitat for the exercise equipment in zero gravity, members of our team created a complete construction study of the mechanism that is supposed to stimulate the astronauts bodies by a centrifugal force of 1,2 G. Its description and technical particles can be found in the forth part of the report. Some problems, such as sleep deprivation or irregular plan of work have been solved via character and construction of the mission itself and more on that can be found in the second and third part of the report. Psychological problems connected with stress and isolation have lost on their troublesomeness through the selection a married couple as the crew. Below we can find a psychological analysis of the human needs. Psychology According to Abraham Maslow s theory regarding the hierarchy of needs which allow humans to progress spiritually and self-realize the ones that have to be fulfilled first are the primitive needs (such as: physiological needs including sexual fulfillment, sense of safety and belonging) [1].Needs concerning relationships with others are the basic human needs. Without them higher needs cognitive, esthetic, self-actualization or transcendence cannot be realized. Only covering the need for belongingness ( including the need to bond with people, affiliation and love ) allows further progress [1]. Isolation limits a lot of these needs. The first one is the limitations of freedom of action and possible contacts with other people which regards the outside of the vessel and the inside (a lot of norms regulating behavior on board). In regard to proxemic observation such small space creates problems within itself because humans are not used to remain in such precisely defined place for such a long time [2]. Another difficulty is the deprivation of possibility of sexual contacts. Another is the deprivation of possibility of making decisions (connected especially with physical activity). In addition, the element of lack of security appears. People are social units, as Aristotile finely called it social animals, and they need each other to enrich their emotional world as via interactions with others people can learn about themselves. People need each other to construct their own I which is possible through nonstop social comparisons they conduct throughout their lives [3]. Isolation is caused also by deprivation of human touch. Research indicates a huge meaning of touch as it allows brain to produce chemical substances such as oxytocin and vasopressin which help people to cope with stress [4]. What is more, a lot of research indicates a positive correspondence between touch and creating and maintaining relationships between people. In research concerning isolation scientists often mention discoveries by a Canadian psychologist Donald Hebb, who, in his controversial experiments, analyzed negative influence of lack of sensory stimulation. His team came to a conclusion that long sensory deprivation in its strongest form may lead to distorted perception, thought processes and memory, as well as emergence of anxiety. Taking all the above conditions into account, we can notice that being close with other people is

5 5 extremely important for humans. The problem disappears when a person has their long-term partner with them. Ensuring all their basic needs is vital, they include: physiological needs, acceptance, intimacy and safety, as well as providing them with comfortable conditions allowing sexual activity. Apart from yielding to basic needs of an organism it will also have a huge scientific potential as there has been no research concerning sexual intercourse in a state of weightlessness. Such research is invaluable in future long-term space missions. Another problem is the sense of confinement. An answer to this issue is to create the biggest living space possible, which could be rearranged and modified therefore should help avoid monotony. This way of thinking is reflected in a few paragraphs further under the title Wet workshop. Factors helping people alleviate their stress and control their emotions will be choice of colors which can be either relaxing or stimulating, and a whole range of digitalized cultural goods which will serve entertainment. Crew selection Taking into account the above conditions, we have decided to abandon the idea of the proposed plebiscite to choose the right marriage for the mission. Such journey, despite providing maximum safety and comfort by us, still requires specialized skills and familiarity with the conditions prevailing in aerospace from the crew. Looking for the right couple among lay people involves a risk of sending into space people who may not tolerate the conditions well, and generate huge costs linked to the recruitment and training. Unverified human agent increases the risk of the mission not being successful and has a significant influence on safety as it is unpredictable in a crisis situation. Because of that we have decided to choose our crew from among married couples of astronauts. Right now there are only few married couples among American astronauts who might fulfill the age criteria for the mission. While analyzing candidates we take into account their skills, experience and space mission seniority. According to these criteria the most suitable couple are Karen L. Nyberg and Douglas G. Hurley. Taking into account the nature of a student contest of our work we did not contact abovementioned persons because we treat them as a model couple. Karen L. Nymberg is working as a mechanical engineer for NASA. She has spent over 180 days in space which is an invaluable asset for the mission Inspiration Mars. In 2018 Mrs. Nymberg will be 48 years old which perfectly fills the contest criteria. Her experience, engineering knowledge and trainings she has taken part in make her the best American candidate for a space flight around Mars. Douglas G. Hurley is a test pilot working actively for NASA. He has spent 28 days in space and in 2018 he will be 51 years old. Having an experienced pilot on board is extremely valued. He is able to control all the maneuvers on LEO, make any corrections to the course of the flight and assure a safe return flight of the capsule to Earth. The mentioned couple is an ideal choice as could help to limit training expenses of the people from the plebiscite, but also thanks to their experience they increase the level of safety of the mission. With them in mind we have designed a habitat based on the Wet Workshop idea. Wet Workshop An idea first proposed by Werner von Braun for the mission Skylab in November It assumes using an empty tank from a used 2 nd stage rocket and adjusting it to human needs. A solution that utilizes a structure that normally is disposed of was supposed to cut down the costs and significantly increase the living space. Unfortunately, technological limitations made it in fact too complex and too

6 6 expensive to adjust an empty liquid hydrogen tank on the orbit. One of the factors influencing the costs was the idea of introducing a rigidly fixed research equipment and other devices into the empty tank. It caused an additional mass to be sent into the orbit as well as assembling and logistic difficulties. This task proved to be too complicated to complete on the orbit. Later, also in the mid 1960 s the idea of wet workshop was suggested for a flight around Venus which was in the designing process at the time. The mission itself has never been realized but its project was very similar to the concept of the Inspiration Mars mission that is why it is worth our attention. In the model suggested for Venus the mission was composed of three stages. Stage A - Combining the habitat with an empty 2 nd stage rocket fuel tank and assessing the possibility of using the tank as a long-term habitat. Stage B - Sending the spacecraft into the high Earth orbit, beyond radiation belts and testing it in the conditions similar to the ones found on Venus trajectory. In case of danger astronauts were able to come down to the save area within a few hours. Stage C manned space flight around Venus. These stages were needed especially taking into account technical limitations in that period. Now, in the 21 st century we have opportunities and knowledge to shorten the time needed and to perform many tests before the launch, without the need of testing the spacecraft on the orbit for a year. Additionally, thanks to miniaturization and optimization of the equipment there is no need to design a rigidly fixed equipment within the tank of the 2 nd stage rocket. Instead we believe in clear, adaptable space and in giving the users great possibilities with as simple assembling as possible. All this is to provide the crew with conditions that correspond with our conception of what a comfortable life during mission in deep space should look like, limiting the costs at the same time. In further technical part, we have elaborated on how we had managed to compromise all these issues. PART II LAUNCH SYSTEM ARCHITECTURE AND TRAJECTORY Preliminary evaluation Obviously, the main parameter of importance for defining an interplanetary spacecraft's capabilities is its maximum allowed mass. Very early in the project we made the decision to base it on a relatively short, free-return trajectory with no big DSMs, for a number of technological, political, medical and psychological reasons. When, after this decision, we concentrated on formulating an early estimate of the payload mass value for our craft and turned to current systems for comparisons, we immediately realized they had certain relevant limitations. Currently used launch vehicles have a stated LEO payload mass no higher than approx. 29 metric tons (DIVH, [5]). This corresponds to a Mars free-return payload no higher than 5 metric tons [6] at least an order of magnitude too little for a manned mission. Systems to be introduced in the future improve this figure to about 10 mt for FH [7] and about 20 mt for SLS with an advanced upper stage, this last configuration having the bare minimum capacity required for a manned Mars mission. However, in January 2018 over five decades will have passed since the first on-orbit rendezvous, made by Gemini 6 and 7 on December 15 th Over fifty years will also have elapsed since the first

