Inspiration Mars. International Student Design Competition WUT. Dream Team

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1 Inspiration Mars International Student Design Competition WUT Dream Team Emilia Węgrzyn Anna Mizerska Katarzyna Woroniak Tomasz Witkowski Błażej Żyliński Wiktor Krzeszewski

2 CONTENTS 1. Introduction Orbital mechanics Environmental Control and Life Support Systems System selection Composition of the system Air Revitalization Subsystem Atmosphere Control and Supply Subsystem Temperature and Humidity Control Water Recovery and Management Waste Management Fire Detection and Suppression Other System Weight Estimation Launcher selection Potential vehicles Performance analyses Conclusions Spacecraft Design Power Management System Power Consumption Battery Pack Solar panels 32 Power output Air mass coefficient Solar cells efficiency comparison Solar Irradiance in space Hydrogen fuel cell Cost Mission Science Program Summary References Inspiration Mars 2

3 1. Introduction This document covers the NASA's pre-phase A. According to NASA System Engineering Handbook it is purpose: "To produce a broad spectrum of ideas and alternatives for missions from which new programs/projects can be selected. Determine feasibility of desired system, develop mission concepts, draft system-level requirements, identify potential technology needs." Nowadays manned mission to Mars is one of the most important goal of human space exploration in XXI century. Nevertheless, politics of space agencies deflate sense of that because of costs. On the other hand space industry is allocate from national space agencies like NASA to private sector. Private sector can introduce cheaper and more elastic solutions based on prepared ones. During mission design process we should remember about many confines. The most important are two of them: 1. time for design and integration of a system - 4 years only, 2. total cost of mission. In our analysis we try to discus every acceptable solution specified by mission requirements. Unfortunately deadline of launch enables the design of new solutions. First of all we don't have enough time for testing new concepts. Mission is very difficult and there is no margin of errors so not tested spacecrafts like Dragon and Orion are not acceptable. The relatively low cost of mission is another argument to use of systems already in operation. Structure of document allows the passage of the thought process of designing a mission to Mars. This is the authors intention. Inspiration Mars 3

4 2. Orbital mechanics The calculations will begin with the idealized case of an interplanetary Hohmann transfer from Earth to Mars and then a rendezvous maneuver in Mars orbit, to evaluate basic parameters and values considering the launch date aspect. Sun Earth (1) Mars (2) r [km] μ [km 3 /s 2 ] 1,32712E inclination of equator to orbit plane [º] 7,25 23,45 25,19 semimajor axis of orbit [km] eccentricity [-] 0,0167 0,0935 inclination of orbit to the ecliptic plane [º] 0 1,85 orbit sidereal period [days] 365, ,99 sidereal rotation period [h] 23, ,62 mass [kg] 1,989E+30 5,974E+24 6,419E+23 distance from Sun R [km] 1496E E+4 Interplanetary Hohmann transfer Let s assume that orbits of the planets involved are circular (eccentricity equal to zero) and coplanar (ignore the inclination). The most energy efficient way for a spacecraft to transfer from Earth s orbit to Mars orbit is to use a Hohmann transfer ellipse. The departure point is at the periapse (perihelion) of the transfer ellipse and the arrival point is at the apoapse (aphelion). When starting from LEO orbit, in this planning phase the differences of distances could be neglected, considering the scale. The circular orbital speed of Earth relative to the Sun is then given as: The velocity of the space vehicle on the transfer ellipse at the departure point is: Therefore the required delta-v at the departure point is: Inspiration Mars 4

5 The delta-v at the arrival point is: For rendezvous to occur at the end of a Hohmann s transfer, the location of Mars in its orbit at the time of the spacecraft s departure from Earth must be such that Mars arrives at the apse line of the transfer ellipse at the same time as the spacecraft does. Phasing maneuvers are not practical for manned missions, due to the large periods of the heliocentric orbits. Rendezvous Time of the transfer from Earth to Mars: Mean motions of Earth (1) and Mars (2): Initial phase angle between Mars and Earth: The phase angle between Earth and Mars when the spacecraft reaches Mars: Minimum wait time for initiating a return trip from Mars to Earth: Total time for a mission would be: Total time for a mission with an immediate return: The above calculations are of course for the very simplified case with no flyby with gravity assist or perturbations included and additional maneuvers necessary to keep the spacecraft in desired orbit. There are several important factors when considering a launch date. The Inspiration Mars 5