7 7 successful docking, made on March 16 th 1966, when Gemini 8, under the command of Neil Armstrong, rendezvoused and docked with an unmanned Agena Target Vehicle. By the end of the decade the rendezvous & docking technology was known well enough to be used in critical phases of Apollo flights. In the 90s, the collapse of the Soviet Union and opening of the Russian space program allowed the Space Shuttle to fulfill one of its original roles building and servicing a space station. Both MIR and ISS missions led to a great increase in skill and knowledge within the space program. As of now, NASA and its international partners have mastered extremely complex, long term on-orbit assembly of structures weighting hundreds of tons out of smaller parts and pieces that can be sent to orbit individually, within the current launch space and weight limitations. Additionally, in 2018 humanity will have over half a century of expertise in conducting rendezvous and docking operations. We can build literally every element of a manned interplanetary spacecraft using current launch and on-orbit assembly technology, except for one: propulsion. There is, currently, no known technology for building rocket stages from parts in orbit. Furthermore, we will not have access to a sufficiently developed inorbit propellant depot technology, either now or in reasonably close future. In consequence, the biggest constraint for deep space missions is upper stage capability we may be able to assemble functional payloads weighting hundreds of tons, but, with such a disproportion in terms of assembly versus propulsion and propellant technology, we cannot send them anywhere. Thus, the approach presented in this report is focused on maximizing upper stage rocket capability (and, therefore, throw-weight and S/C capability), within limitations imposed by current technology and the constraints related to running the mission in SLS carrier of the extended Large Upper Stage NASA's Space Launch System with a DUUS/LUS class upper stage would be the most capable launch vehicle that could be available for a 2017/2018 launch opportunity [8]. However, as stated in that source, creating an upper stage fit for an Inspiration Mars mission would require a significant increase in development speed. Given this requirement, we consider the development of a generic multi-purpose rocket/upper stage combo for a manned mission program an architectural mistake, for reasons explained above. The best concept for short and mid-term deep space exploration without in-space depots, including a Mars flyby of this sort, calls for using an SLS rocket as, essentially, a carrier for a big, cryogenic, maximum-performance LH2/LOX upper stage, which, if the mission is planned for the near future, can just be a modified DUUS/LUS with stretched tanks, as presented in this paper. The following section contains a summary of the most important parameters and models relevant to our modified or extended LUS scenario (the question of its payload and the way of transferring that payload to the rendezvous point in this case into LEO is discussed later in the document). However, it must be borne in mind that, as a team based in Central Europe, in a country with virtually no launch-oriented space sector, we are limited practically to publicly available data only. Furthermore, it seems probable that much of the information needed for a professional analysis of such a mission concept is ITAR restricted. In consequence, what follows is as good an analysis of an extended LUS scenario as we could create in our information and tool restricted position. LUS parameters We chose the Boeing Large Upper Stage as the basis for a big upper stage concept, one we then named extended LUS or ext-lus, mainly because its key parameters were made publicly known in the publication The Space Launch System Capabilities with a New Large Upper Stage [2]. A summary of known parameters of a RL10 and a MB60 LUS is given in table 1.

8 8 LUS 4xRL-10 LUS 2xMB-60 LH2 (TLI) 15573,64 kg kg LO2 (TLI) 91105,39 kg kg RCS prop 649,54 kg kg Dry mass 11853,73 kg kg Stage total mass (TLI) ,30 kg kg Usable propellant (TLI base) kg kg Usable propellant (LEO) kg kg Residual propellant 1679,03 kg kg LEO parking orbit m m LEO payload kg kg Tanks propellant load max kg kg Isp 462,5 s 465 s Thrust lbf lbf Table 1. Boeing LUS parameters The information about the 4x RL10 configuration is fairly detailed. The higher performing 2xMB60 is something more of a mystery, while 1x J2X is not presented here at all, due to its poor high-c3 performance. Given the data, we chose RL10 as safe, base option, but also present calculations for MB60, considering that MB60 LH2/LOX liquid rocket engine development is to be undertaken by JAXA in support of the Space Launch System (SLS) Program [9]. Its design is heavily based on finished work, with over 90 % of the components being complete [10]. Scaling from LUS to ext-lus It must be noted that the scale up model is limited in its finesse. That is due, firstly, to the fact that we do not know the exact parameters and behavior of an SLS rocket and, secondly, because even if the data were available, writing a reliable numerical rocket launch simulator would probably be impossible within the project completion deadline. The change in lower stack performance due to change in mass of upper composite is unknown, and would probably be significant, given the SLS's hydrolox core. In consequence of all those complicating variables, we simplified the model by making the ext-lus wet mass fixed at one specific mission point based on the Boeing's LEO mission scenario. The equation is as follows: LUS ext gross mass = Usable prop. (LEO) + Dry mass + RCS prop. + Residual fuel + LEO payload The result is kg for fully fueled 4xRL10 ext-lus. We assume that, given no change in gross mass, the amount of propellant consumed during ascent to LEO is the same as in Boeing's reference, i.e. 63 metric tons. This gives kg mass in LEO. For a 2xMB60 ext-lus, RCS propellant and residual propellant values were assumed to be the same. although, in reality, the second value might be a little lower, due to the reduced number of fuel lines in the thrust structure area. Precise change of dry mass is unknown and educated guess estimates had to be made. Two MB60 engines weight about as much as four RL10 engines (2604 lb vs lb), while a single J2X engine is about two times heavier. However, subtracting 4xRL10 mass from LUS dry mass and adding 1xJ2X mass gives a result 1334 lb higher than the stated J2X LUS dry mass. From there, quick Scientific Wild-Ass Guess was made. Assuming the savings come from eliminating three out of four per-engine stage structure components, one per-engine stage structure component weights about 200 kg. 1 Thus, 2xMB60 LUS dry mass would be about 400 kg lower than in a 4xRL10 LUS. That means kg for fully fueled 2xMB60 ext-lus and kg after ascent to LEO. The next step in the calculations involved scaling the stage up, to accommodate all this mass. We 1 Note: the team is aware that the RL10 and J2X structures are, in reality, quite different this is a simplified calculation assumption.

9 9 assumed that dry mass will scale proportionally to the propellant tank capacity stretch, except for some elements with constant mass (engines compartment, intertank, etc). Since, in either configuration, an engine weighs about 1200 kg and an individual engine structure component weighs about 200 kg, this constant mass was assumed to be 3000 kg for 4xRL10 and 2600 kg for 2xMB60. The original LUS has to carry the same gross mass, just distributed differently it may be assumed that there would be no significant weight penalty from the different distribution that our model does not account for. Residual fuel mass was assumed to scale proportionally to the extension in tank capacity. Values obtained under such assumptions agree with historical data [11]. The RCS fuel value was assumed to be constant, i.e. the same as in a normal LUS, and was treated as an inert mass, unconsumed till the end of the TMI burn, like a ballast. Furthermore, later in the design process of our ext-lus, we added additional hardware to the stage its description and the rationale for its inclusion is presented in later chapters of this document. Here, we provide just a sketch of its mass breakdown for calculation clarity, since we present the final figures for our ext-lus configuration below. The stage modifications mass breakdown is as follows: Docking compartment: Hatch door Airlock + thermal insulation NDS Tank interior accomodations: ECLSS pipes wall attachments g-bike attachment Total stage modifications mass: Table 2. Stage modifications mass breakdown 80 kg 200 kg 370 kg 50 kg 1,5 kg 0,5 kg 702 kg The rest of the mass, up to the limit based on the data point, is fuel. The model in its mathematical form is presented in Equations 1 below. Due to the nature of this set of equations, a python script was written for iterative solving. The results are presented in table 3 and used as the basis for an ext-lus payload analysis. 4xRL10 ext-lus Stage burnout mass: 17408,4 kg 17315,3 kg Residual fuel 2081,4 kg 2130,3 kg Usable prop: 89873,9 kg 93467,0 kg Tank stretch: 23,964 % 26,878 % Total in-leo mass: ,3 kg ,3 kg Table 3. ext-lus stage data 2xMB60 ext-lus {LUS burnout mass = LUS dry mass + RCS fuel + LUS Residual fuel LUS usable LEO prop. = LUS LEO payload + LUS burnout mass LUS burnout mass ext ext LUS usable LEO prop. + ascent prop. + LUS Residual fuel LUS max prop. load ext ext Tank stretch = LUS max prop. load LUS Residual fuel = LUS residual fulel (1 + Tank stretch) ext LUS burnout mass = LUS dry mass + (LUS dry mass mass const.) Tank stretch + Stage mods ext + LUS Residual fuel + RCS fuel ext Equation 1. Scaling-model equations

10 10 Trajectories and ext-lus payload An ideal trajectory for a Mars flyby should have a number of properties, which can, in reality, be mutually exclusive, due to celestial mechanics. The most important ones include: No big DSMs. They not only add risk and complexity to the mission, but also increase deepspace-stack mass in an ineffective manner, requiring it to carry a relatively low Isp propellant (currently, it has to be hipergolic) and expend it without the help of the Oberth effect (excluding a close Mars flyby itself). The shortest flight time possible. Flight time has a significant impact on ECLSS design requirements and parameters (i.e. mass), the crew's psychological and medical safety, as well as the chances of completing the project (political and public perception and support). Minimal required C3. Characteristic energy (C3) is twice the specific orbital energy of an object escaping the Earth gravitational field, therefore it has a direct impact on potential payload mass. Escape trajectory asymptote declination equal to or lower than the launch site latitude, and permitting flight within range constrains. Declinations equal higher than launch site latitude require ascent to orbit with an inclination higher than minimal, thereby causing payload reductions. A launch azimuth exceeding range constrains could make the mission impossible or cause serve payload reductions due to the dogleg maneuver. To comply with IM competition rules, the launch window should lie in After weighting each point and discussing the relevant pros and cons, a fast, free-return trajectory, very similar to the IM reference trajectory, was chosen as the most compatible with the above goals. Therefore, we present, first, an analysis based on the data provided in the Inspiration Mars mission paper and in its sources, and then, based on independent deep-space simulations. ext-lus payload on IM reference trajectory An authorial tool was written in the python programming language for this part of the analysis. Relevant figures were imported from the stage properties modeling module used for an earlier part of the analysis. To properly estimate the payload, a numerical simulation of a burn from 130nm parking orbit was ran, using NumPy and SciPy and a spherical Earth gravity model, to account for gravity losses, since a RL10 or MB60 stage has a relatively low thrust-to-weight ratio. To optimize performance, a twoburn sequence with single coast period in between, on an elliptical, high apogee orbit, was selected. The radiation risk associated with this choice did not noticeably increase the mission's overall radiation budget. Propellant loss due to boil-off was taken into account, but had to be estimated, rather than calculated precisely, due to data availability limitations. Specifically, no exact figures concerning LUS/DUUS predicted in-space boiloff rate were found. Current LH2/LOX upper stages, like the Centaur, are prepared only for missions lasting no longer than a few hours and involving carrying payloads targeted at GEO. Consequently, they have either minimal or no in-space insulation, which results in boil-off of approximately percent of full fuel load per day [12]. If one of those existing stages were modified accordingly, it could reach a 0.X %/d value range and with passive systems it could go down within that range, to as little as 0.1 percent per day. A purpose built stage could further improve this figure by about an order of magnitude [13]. There is no publicly available information about planned LUS insulation. To accommodate crew habitat requirements, we assume an ext-lus would have a modified, habitat-rated insulation similar to the one used in a DUUS. Namely, a dual SOFI+MLI

11 11 insulation, under a MLI aerodynamic ascent shield, as presented in DUUS paper [14]. A DUUS has a total stage life of seven days, but this figure also includes other constrains. Since a DUUS's performance parameters usually match or exceed LUS's parameters and no other information is publicly available on this issue, no special ext-lus estimate was made. The time between an ext-lus's arrival at LEO and TMI is assumed to be four days, which is considered the absolute minimum for in orbit operations in terms of architecture design, and equal to planned LEO loiter time of Constellationera Ares's V EDS [15], designed to perform a very similar mission. In this document only the two most important charts are presented. For purposes of the analysis, the C3 achievable in subsequent days of LEO loitering was plotted, using Matplotlib, against an IM trajectory C3 curve taken form [8]. Maximum payload mass was automatically optimized. The tangential angle of the achievable C3 curve suggests that the boil-off rate should be no greater than 0.3 %/d, and rates lower that 0.2 %/d are not necessary. Thus, a 0.2 %/d boiloff rate was taken as baseline. Plugging boil-off rates of existing stages, however, has a serve impact on the achievable payload. Figure 1. C3 chart for RL-10 based ext-lus For 4xRL10 ext-lus the maximal payload is equal to kg at TMI date (figure 1). The engines fire for seconds to inject the stack into x km transfer orbit with a period of 5.18 hour. Before completing a one revolution the engines fire again for seconds to send the stack to desired departure hyperbole. Other departure dates could be examined as well, for example at TMI date achievable payload is kg.

12 12 Figure 2. C3 chart for MB-60 based ext-lus For 2xMB60 ext-lus the maximal payload is equal to kg at TMI date (figure 2). The engines fire for seconds to inject the stack into x km transfer orbit with a period of 5.20 hour. Before completing a one revolution the engines fire again for seconds to send the stack to desired departure hyperbole. Other departure dates could be examined as well, for example at TMI date achievable payload is kg. Note that in those payload figures, 702 kg of stage modifications are treated as the stage dry mass and not as payload. With stage modifications mass set to zero, the payload for the TMI is kg for a RL10 and kg for a MB60 based ext-lus, respectively. Independent trajectory modeling In this part of the analysis another authorial tool was written and used for trajectory simulations. A trajectory similar to the IM reference one was modeled and other options were also examined, with the aid of PyKEP and PyGMO libraries. A problem, based on modified PyKEP's MGA_1DSM built-in module was constructed and optimized using JDE and NSGA II evolutionary algorithms. Due to MGA_1DSM heritage, a practical total mission ΔV limit of 1 cm s -1 was used to classify a trajectory as essentially free-return. This is two to three orders of magnitude less than total TCM budget of a current one-way unmanned Mars mission [16] [17]. PyKEP's built-in plotting function was used for visualization. For payload mass calculations, LEO loiter time was set to four days. Below, we present the most important results, selected from among many calculation runs. Recreating a 500-day class trajectory Evolutionary algorithm optimization yielded a slightly lower minimal C3 energy requirement with reentry speed cutoff set to m/s. Table 4 presents the minimal C3 free-return trajectory parameters found, figure 3 and 5 shows its visualization. Declination of the outgoing asymptote allows for a launch from KSC to LEO with the lowest inclination possible. Entry velocity is on upper end of Orion ERP capability. The simulation is somewhat sensitive to the minimal Mars distance allowed during swing-by. There is number of inconsistent figures about the planned IM flyby distance, ranging from 100 km, through 100 miles, to closer than ISS is to Earth. For purpose of the minimal C3 evaluation, this distance was set

13 13 to 100 km as in [7]. Computed C3 equals to km 2 s -2 and the corresponding ext-lus payload values are kg for the 4xRL10 version (with a x km parking orbit) and kg for the 2xMB60 version (with a x km parking orbit). For the 2xMB60 based ext-lus with stage modifications mass set to zero and following the presented trajectory, the maximum achievable payload mass peaks a little under 32.5 metric tons ( kg, to be precise). Figure day class trajectory visualization Leg d Earth depart: 2018-Jan-01 16:47: C km 2 s -2 Dec deg Rt Asc deg Leg d Mars flyby: 2018-Aug-19 08:17: Flyby dist km Earth arrival: Table 4. min-c3 trajectory Vinf Entry Vel 2019-May-17 03:20: m/s m/s An example of a free-return trajectory with a larger margin, with flyby distance set to 210 km and C3 set to km 2 s -2 is presented in table 5. Its visualization looks pretty much the same as previous one, therefore it is omitted. Note that in this variant the launch date changes to the December The ext- LUS payload values for this trajectory are kg for the 4xRL10 version and kg for the 2xMB60 version. The parking orbit differs mainly in apogee height, being about 380 km higher in both cases. If a payload trade-off were considered acceptable, with a higher C3 and/or a closer Mars approach, Earth reentry speed could be further reduced. A reduction is also possible if some small deep space maneuvers are added. Table 6 (with figures 4&6) presents a trajectory optimized for Earth reentry speed. The speed was reduced by almost 470 m/s, with C3 boundary set to 45.1 km 2 s -2, a shortest distance to Mars of 100 km and a cumulative DSM magnitude limit set to 10 m/s of dv, a value comparable to the nominal TCM budget of current unmanned Mars missions [16] [17]. The ext-lus payload values are kg for the 4xRL10 version and kg for the 2xMB60 version. Other trajectory options besides the 500-day variant were also studied. They were discarded for the present mission, but two variants are presented in a following subsections as alternative mission options.

14 14 Leg d Earth depart: 2017-Dec-24 12:29: C km 2 s -2 Dec deg Rt Asc deg Leg d Mars flyby 2018-Aug-17 18:49: Flyby dist km Earth arrival Vinf Entry Vel Table 5. Higher margin 500-day class trajectory 2019-May-11 0:55: m/s m/s Leg d Earth depart: 2017-Dec-04 20:43: C km 2 s -2 Dec Rt Asc deg deg DSM after days DSM mag m/s Leg d Mars flyby: 2018-Aug-13 08:07: Flyby dist km DSM after days DSM mag m/s Earth arrival: 2019-Apr-23 10:58: Vinf Entry Vel m/s m/s Table 6. Trajectory optimized for Earth reentry velocity Figure day class trajectory optimized for Earth reentry velocity

15 15 Figure day class trajectory - Mars plane view Figure day class trajectory optimized for Earth reentry velocity - Mars plane view Alternative 730-day class trajectory The parameters of a more classical, two-year free-return flyby trajectory were also investigated, despite the fact that flight length alone would make this an undesirable choice for the present mission. A launch window was found in 2018 (around May). Example trajectory parameters are shown in table 7, figure 7 shows its visualization. Earth depart: 2018-May-09 20:21: Leg 1 C km 2 s d Dec Rt Asc deg deg Leg d Mars flyby: 2018-Aug-14 13:57: Flyby dist km Earth arrival: Vinf Entry Vel Table 7. example 730-day class trajectory 2020-May-3 7:04: m/s m/s It is clear from the data, that this trajectory would also have a few other drawbacks. First, the flight time is not only significantly longer, but also divided unevenly between legs, the second one being six and a half times longer that the first. The ship would arrive at Mars about three months after launch, and then the crew would have to wait 21 months to come home. This is undesirable for both psychological and political reasons. Secondly, the Mars flyby distance is large, usually tens of thousands of kilometers,