6 most important is the alignment of the planets. Launch dates for which the orbit of the spacecraft intersects Mars orbit when Mars is at that location occur about approximately about every two years (2,13 years). The other important factor, as far as the launch date is concerned, is solar activity, especially for manned missions, where we already have higher radiation during the voyage, and potential communication problems could severely affect the mission. The minimum solar activity is said to be on 27 th January 2017, while the ideal phase angle between Mars and Earth occurs on 10 th November 2018 (44,3 according to the geocentric ephemeris) and we shall consider the second date further, which gives us the following schedule: Launch date of spacecraft as a rocket s payload from Earth such that it would reach the transfer circular orbit (additional Hohmann transfer needed from the injection orbit to the parking orbit depending on the vehicle lifting our spacecraft) a few days before 2 nd November 2018, which gives us plenty of time to choose launch window in case of any problems along the way. According to the above calculations, transfer will last about 259 days, which means that we will flyby Mars on 27 th July Without landing on Mars and following the wait time for a proper position of planets, we could assume the return to Earth on 11 th April 2020, which gives us the total flight time of 518 days. However, there is a way to use a gravity assist of involved planets and shorten the time and total delta-v needed for the two-way transfer. Spheres of influence Near a given planet the influence of its own gravity exceeds that of the Sun. The radius of the sphere of influence is given by the equation: Within the planet s SOI the motion of the spacecraft is determined by its equations of motion relative to the planet. Outside of the SOI, the path of the spacecraft is computed relative to Sun. It is not an exact quantity, but a reasonable estimate of the distance beyond which the sun s gravitational attraction dominates that of a planet. The calculated spheres of influence for Earth and Mars are respectively: To study an interplanetary trajectory we assume that when the spacecraft is outside the SOI of a planet it follows an unperturbed Keplerian orbit around the Sun. Due to the vastness of interplanetary flights, we may neglect the size of the SOIs for heliocentric orbits and consider Inspiration Mars 6

7 them, like the planets they surround, to be just points in space coinciding with the planetary centers. Within each planetary SOI, the spacecraft travels an unperturbed Keplerian path about the planet. From the point of view of the planet, the SOI is very large and may be considered to lie at infinity. Planetary departure from Earth In order to escape the gravitational pull of a planet, the spacecraft must travel a hyperbolic trajectory relative to the planet, arriving at its SOI with a relative velocity (hyperbolic excess velocity) greater than zero. On a parabolic trajectory ( spacecraft will arrive at the SOI with a relative speed of zero. In that case the spacecraft remains in the same orbit as the planet and does not embark upon a heliocentric elliptical path. Let s consider a spacecraft leaving Earth. The assumption is that the spacecraft was lifted to h = 300 km circular parking orbit. At the sphere of influence crossing, the heliocentric velocity of the spacecraft is parallel to the asymptote of the departure hyperbola as well as to the planet s heliocentric velocity vector. These velocities must be parallel and in the same direction for a Hohmann transfer such that is positive. is the hyperbolic excess speed of the departure hyperbola: A space vehicle is ordinarily launched into an interplanetary trajectory from a circular parking orbit. The radius of this parking orbit equals the periapse radius r p of the departure hyperbola: where h is the angular momentum of the departure hyperbola (relative to the planet), e is the eccentricity of the hyperbola and is the planet s gravitational parameter. Since the hyperbolic excess speed is specified by the mission requirements, choosing a departure periapse r p yields the parameter e and h of the departure hyperbola. From the angular momentum we get the periapse speed: The speed of the spacecraft in its 300 km circular parking orbit: Inspiration Mars 7

8 Finally, the required delta-v to put the vehicle onto the hyperbolic departure trajectory: The location of periapse, where the delta-v maneuver must occur, is found this way: gives the orientation of the apse line of the hyperbola to the planet s heliocentric velocity vector. If the mission is to send a spacecraft from an outer planet to an inner planet, then the spacecraft s heliocentric speed at departure must be less than that of the planet. That means the spacecraft must emerge from the back side of the SOI with its relative velocity vector directed opposite to. The above calculations are only a preliminary phase of mission design, but it should be noted that the most essential part during the flight will be midcourse maneuvers along the elliptical orbit, since the small errors which are likely to occur in the launch phase may change the target radius even by km. Planetary rendezvous(mars Orbit) A spacecraft arrives at the sphere of influence of the target plane with a hyperbolic excess velocity relative to the planet. In our case, a mission from Earth to Mars, the spacecraft s heliocentric velocity is smaller in magnitude than that of the planet,. Therefore, it crosses the forward portion of the sphere of influence. For Hohmann transfer, these two velocities are parallel, so the magnitude of the hyperbolic excess velocity is simply: If the intent is to go into orbit around the planet, then the aiming radius Δ must be chosen so that the delta-v burn at periapse will occur at the correct altitude above the planet. If there is no impact with the planet and no drop into a capture orbit around the planet, then the spacecraft will simply continue past periapse on a flyby trajectory, exiting the sphere of influence with the same relative speed it entered, but with the velocity vector rotated through the turn angle. Inspiration Mars 8