16 16 and potentially almost km. At that distance, Mars would be only around 7.5 times larger than the Moon as seen from Earth. Thirdly, a large portion of the second leg is spent further away from the Sun, behind Mars orbit. The last, minor drawback is higher a outgoing asymptote declination, a bit larger than the KSC latitude The trajectory has two main advantages, however. Firstly, a lower required C3 launch energy less than 27.5 km 2 s -2, which allows for a bigger payload. The ext-lus throw-mass is calculated to be kg for a 4xRL10 version (a x km parking orbit) and kg for a 2xMB60 version (268.6 x km parking orbit), with ~0.5 deg parking orbit inclination increase unaccounted for in the calculation a half a degree change should not, however, significantly decrease the payload. The second advantage is a much lower reentry velocity about m/s bringing reentry conditions back into known territory. Figure day class trajectory visualization 2021 Venus-Mars flyby trajectory A double flyby trajectory for 2021 was investigated as a backup, in case of schedule slips, providing a nearly four year time margin, with a late-november 2021 launch window. Example trajectory parameters are given in table 8, while figures 8 show its visualization. The minimal Mars flyby distance was set to 100 km. This family of trajectories is characterized by high departure asymptote declinations, approaching 60 deg (over two times higher than the KSC latitude), which significantly decrease in-leo payload, and may even necessitate a dog-leg maneuver during ascent from KSC. The SLS+LUS payload delivery capability to such orbits is unknown, thus no ext-lus payload calculations were attempted, since the methodology used in earlier sections could not be reapplied here. An educated guess estimate was made instead, based on taking the known payload reduction in other vehicles launching from the Cape to high inclination orbits and applying it, appropriately scaled, to the LUS LEO payload mass. The result indicates that the ext-lus payload mass for a C3=18.5 km 2 s -2 and a i~=60 deg trajectory should, at worst, be comparable to the mass in a 500-day high-c3 trajectory scenario, or higher. Since the Stardust Sample Return Capsule survived Earth reentry at 12.9 km/s [18], the entry velocity of the capsule as calculated for this variant should also be within the currently known manageable range.

17 17 Leg d Earth depart: 2021-Nov-20 07:41: C km 2 s -2 Dec deg Rt Asc deg Leg d Leg d Venus flyby: 2022-Apr-05 22:50: Flyby dist km Mars flyby: 2022-Oct-09 15:05: Flyby dist km Earth arrival: Vinf Entry Vel m/s m/s Table 8. Double flyby trajectory Summary Figure 8. Double flyby trajectory The modification of the upper stage, described at the beginning of this analysis, is pretty simple in principle stretching the propellant tanks to accommodate more propellant. Due to our limited tools and access to data, a non-optimized scaling model with maximum reasonable simplicity was used to scale the stage up to the fixed data point which, as mentioned for example in [6], may not be optimal itself beyond a certain point. more fuel does not results in more payload. A fast, free-return flyby trajectory was chosen, allowing us to reuse IM findings in first part of the ext- LUS payload-trajectory study, giving us apples-to- The ext-lus on top of the SLS carrier rocket, with short aerodynamic shorud used.

18 18 apples C3/reentry speed comparisons in a second part, and useful payload mass comparisons in both parts. Using an ext-lus instead of a LUS/DUUS results in a considerable increase in payload mass for the same C3, even around 50% for the conservative, RL10 based, stage. Independent trajectory simulations further lowered the requirements and, together with MB60 engines, resulted in a payload mass increase of over 60%, up to a whopping 67% (when using both the ext-lus maximum achievable payload and a SLS/DUUS figure from [8]). Thus, a RL10 based stage with an IM reference C3 was chosen as our safe-bet option (resulting in a final payload of kg), and a MB60 based stage with the independently found minimum-c3 trajectory was chosen as SIM-achieved (resulting a final payload of kg). Both variants were used equally in the payload architecture design process, as well as plugged directly into payload mass margin calculations. It can easily be concluded that the payload, although much bigger than in an unmodified SLS+LUS/DUUS configuration, could be launched in one go by a Falcon Heavy, which has a LEO capacity of 53 metric tons [19], and then rendezvous and dock with the loitering stage. With over half a century of experience in conducting rendezvous and docking operations, this should not be considered a disadvantage of the presented concept. If it were a disadvantage, how many centuries would we have to wait for a manned Mars flyby mission? PART III PAYLOAD Falcon 9 Heavy Payload Falcon 9 Heavy will carry to LEO Earth Reentry Pod, Crew, Launch Abort System, modified Cygnus Service Module and Dry Workshop. Falcon 9 Heavy is equiped with custom fairing allowing for mounting LAS system to the ERP capsule. This solution in necessary because of crewed character of the mission. If telemetrics suggest any problem with the Falcon 9 heavy rocket during ascent to low earth orbit the LAS will use it s solid fuel rockets to separate crew in the pod from rest of the spacecraft[p1]. Dry workshop module have docking interfaces on both ends of it. Dry Workshop Dry Workshop is a pressurised container 3167mm long and 4762mm wide. Its construction is simmiliar to Pernament Multipurpose Module Leonardo and weights around 2000kg. Outside the Dry Workshop are tanks with 1,5m 3 liquid oxygen and 4,5m 3 nitrogen, heat radiators, and the antenna. At both ends it is equipped with Nasa Docking System. Internal equipment consists of ECLSS life support system, the hygiene module containing toilet and a shower, a 2,5m 3 water tank, the hardware module and a storage space. The hardware module includes a 3d printer, a server, and a radiation measurement apparatuses. Rest of the Dry Workshop's cargo consist of mission supplies and equipment needed for adaptation of the ext-lus/wet Workshop interior. Mission Supplies: -1.6 tonnes of food supplies -collapsible cardboard cabinets -projector -personal belongings -plexiglas containers filled with regolith -seedlings -reagents -fertilizer portable electronic equipment

19 19 Earth Reentry Pod "orion"erp - is a Earth Reentry Pod that was presented in the official materials for Inspiration Mars mission by Inspiration Mars Foundation. It is a scaled down version of the Orion capsule designed by NASA. Version proposed for the needs of our mission is much smaller than the original and is designed for two people. The diameter of the capsule was reduced by 20% to span equal 4m, what allowed to lower weight by almost 40% and as a result giving weight of 5.5 tons with fuel included. Cygnus Service Module Attached to the ERP Cygnus service module has been modified to the needs of the mission. It is slightly larger than the basic version. The main change is the increased hydrazine tank, and increased solar panels ATK Ultra Flex. For the purpose of the mission is expected to also increase the capacity of the batteries, and calibrate avionics.

20 20 Ext-LUS modifications Ext-LUS had to include further modifications in order to meet Wet Workshop requirements:

21 21 Airlock Concept of wet workshop creates many technical challenges but the most significant modification implemented In the Space is More design is designing airlock with a Hatch. After docking with dry workshop in orbit the astronauts would fill the 370 cubic meters of dry LH tank with atmosphere. Nasa Docking System have necessery pumps and all hydraulics needed for this task but docking interface is not designed to withstand cryogenic environment of the tank during the ascend to the orbit and then waiting there for randevous with ERP. We decided to solve this problem by placing an airlock with hatch on the tank aft bulkhead so it protects the port from liquid helium. The airlock would be 63 inches in diameter so it would be wider than the autside diameter of the port allowing for pumping atmosphere to the airlock. The hatch is the same as on the space shuttle. It can be both manually and automaticly opened and shut. On the tank inner walls are installed thin-walled steel pipes that withstand cryogenic environment. They will be used as installation pipes for water, air conditioning and electricity. They are provided with sealed sockets for given purposes, that can be unsealed once tank is filled with atmosphere via airlock. Inner walls are also provided with snap hooks for ropes. Ropes will serve as communication assistance, and also as a frame for interior structure. There is one special hitch, that is adapter intended for g-bike's localization. All additions are designed with uniform weight distribution in mind and their totall mass sums up to 700kg. Mass Margin The total mass of deep-space-stack minus the ext-lus dry mass was calculated to be no greater than about kg, without the Cygnus SM fuel. Fuel load depends on a required V and precise required V for correction maneuvers is unknown. The propulsion requirements were estimated using IM stack mass from [8], equal to kg, default Cygnus SM max fuel load of 900 kg (from [20]) and Cygnus SM main engine Isp equal to 329 seconds [21]. Two IM-reference mission scenarios are

22 22 investigated, with and without the upper stage. Calculated mass margins are presented in table 9. IM Upper Stage state: Detached Attached IM stack V [m/s] SiM stack mass for above V [kg] SiM mass margin: Safe-bet ext-lus&traj. 79,8% 92,6% SiM achieved ext-lus&traj. 99,1% 113,3% Table 9 SiM mission mass margins PART IV MISSION ARCHITECTURE Docking Manned Mars flyby mission requires launching three different elements into low earth orbit on which on-orbit assembly is performed. Once assembly is completed the main burn takes place and spacecraft is placed on desired trajectory. Three crafts adding up to final spacecraft are to be launched in two rockets. First, SLS rocket which will carry it s modified second stage to LEO and second, Docking timeline Once Falcon 9 Heavy is in orbit the LAS system will jettison from ERP and the capsule with service module will be decoupled from rest of the spacecraft. Once it drifts of to a safe distance the crew will use RCS engines to rotate craft 180 degrees and push themselfes In direction of dry workshop attached to second stage of F9H. Once docking is confirmed second stage tank will be decoupled using explosive charges. At this stage crew waits for the burn which will put them into orbit that passes near ext-lus and during this time they will open NDS door and briefly inspect the state of this part of spacecraft. Once the inicial burn is complete the craft is on collsion orbit and will perform second burn that will put them in a stabile parking orbit near ext-lus. Using RCS in service module of the ERP crew will perform docking to the ext-lus LH tank. After confirmed docking crew will remove palets from docking interface to increase diameter of the port.

23 23 Docking Interface Manned character of the mission creates rigorious technical specification that docking interface have to withstand. Long duration of the mission, no possibility of aborting in case of any problems and mechanical forces that docking ports will experience during main burn and then changing orientation. Extremely high changes of temperature, radiation and hazard of micrometeorites have to be taken into an account. Docking interfaces In use today were not designed for a interplanetary mission but rather for docking on low earth orbit to ISS. The only docking system designed for interplanetary missions is NASA Docking System. First flight using this system is scheduled to This system is technological heritage after canceled Constellation Mission to Mars therefore it is perfect for Inspiration Mars mission. Technical Specification Exhibit extremely low-leakage of less than lbm/day (0.001kg/day) Meet extremely high reliability requirements for man-rated vehicles Accommodate off-nominal conditions including gapping and axial misalignment Exhibit low sealing compression and adhesion forces Withstand long term exposure to the following space environments without excessive damageor loss of sealing ability: Atomic oxygen (AO) Ultraviolet (UV) radiation Ionizing radiation Micro-meteoroid and orbital debris (MMOD) Vacuum conditions Thermal cycling: Operate at temperatures from -50 C to +50 C including thermal gradients Endure long mating periods (~7 months) and repeated docking NDS dimmensions: 57 inches of outsider diameter, 27/32 inches of inside diameter Solar Arrays The most advanced, scalable and cost competitive solar array system is the Next Generation UltraFlex developed for NASA [22]. It has proven it's reliability on the Mars 01-Lander System where first generation of NGU was used. The system is capable of working in deep space environment up to 5 A.U. The reason why using this system for the mission is the scalability of the array area, thereby generated power. It can be scaled to produce from 7 kw of power up to 15 kw and beyond (the system is still under development). We have rough estimations of the power consumption of the systems operating on the spacecraft and NGU is flexible enough to meet even the highest energy requirements during Mars flyby where the spacecraft will be furthest from the Sun and solar array will provide only 50% of power available on a near Earth orbit.

24 24

25 25 Radiation problem (SPE and GCR) Over the duration of a manned space missions, it is necessary to take care of the crew s health. A huge hazard for human life is space radiation, which can be divided into two groups: constant (in approximation) galactic cosmic rays (GCR) and persisting from a few hours to a couple of days Solar Particle Events (SPE) usually strictly correlated with Solar CMEs and Flares. 2 GCR is characterized by low flux and high-energy particles, while SPE has high flux and low-energy particles (protons mainly). Because of their different reactions for shielding (Chart 1), there have to be used two different strategies. As it can be seen on the above charts, effectiveness of shielding materials is several times weaker for GCR in comparison with SPE. There hasn t been developed any structural material more effective than polyethylene, so far. Additionally, mission will be taken during the solar minimum, which means that the crew will be exposed to bigger doses than during solar maximum (example on Chart 2), because the Sun generates Interplanetary Magnetic Field, therefore an effective dose in solar minimum can reach 300% value of the effective dose in solar minimum[9] Chart 1a, 1b. one-day effective dose vs. thickness of two different shielding materials during SPE 1972 and Annual GCR effective dose vs. thickness of two different shielding materials for solar minimum 1977 [23]. GCR shielding strategy Using LH2 tank as habitat shall lead to a strong and negative influence on GCR shielding strategy. However, the team unanimously decided on higher importance of the huge working and living space resulting from using this solution over the disadvantages, one of which is the covering cost of about 2,6 t stage dry-mass for every 1 g/cm 2, which make uncoverable, cause of gaining too much mass for any rocket today and in the near future. Therefore it is vital to concentrate on things like optimized crew selection and estimation of radiation distribution in the whole habitat. Chart 2.. Effective dose vs. thickness of shield material for male in deep space for solar minimum and solar maximum. [9] 2 Van Allen belts issue has been skipped because trajectory models do not provide long enough exposures for highest fluxes of electrons and protons.

26 26 Chart 3. Estimate of GCR radioation in habitat Average man will be exposed to 82 rem effective dose during 500-day mission, which gave us ~1,37% chance for fatal cancer based on BEIR estimates Estimation was drawn by putting 158 grid points on longitudinal section of the habitat, full symmetry was assumed and highly non-symmetric parts (antennas, solar arrays, etc.) were omitted. 24 rays were started from each point, each had assigned 1/24 of the equivalent dose (E) without any shield (based on ICRP model). Next, basing on the equivalent dose vs. depth in material charts, value of each ray was calculated. 24 summed up rays gave a complete weakened equivalent dose (we). Finally, proportion we/e was multiplied by the effective dose without shielding so (we/e)*(97 rem) (Annual dose for deep space GCR in solar minimum is 700 msv without any shielding (ICRP 2007 Q/Wt), so assuming that mission shall last 501 days which equals ~1,373 y it gives us (70 rem /y)*(1,373 y)=(97 rems/mission). SPE shielding strategy. Frequency of SPE occurs much less often in the solar minimum than in the solar maximum, but it s still a huge hazard to the crew because SPEs are a poorly understood occurrence, they are unpredictable and have wide-spectrum huge energies carried by protons in a very short time. Nevertheless, it s still a less problematic issue than GCR. High energy flux has a fixed direction so the shield can be concentrated in a little space of the spacecraft to completely reduce the effective dose to less than 1 rem. The key solution against SPE is the position of the spacecraft towards the Sun, the crew must be separated from it by an amount of constant mass that is as large as possible i.e. a whole stage construct, so the spacecraft will be oriented with Cygnus towards the Sun. Table 10. List of the most dangerous SPEs since Arranged in relation to the value of 30MeV proton flux [9] Chart 4. BFO (blood formed organ) Equivalent dose vs. aluminum shield thickness for three of top10 cases (wide spectrum of influence)

27 27 This solution overcomes almost every SPE, but in the case of the biggest SPEs (Table 1) of which 30MeV proton flux exceeds the value of 10 7 (particles/cm 2 ) there was designed a two-compartment water-waste 3 sliding container. SPEs fluxes have different reactions to shielding and points out the necessity to use a system that can protect the crew from that phenomenon [Chart 4, 5, 6]. Data about all the potentially lethal SPEs is collected by GOES managed by NOAA, so in case of danger the crew would be informed directly by it and could start the procedure of positioning the sliding container into the center of dock-part of the storage and find shelter behind the container. Additionally, spacecraft will be equipped with RAD (radiation assessment detector) adopted for GCR and SPE detecting which would also inform about hazards from the Sun activity. Chart 6. Dose equivalent vs. depth in material for aluminium equivalent group for King SPE [24] Chart 5.Dose equivalent vs. depth in material for polyethylene equivalent group for King SPE 1972 [24] Container will generate 185 g/cm 2 of shielding, completely eliminating the SPE radiation problem. Probability of coercion to use it is small because of the solar minimum and the generally weakened solar activity during recent years. Though, prevention is necessary because of the possibility of ARS and weakly studied effects of small acute doses. *wastes are assumed as complex of carbon based materials. 3 wastes are assumed as complex of carbon based materials.