9 The aiming radius: The purpose of the mission is to enter an elliptical orbit (160 km flyby) around the planet. This will require a delta-v maneuver at periapse P, which is also periapse of the ellipse. Orbital period of 160 km height orbit: Semimajor axis of the capture orbit: Eccentricity: The speed in the hyperbolic trajectory at periapse is: The velocity at periapse of the capture orbit is: Hence, the required delta-v is: At the point of Mars flyby, we will immediately start the return mission and go back to Earth on the similar hyperbola. Planetary departure from Mars Let s consider a spacecraft leaving Mars. The assumption is that the flyby was at h = 160 km parking orbit. The hyperbolic excess speed of the departure hyperbola: Inspiration Mars 9

10 The speed of the spacecraft in its 160 km parking orbit periapse is: The required delta-v to put the vehicle onto the hyperbolic departure trajectory: The location of periapse, where the delta-v maneuver must occur, is found this way: Planetary rendezvous(earth orbit) For the mission from Mars to Earth, is greater than and the spacecraft must cross the rear portion of the sphere of the influence: The intent is to go into orbit around the planet (preferably the same orbit as before, that means 300 km height). Orbital period of 300 km height orbit: Semimajor axis of the capture orbit: Eccentricity: The aiming radius: Inspiration Mars 10

11 The speed in the hyperbolic trajectory at periapse is: The velocity at periapse of the capture orbit is: Hence, the required delta-v is: Total delta-v Summarizing, the total delta-v required for the mission for Hohmann transfer would be: and with use of a gravity assist near the planets: Inspiration Mars 11

12 3. Environmental Control and Life Support Systems Key component of the mission is a system of supporting life onboard the spacecraft against the hostile environment of space. Due to the lack of atmosphere in space it is necessary to artificially control both the environment as well as provide enough sustenance to assure survival of the crew. For that purpose, Environmental Control and Life Support Systems (ECLSS in short) needs to be installed and the amount of payload necessary to keep the crew alive needs to be estimated. 3.1 System selection Unfortunately, due to the launch time restraint of four years it is not possible to reliably design, manufacture and test any new life support technology for the mission. Therefore, the ECLSS system will be based on existing designs. When considering the choice of ECLSS system for our mission, two design parameters were taken into consideration. The system reliability as well as experience in space environment are critical since any failure can cause loss of life aboard the spacecraft. Second design consideration is efficiency of the system. Higher efficiency of the life support systems reduce the necessary launch payload allowing for wider variety of launch options as well as reducing the overall cost of the launch. The only system that fulfills all the criteria for the necessary period of time is ECLSS aboard the International Space Station. Therefore, all further consideration will be based on ISS ECLSS data. It is worth noting that the ISS ECLSS system is designed for six-man missions instead of our design requirements of two-man mission. Due to scalability considerations as well as possibility of decreasing the systems reliability while redesigning it, the system will be adopted as is, without any significant changes. 3.2 Composition of the system Environmental Control and Life Support System consists of seven main subsystems: Air Revitalization Atmosphere Control and Supply Temperature and Humidity Control Water Recovery and Management Waste Management Fire Detection and Suppression Other However, due to the longevity of the mission we need to make an addition to the overall system. Exposure to extended periods of microgravity can cause bone loss, muscle volume Inspiration Mars 12

13 decrease and increase in calcium levels in bloodstream. In order to minimize these effects it is necessary to include aerobic as well as resistance training devices to reduce the negative effects of microgravity on the crews health. Below are short descriptions of each module and its purpose. Air Revitalization Subsystem During the duration of our mission, the crew will live in an artificially controlled atmosphere aboard the spacecraft. The composition of the atmosphere was chosen based on the atmosphere requirements for Space Station Freedom operational composition. Parameter Unit 90-Day Operational Status CO2 Partial Pressure Pa 400 max Temperature K Dew Point K Ventilation m/s O2 Partial Pressure kpa Total Pressure kpa Oxygen/Nitrogen Air Ratio %/% 21/78 The main purposes of the Air Revitalization Subsystem are CO 2 removal, O 2 generation, trace contaminants control and O 2 recovery from CO 2. During normal human metabolic process, significant amounts of CO 2 are generated. In a closed environment of a spacecraft that could lead to a rapid increase in CO 2 concentration in the atmosphere. In order to control the CO 2 levels in the atmosphere a molecular regenerative sieve (such as zeolite) can be used. It has a crystalline structure of with an extremely large surface area due to large pores. CO 2 is trapped in the pores while O 2 and N 2 can pass through. The sieve can repeatedly absorb and de-absorb CO 2. The de-absorption process can be induced by either exposing the sieve to space vacuum or heating and reducing the pressure with a heat pump. Another important parameter to control is the O 2 Partial Pressure. The effects of insufficient oxygen pressure can include decreased night vision, impaired memory and coordination, unconsciousness, convulsions and death while too high partial pressure can lead to lung irritation. It is, therefore, critical to keep the oxygen levels close to the sea levels value of 21.4 kpa. The oxygen necessary to keep the stable atmosphere will be supplied to the system via a process of waste water electrolysis. Moreover, through the application of Sabatier Reaction, approximately fifty per cent of oxygen can be recovered from carbon dioxide. Inspiration Mars 13