28 28 Communication An extremely important part of any space mission, especially human flight, is a communication system. We want to send flight parameters to the mission control center on Earth, provide frequent communication between the astronauts and the control center, send some selected science experiment results and first photographs of Mars taken directly by people. Considering aforementioned needs we established several conditions for our system that must be met: At least basic communication must be possible at all times; we would be glad if there were no periods when communication is impossible. The system should be well-proven and reliable. It should work from any distance from Earth from the very beginning of the flight, through the approach to Mars and during the return flight. It should allow to transfer all data we need. We decided to use NASA Deep Space Network and Space Network/Tracking and Data Relay Satellites System. DSN and TDRSS systems Deep Space Network (DSN) is a network of huge antennas located about 120 degrees around our planet providing communication with a spacecraft uninterrupted by its motion. It has been allowing contact with many kinds of space missions exploring Solar System (including Mars) for over 50 years. The picture shows the DSN antennas range. They are located in Goldstone, California, near Madrid, Spain and near Canberra, Australia. Since the distance is about 30,000 km from Earth there is no single point where spacecraft could not connect to the network. However, this kind of problem appears for shorter distances from our planet. Because both launch and returning approach to Earth are the key parts of the mission we decided that another network providing constant communication during these moments will be needed. Space Network/Tracking and Data Relay Satellites System (TDRSS) is commonly used to communicate with objects on Earth orbits (low Earth orbit satellites, International Space Station Hubble Space Telescope, formerly also space shuttles). TDRSS will be an excellent complement which fill the gap in communication near Earth. Using DSN and TDRSS we ensure that our mission has constant contact with Earth. Both systems have been used commonly for long time therefore we can consider them as proven and dependable. Moreover, TDRSS usually operates above 99.9%. Both DSN and TDRSS provide precise tracking of the position of the spacecraft which is their advantage that we will use. Antennas Figure 9. Field of view of the Deep Space Network antennas, looking down from the North Pole.B y LouScheffer (Own work) [CC-BY-SA3.0 (http://creativecommons.org/licenses/bysa/3.0)], Wikimedia Commons To establish connection with TDRSS we will use an antenna which Cygnus spacecraft is equipped with by default. It is a quasi omnidirectional antenna formed using two passively coupled, circularly polarized, hemispherical low gain antennas, one on the zenith and one on the nadir side of the via

29 29 spacecraft [25] It is designed to communicate with TDRSS therefore we do not need any other antenna. For communication with DSN we plan to use a 1.3 meters parabolic dish X-band high gain antenna (HGA) (similar to one that 2001 Mars Odyssey Orbiter is equipped with) and two X-band low gain antennas (LGA) mounted on the HGA dish on the front and the back sides creating a construction similar to one that Mars Reconnaissance Orbiter is equipped with. It will provide communication regardless of position of the spacecraft and point of the trajectory. The HGA needs to be pointed accurately to Earth therefore it will be steered using a precise gimbal mechanism. It is our main antenna which will be used in most cases. The LGAs provide much lower data-rate capability but they do not need to be pointed to the target precisely. They will be used in some special cases when the HGA cannot be used. We also plan to use two (which one is active at a time) Small Deep Space Transponders (SDSTs) and two redundant Solid-State Power Amplifers (SSPAs, output power: 15 W), as they are proven solutions from 2001 Mars Odyssey.[26] [27] Other equipment: According to the telecommunication system from 2001 Mars Odyssey, we must use also: - A 3-dB hybrid coupler, to allow either SDST to drive either SSPA without requiring active switching; - A bandpass filter (BPF) on each of the 2 transmit paths. The BPF filters the SSPA output for increased isolation at Odyssey s receive frequencies; -A waveguide transfer switch (labeled S1 on the block diagram of Fig. 3-1) for transmit path switching. Depending on the S1 position, one SSPA is connected to one transmit antenna, while the other SSPA is connected to the other. - A coaxial transfer switch (labeled S2 in the block diagram). S2 connects one antenna to one SDST, and the other antenna to the other SDST; - A notch filter (NF) on each of the two receive paths. The NF passes signals only at the Odyssey receive frequency band [27] Data rate Through duration of the whole mission the Sun will not lie between Mars and Earth therefore the communication will not be disrupted by it at any time. The chosen trajectory also provides the spacecraft with relatively small distance from Earth through the whole flight. When the spacecraft reaches the orbit of Mars the distance between Earth and Mars will be nearly the smallest possible. As we use an antenna similar to Mars Odyssey's antenna and a powerful amplifier we expect to reach data rates similar to Mars Odyssey's data rate in these conditions 110 thousand bit per second [28]. Second data transmission system Three orbiters revolve around Mars: Mars Reconnaissance Orbiter (MRO), 2001 Mars Odyssey (MO) and Mars Express Orbiter (MEO). They collect data from rovers and send it to Earth using DSN. Communication through the orbiters would be much more efficient than direct communication. Using them, especially MRO, we would reach higher data rate. We considered about using orbiters (mostly Mars Reconnaissance Orbiter, it has the fastest downlink (to the Earth) transfer) to communication, as an extra communication system. It would require more onboard equipment (UHF transmission system), and it could be unprofitable. UHF receive systems, like this on the Mars Reconnaissance Orbiter, were designed to communicate chiefly with landers and rovers. The UHF antennas, that are mounted on the orbiters are directed to the surface of the Mars. The period of time, in which we will be at a sufficient distance to be able to communicate with orbiters in the UHF band, is in our opinion too short, to use second data transmission system in our mission.

30 30 Life Support Studies In a considered first manned Mars mission, the highest priority should be emphasized on the psychical state of the crew. Such a long duration flight, as is presented in this work has not been previously practiced. Long isolation time in zero gravity conditions in limited space can have high impact on the psychical state of the crew members and requires long training [29]. Due to this fact crucial is the selection of people capable to take part in such a mission. Crew selection should be performed from the group of people, which are in good personal relations and can be described as mental stable. In order to keep crew in good mental health there need to create opportunities and provide some duties to fulfill. Instead of all systems maintenance crew have to take part in scientific studies using a unique opportunities to stay in space for so long time. In our project we propose preformation of studies as presented in the following paragraphs. PRODUCTION OF NUTRIENTS Possible occurrence of conflicts is highly possible during long duration flight. This should be understood and regulated during the fly by the studies performed in zero gravity conditions over problems, which solved, could have the great importance on the future manned missions to the Mars surface and the other long duration flights. There should be performed studies concerning the problem of nutrients production over the flight time. There are numbers of studies considering this problem. We would like to limit the discussion only to the problem of algae growth [30]. Intensified studies are being done on algae, this material has a great potential and can become biocatalyst capable to degrade products of metabolism and produce edible proteins. Proposed habitat is equipped with infrastructure adequate to perform studies on biologically originated materials. Due to low available volume of habitat only simple studies could be performed in flyby conditions. More sophisticated studies cannot be done with specialized devices and trained staff, but simple studies can be done in zero gravity conditions. Sealed tanks filled with algae need to have transparent coating to allow the light to penetrate the interior of containers. Devices attached to the tank are sufficient to filtrate cell suspension and living samples to further studies. The studied system need to be isolated from crew members to reduce impact to life support system. In such an attractive conditions there should be performed simple studies like growth yield. But also interesting from life science point of view is the problem of cell polarity and diversification of higher order living structures [31]. This goal is possible to achieve, but the first step to determine is the algae growth in zero gravity conditions. To fulfill these objectives algae samples collected from tank have to be delivered back on Earth in order to perform detailed studies about cells morphology and any possible changes observed due to flight conditions like zero gravity or space radiation. Such an approach has important scientific impact and can become a solution to arising problems. This is the step to be done in order to better understand life in space and on the Mars surface. The other problem that should be studied during the mission are social behavior and psychical condition of crew members. What should help in understanding of all situations is writing and analysis of logbook of the journey. Analysis of human behavior should be correlated with known behavior profiles observed at similar conditions, in comparison to previous manned space missions, long duration ship cruise or imprison behavior. Crew should be provided with psychologist help available online at any time when needed. Keeping the crew in good mental health is essential for the success of the mission. Life support system The next element from which depends the success of the mission is providence of life support system [32]. Safe and precisely tested system need to be applied in described project. Nowadays such system is present and used with great success in International Space Station. Environmental Control and Life Support System (ECLSS) is the most reliable, tested and ready to apply system which can be

31 31 directly used in long duration orbiter. ECLSS was introduced in 2001, designed to provide basic needs of six crew members. Complexity of the system supports all basic human needs to water, air, food and waste reclamation, moreover the system stabilizes the temperature in crew cabin [33]. The most important parameters are monitored online and system optimizes conditions continuously. Due to the short preparation and realization mission time, life support system should be known and validated, another high-end system cannot be performed and optimized in such a short time left to flight. Also systems based on artificial biosphere cannot be applied due to low reproducibility and high volume needs. Figure 10. Schematics showing all connections between sub-systems in ECLSS. ECLSS combines together numbers of sub-systems closed in connected loops [34]. System force matter flow and its renewal on catalyzers. In such a way is regenerated water, oxygen and discarded waste metabolites. ECLSS for Mars mission have to work in closed loop, due to restrictions of available mass and volume taken from Earth. All basic human needs have to be satisfied using this system. Standard ECLSS designed for six crew members has very good performance coming up to 88% for water recovery and 98% of oxygen recovery. This system can be optimized for two crew members, which would result in reduction of needed area mass of the systems. CREW NEEDS All calculation done for ECLSS system were based on assumption of two crew members, having average mass m = 70 kg, and basal metabolic rate equal P = 10.5 MJ/d. In such conditions organism requires intake of oxygen of mean value equal m O2 = 0.88 kg/d, producing at the same time carbon dioxide of mean mass equal m CO2 = 0.98 kg/d. Potable water need is equal m H2O = 4.0 kg/d, to the water consumption have to be added water needed for hygiene purpose, the mean lowest possible mass needed is equal m H2O = 0.75 kg/d, however this can be raised up to kg/d when assuming shower in the system. In this work we assume basal need of crew just to live and perform tasks required for mission success. Dry mass food has been assumed as m food = 1.0 kg/d, however literature data are ranging from 0.5 up to 1.5 kg/d.

32 32 For presented assumptions faeces weight was taken as m = 0.35 kg/d. Mass balance for statistical crew member and total masses are composed in table below. ECLSS DESIGN ECLSS system is composed of a few sub-systems which are fulfilling their roles in closed relay to the other systems, as a whole forming complete and functional devices array. Sub-systems have been divided to small subunits easily to exchange and servicing. There are listed and shortly described basic modules forming ECLSS [35][36][36][37][38]. Water reclamation system is composed of two parts. Urine Processing Assembly (UPA) cycling urine and flush water coming from WHC. It is a first step in water purification. Second system is called Water Processing Assembly (WPA), which process UPA distillate. This system is responsible for iodinated water production and its deliver to oxygen production. Next step is Oxygen Generation System (OGS). The main part of this system is the electrolyser, having a task to form oxygen and hydrogen from water. Hydrogen is next transferred to Sabatier, but oxygen is given to the cabin. Water needed for electrolysis is stored in tanks and possessed from WPA system. System responsible for removal of carbon dioxide Carbon Dioxide Removal Assembly (CDRA) is based on reversible reaction like on activated carbon or sorbents, however CDRA system is also equipped with lithium hydroxide assembly to remove carbon dioxide in irreversible way in state of emergency. To the described system is closely related system of Air Contamination Control (ACC) which in online monitoring contamination of air using mass spectrometer and removing waste and toxic compounds. Ari management is also regulated with pressure control with supported system MCA introducing oxygen to cabin. The following system belonging to the ECLSS is Temperature and Humidity Control (THC). It is responsible for stabilizing temperature in the cabin and forces air flow. Water condensed in the system is directed to WPA. Fire safety is realized by Fire Detection and Suppression (FDS). Crew cabin is equipped with fire extinguish and CO 2 based system. Fire is detected with laser sensors. ECLSS adopted to Mars mission would not, as it is on ISS, be so diversified. It is possible to reduce each components and optimize them to only two crew members. With such modification sub-systems connected to ECLSS can be localized in loading space and distribute air and water into infrastructure present in wet workshop. Modularity is great advance of this system, this allows to design habitat without any problems. Figure 11. Consumption and expulsion for statistical crew member of weight 70 kg and normal basal metabolic rate. Total value represent summed value for whole trip, without recycling. ATMOSPHERIC CONTROL The main function of ECLSS is production and stabilization of atmosphere. The system allows formation of atmosphere corresponding to Earth atmosphere to the composition and pressure in cabin. Atmosphere contains 78% of nitrogen, 21 % of oxygen and 1% of other gasses in normal pressure. Gasses before filling wet workshop are stored in form of liquid oxygen and liquid nitrogen in high pressure tanks outside of the vehicle. After emptying of fuel chamber oxygen and nitrogen are inserted

33 33 into free space heated up with thermal control system up to nominal temperature for the system (25oC). Nitrogen have to be inserted as first to remove hydrogen remaining in tank and to reduce oxidative property of pure oxygen. Oxygen and nitrogen are heated during entering the chamber, this also have to minimize oxidative property of liguid oxygen. Air circulation is made by fans distributing air over the whole wet workshop. Carbon dioxide is removed from the habitat with sub-systems described previously. Heat exchanger are placed as presented in schematics on [figure 11] outside the vehicle. Supply of O 2 is done by electrophoresis. Oxygen recharge is done with rate 7.2 kg/h in standard considered system. To the emergency services is counting fire detection and extinguish system with continuous monitoring of all compounds in air. Air leakage in long term timescale is possible, but pressure reduction turns on air supplementation procedure from main oxygen and nitrogen tanks. Figure 12. Air circulation system in ECLSS connected with heat transfer module. WATER CIRCULATION Water management for ECLSS is based on water stored in tanks and collapsible bags diminishing occupied area. Emptied bags can be refilled from the system or stored on the walls of fuel chamber. Water tank is also forming inner walls in loading space. This water jacket is the main reservoir of potable water in described project. Water is microbiologically controlled with iodine in off-line system. Organic carbon is also monitored in off-line mode, informing about malfunction of waste treatment subsystem. Online monitoring is reduced only for conductivity as a good remark of total contamination. Water management system works with yield calculated as 88%. Volume of water jacket, which is at the same time main water tank is quite high for the purpose of this mission but it is recommended to reduce showers frequency during the flight.

34 34 SUPPLIES AND WASTE Food supplies necessary for crew survive are stored in dry food form. For the food preparation is needed addition of water in specially designed place. Food packing can be recycled with biomass, when using biodegradable materials or stored pinned to cabin walls under masking coating. Waste management is divided to urine treatment and biomass treatment systems. Urine is filtered and recycled for potable water, biomass can be rejected after dehydration. Waste Management System is placed in loading space. Important part of this system is odor and bacteria filter, and vacuum vent for gasses removal. Description of the interior design Issues Designing the utilisation of the empty liquid hydrogen container, the designer had to face the following problems: the limited loading space of Dry Workshop, which was to accommodate the whole future equipment of Wet Workshop mass restrictions that result from the carrying capacity of rockets ease of assembly adaptability dependency of used solutions Taking into account all these requirements, we have decided to design a room divided with rigid, endurable PCV-coated membranes. The rolls of membrane coiled in Dry Workshop take up little space and have a small weight. Inside Wet Workshop the sheets of the material will be fixed to lines that will also serve as communication assistance. This space offers a huge compositional and functional potential. The membranes and lines make up the floor, walls, screens and tops. All this being of a small volume and weight of the package and high durability. The space enclosed by the membranes satisfies all the requirements and provides the crew with the possibility of modifying and adjusting the space so as to suit their needs. It also corresponds to the cultural background of our idea and draws on the creativity of modernist artists as inspired by the works of Mondrian. Before we lay down the functional analysis of our design, we will first describe its assembly. Assembly First Wet Workshop is filled with air and heated up, then it is supposed to be fitted with the equipment that is contained in Dry Workshop. Step one is to unseal electric and ventilating sockets and then to assemble the lighting and target ventilating systems. Only then can the crew go on to do the fitting in the space of the container. The crew must at first fit and tauten the communication lines. At this time the crew moves about the