14 A spacecraft environment has several sources of contaminants within. Offgassing from materials, crew metabolic byproducts, food preparation as well as scientific experiments all contribute to the generation of contaminants within the spacecraft. However, even small amounts of contaminants can lead to a potentially hazardous situation. It is, therefore, critical to monitor and remove excess contaminants. Careful selection of materials to minimize offgassing and dust generation can significantly reduce the overall amount of contaminants needed to be actively removed. The systems applied on International Space Station to monitor and remove excess contaminants are respectively Volatile Organic Analyzer, consisting of Gas Chronograph and Ion Mobility Spectrometer, and High Efficiency Particulate Atmosphere (HEPA) filters. Additionally, HEPA filters placed in strategic positions in the ventilation system allow for removal of microbial organisms as well as contaminants. Atmosphere Control and Supply Subsystem There are four major elements of the spacecrafts atmosphere H2, O2, H2O and CO2. It is necessary to continuously monitor their levels. One possible solution is to place a separate analyzer for each element. Another method is to use mass spectrometry to monitor the level of gases aboard the spacecraft. Mass spectrometer, however, can't distinguish between elements of the same molecular weight so in order to identify these compounds it is necessary to include additional, specific monitors. A combination of both approaches is therefore the most applicable choice, providing a benefit of additional redundancy of the system. While most of the oxygen used during the mission will come from the process of water electrolysis, it is necessary to also include oxygen in other forms in case of an emergency or maintenance of the oxygen generation system. Pressurized tanks can be used to store the back-up oxygen supply. Apart from oxygen, another constituent of spacecrafts atmosphere is nitrogen. While early space programs utilized pure oxygen atmosphere aboard the spacecraft, modern designs steer towards a mixture of oxygen and nitrogen similar to Earth sea level air composition at 21% and 78% per cent respectively. Such composition has the benefit of inhibiting fire spread in case of emergency as well as preventing physiological effects of a pure oxygen environment on the crew. Additionally, it simplifies the design and test of equipment prelaunch. The nitrogen can be stored in the form of hydrazine (H2N2) and separated into hydrogen and nitrogen, in the process of catalytic dissociation, when needed. Another important goal of the ACS system is to control the total pressure of the spacecraft atmosphere. Pressure below the design limit can negatively affect the temperature and humidity control systems and can lead to physiological effects on the crew. Underpressure Inspiration Mars 14

15 can be dealt with by having makeup gasses stored in gas tanks while overpressure can be solved by either compressing the excess gases or venting them into space. Temperature and Humidity Control Due to the closed character of a space environment the spacecraft would quickly overheat without a proper system to control the temperature within. Electronics, lighting, solar heating of the spacecraft as well as crew metabolic processes act as heat sources inside the space craft. The goals of the THC system is to continuously adjust inside temperature and humidity to a degree comfortable for the crew while preventing condensation from forming on windows, walls and equipment. The preferred atmosphere conditions vary between people from 18 ⁰C and 27 ⁰C for temperature and between 25% and 75% for relative humidity. The method of controlling the temperature aboard the spacecraft is through a forced ventilation system. Ventilation flow moves the spacecraft atmosphere through condensing heat exchangers. The heat can be then radiated into space. Moreover, excess moisture in the spacecraft atmosphere can be removed through heat exchangers via an addition of "slurper bars" or wick-type devices. Water Recovery and Management Without water, an average person can die due to dehydration in three days. However, due to the longevity of the planned mission it is not feasible to carry the full amount of water on board. It is, therefore, critical to include a high efficiency water recycling system aboard the spacecraft to cut into the necessary amount of water at launch. Thankfully such systems are already in use aboard the International Space Stations Destiny module. With recovery rates of 100% and 83% for moisture and urine water recovery respectively it will be a key element in a long-distance space mission. In such a system, water is recovered from urine during a low pressure distillation process. Then it is combined with the rest of wastewater and moved through Water Processor, in which free gasses and solid objects are separated and subsequently the water is further filtrated through a series of multi-filtration beds. Any remaining organic contaminants and microorganisms are removed by a high-temperature catalytic reactor assembly. Finally, the water is checked by electrical conductivity sensors and clean water is sent to storage tanks. Waste Management There are four types of wastes of interest in terms of Environment Control and Life Support: metabolic wastes consisting of moist solids, other solid wastes, liquid wastes consisting of urine and waste water and gaseous wastes. Wastewater and gases recycling were discussed earlier in the chapter. Unfortunately, to-date there was no metabolic waste recycling device aboard any space station. The systems on International Space Station store the metabolic wastes until they can Inspiration Mars 15