35 35 place using installation pipes that are in the container. One end of the lines is fitted by means of a carabiner and a hook; the other end is provided with a simple tautening mechanism. The line is fixed to a spring whose other end is to be attached to another hook on the wall. To tauten the line, the crew uses the turnbuckle that must be turned so long till it allows for the spring to be fitted on the hook on the wall. Now the line is taut and the spring keeps it taut during the mission despite the fact that the line is being constantly stretched. The following step is to stretch the main membranes that constitute both the floor and the partition between the living and the study rooms. The fabric used for this purpose is strengthened with polyester yarns. It is especially tear-proof, stable and climate resistant (-30 C to +70 C). A special acrylic paint makes the coated material extremely dirt and decay resistant and therefore extremely lowmaintenance. It is flame resistant in accordance with M2, CI1 and DIN 4102 B1. The fabric (420 g/m²) is ideal for applications on large surfaces. Due to its thickness of only 0.45 mm, the fabric can be easily rolled up and unrolled. When stretched on a tensile structure, it receives rigidity similar to conventionally constructed partitions. Fabric sheets are stored in Dry Workshop in a form of 2.5m wide rolls. The assembly of the membranes is fairly simple and even simpler in zero-g environment. Four segments of the main membrane are stretched on steel ropes tautened with turnbuckles. It should take around two hours to complete this task and after this Wet Workshop will consist of two separate spaces. Accomodaction These are the spaces ready to be used as living rooms and study rooms. First the crew must prepare a place to live in, where they will be able to have a rest after they have completed the stressful beginning of the mission. To this end they stretch additional polypropylene lines and between them they spread the membranes that make walls, partitions and tabletops. In the living space there will be a water dispenser close to the table tops for foodstuffs and for the preparation of food. LIVING QUATERS The furniture designed as a store is made of cardboard. It is light and it takes up little space when unassembled. It will contain stockpiles of food and utensils of everyday use brought from Dry Workshop, and it will also serve as a store of disused packaging and wrappings. The pieces of furniture that will serve as seats and beds will be pneumatic. Although in zero-gravity there is no need for recreational equipment, still it is recommended that the space be filled with things that serve the purpose of having fun and rest, maintaining the cultural and behavioural context. The fact that in the living room there will be utensils and objects that are familiar from an ordinary home on earth will allow the crew to better identify with the new place of living; additionally it can be observed what purpose will those objects serve during the crew s free time. PRIVATE SPACE The most important thing appears to be the inflated bed that will be used for having a rest and a sexual intercourse. This bed is fixed to the lines of the main membrane by means of straps. The mattress is made up of three layers with three segments. The top is provided with straps and buckles for sleeping bags and for fixing the bed so as to immobilise it. If necessary, there can be created a space between two layers of the bed by putting up two cushion layers that face each other. This will create a quadrangle-like space with a width of two and a height of one cushion. This constitutes a soft, separate space for the intimate relationship in weightlessness. The limited space gives support for the body, and

36 36 its elasticity does not hinder body movements. This piece of furniture will provide the crew with the basic comforts of body and mind, but not only this. If people are supposed to travel in space for a long time, they must answer the question of how to have sex in weightlessness. Up to now this problem has not been addressed by scientists. The mission to fly around Mars provides an occasion to do away with this taboo theme, since the lack of sex will count as deficiency in the basic needs of an individual. This is something we have addressed in this project. The observation of this aspect of life carried out by the crew will be discussed in Part Five of this report, where also the research and tasks that the crew will be entrusted with will be presented. After the fitting of the living room has been completed, the crew should have one day off, during which they may make use of the space and provide necessary adjustments to suit their tastes. The space solution proposed by us here has the following advantages. First, it takes into consideration the distribution of the Galactic Cosmic Rays in the container. Hence the living room, where the crew is going to spend a significant amount of their time, is placed at the end of the container, where the amount of GCR to which the crew will be exposed during the flight will be by 13% smaller. This solution let us design the zoned space. Starting with the hatchway from Dry Workshop there is a space accessible to all, the room for daily activities and the canteen. Next we enter the corridor from which the crew members may enter their own, separate rooms. They serve the purpose of providing privacy and separateness, which is of importance when people are forced to live together and share a limited space. The last zone is the sleeping zone, which has the most intimate character. WORKING SPACE The interior design of the space devoted to work ought to begin with vertical partitions that divide it into a study and a gym. Having regard for the health of the crew, it is recommended that the gym be prepared as first, which means the assembly of sports equipment such as human centrifuge, treadmill and expanders. The study will contain tabletops where Plexiglas terrariums for the research into the cultivation of algae will be placed. At the centre of the space below the centrifuge, there will be an atrium where in a cardboard flowerpot filled with soil research into the cultivation of some plants will be carried out during the mission into Deep Space. This will have not only a scientific value, but it will be a positive incentive connected with the contact with nature, and may provide the crew with possible additional foodstuffs. For the description of this research see part Four of the Report.

37 37

38 38

39 39 The centrifuge the G-bike Microgravity negatively affects bone tissue. Experience shows that astronauts lose some 1-3% of bone tissue after a month spent on orbit. Bones decalcify and become weaker, easy to break. In the case of the second space flight presented by our team, the changes that might occur in the bone tissue of the astronauts would be irreversible and so they might be life threatening and thus endangering the whole mission. So we decided to consider creating artificial gravitation. Gravitational force may be replaced with centrifugal force. Unfortunately, it is not all that easy to make the whole habitat rotate; it is much easier to design rotating equipment. It is also worth remembering that astronauts move using hands first of all, so the trainings should target legs. Let us consider a design of such a bike that driven by the astronaut s legs could rotate around its own axis. The gear is to transmit the rotational movement of the pedals onto the movement of the whole centrifuge around its own axis. As the figure shows, the centrifuge is made up of 4 belt-drive cogs and 3 cogwheels. We assume that cogwheels have the same radii. The cogwheel fitted on the axis of the centrifuge is immobile, the other cogwheel moves around it. We make an initial assumption that the acceleration at a distance of 170 cm from the axis of the centrifuge equals 1,2g. This will be centrifugal acceleration. Let us calculate the angular velocity with which the centrifuge must rotate: If we assume that u2 =u1=u we will obtain:

40 40 Additionally, we assume that the cyclist does 80 rotations with the pedals per minute. Thus: So the transmission ratio of both chain drives will equal 1.8. The centrifuge will perform 25.2 rotations around its axis per minute. To this purpose the bearing marked as PRT from Igus might be used. Let us assume that the framework will be made of steel profiles. Let it be a rectangular closed profile with the dimensions 40x20x2.5. The parameters of the profile needed for calculations are: * cross-sectional area * moment of resistance Let us further simplify the equation for such a framework: we ignore the relatively small weight of the profiles (approx kg/m) and we assume that the force is applied at the very end of the lower bar. 1. Bending: 2. Bending

41 41 3. Bending Let us thus apply constructional alloy steel for carburisation Let each connection be made by means of two fitted bolts. We are looking now for the minimal diameter of the bolts: Thus we can accept bolts (d=10mm). Let us check on the clamp on the hole in the sheet metal: Where g gauge of the sheet metal: Where g gauge of the sheet metal: The transmission ratio obtained from the calculations amounts to 1.8. its exact value, as in the case shown in the figure, equals. This solution enables the training of one astronaut who drives the centrifuge. In the event that two people want to take exercise at the same time, then another bike must be fitted to the centrifuge, which, however, will not drive the construction. In order to diversify the exercises derailleur gears could be added.

42 42 In order to limit the feeling of dizziness while performing the exercise it might be helpful to do it blindfolded. Unfortunately, this is by far an unhappy solution. A better solution to prevent dizziness from occurring and thus contribute to a better psychological condition is to use the Oculus Rift google glass. They convey a stereoscope picture, so they could prove good as projectors of virtual cycling trips as Figure 13. Selection of belt-drive cogs. Z number of teeth. carried out on Earth. The current technology allows for ever better immersion of the human being into virtual reality, which might render inestimable help during long missions. Together with the centrifuge it will make up a complete set of stimuli, both physical and psychological. PART IV MISSION This type of enterprise calls for precise schedules. In this case the following three arrangements can be singled out: preparation of the mission performance of the mission daily schedule of the crew activities Preparation of the mission The previous parts presented the kind of technology that will be used and the method of its use. The schedule for the preparation of this technology and the equipment for the 2018 mission will be presented together with the cost estimate made so as to match the time of the completion of the whole. The cost estimate covers the period from now, i.e. from Q to the safe landing of the crew in Our budget is noticable lower than the one presented by Inspiration Mars Foundation. We decided to not include a habitat. Instead we use small starege mosule, Dry Workshop. Also we doesn t plan to use any other spacecraft to deliver our crew to the orbit. Final factor is very low mass that need to be carried into space thanks to utillizating empty LH2 tank for living purpose. We belive it is the most cost efficient, simple, and comfortable way of interplanetary travel.

43 43 Performance of the mission The performance of the mission shows what occurrences may the crew encounter flying along the trajectory, and how their flight will look like. We have divided the mission into the so called key points. These are the turning points of the expedition, between which the crew fulfil their daily schedule. launching r SLS Start D9H Start Docking launching into trajectory development of Wet Workshop daily schedule of crew activities. adjustment of trajectory. daily schedule of crew activities passing Mars daily schedule of crew activities landing Daily schedule of the crew activities Every day the crew is obliged to take exercise and to make experiments. This is supposed to fully employ the mental capabilities of the mission and provide the members of the crew with activities. Research has shown that a specific daily schedule plays an important role for the psychological health of a separated human being. Below is the timetable for daily activities with explanations: 8.00 am am. Getting up 9.00 am am. Training am pm. Research pm. Lunch 1.00 pm pm. Research 4.30 pm pm. Training 6.00 pm Supper 7.00 pm Free Time

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