16 be transported away from the station via Space Shuttle or Progress vehicle. This approach is not applicable in a deep-space manned mission since biological solid waste is generally unstable due to its water content. Storing of such wastes for extended periods of time would result in decomposing, leading to growth of undesirable microorganisms, generation of noxious gasses and create foul odours. Currently a process of recovering water from the wastes creating dry, plastic-encapsulated bricks which can be used for radiation protection is being developed. However, due to a narrow launch window for our mission, we will assume the primary method of biological waste management as periodic jettisoning into space. Fire Detection and Suppression It is important to detect fires early aboard the spacecraft. An undetected fire can be disastrous to the mission. Using fire resistant materials in the design and designing the spacecraft environment to slow the propagation of fire can reduce the possibility of a fire. Even so, the detection systems are necessary aboard the spacecraft. These systems include flame detectors based on visible, infrared and ultraviolet emissions as well as smoke detectors based on detecting the particles emitted by burning materials. Following the detection of a fire, it can be suppressed with CO2, N2 or by depressurizing the environment. Any fire will create hazardous byproducts. The ability to remove these byproducts from the environment of the spacecraft is an essential part of the FDS module. The approach, however, varies based on the scale of the fire and the amount of byproducts produced. Large fires producing a lot of smoke are best dealt with by venting the contaminated atmosphere into space. For smaller, localized fires a portable contamination control device can be used to remove most of the contaminants so that Air Revitalization System can remove the rest. Other While technically not a part of the ECLSS system some other elements and requirements of the flight need to be considered. One of these concerns is the amount of payload needed at launch for food, water and oxygen. The International Space Station rules allow for a maximum weight of daily food portion of 1.71kg including 0.45kg of packaging. We will assume these values for our mission. Throughout a 500 day mission the total amount of food-related weight will total to: In case of emergencies additional 103kg, an equivalent of 30 days of food consumption, will be added to the total value of 1806kg. Due to close to 100 per cent efficiency of the moisture recovery of the ECLSS system water losses on the spacecraft occur mainly through urination and feces. Assuming an average volume of daily urination per person to be 1500ml and 83 per cent water recovery rate from Inspiration Mars 16

17 urine we can easily estimate the water losses through urination during the course of the mission: Moreover, additional 100ml of water is lost daily per person through feces and, since our mission consists of a male and a female, another 50ml through vaginal secretions: Therefore, the minimum amount of drinkable water necessary for the flight is 380L. However, another 30L, equivalent of thirty days of water loss, will be added in case of emergencies or need of maintenance of the system to a total of 410L. Oxygen consumption is the most problematic to accurately predict. Due to effects of microgravity, astronauts aboard the International Space Station spend two hours daily exercising to minimize the effects of bone loss and muscle volume decrease. The exercise consists of resistance training, treadmill running and stationary bike riding. The best method of estimating oxygen consumption throughout the exercise would be to perform a clinical test with measuring devices. However, due to unavailability of such measures, constant oxygen consumption levels during the exercise were assumed and the American College of Sports Medicine treadmill running equation was used. The reference values of speed and inclination of the treadmill were 12 km/h and 0% respectively. where, V - speed in meters per minute G - treadmill inclination VO2 - gross oxygen consumption in ml/kg/min Assuming a crew of two, with an average weight of 70.8kg we reach the total oxygen required: However, oxygen is burned also outside of the exercise. American College of Sports Medicine reports the value of resting oxygen consumption to be at a level of 3.5 ml/kg/min. That leads us to the amount of oxygen burned during the day, outside of exercise: Inspiration Mars 17

18 The total amount of oxygen consumed during the mission is then: However, not all of it is necessary for the mission. Thanks to the Sabatier Reaction, the ECLSS has the capability to recover 50% of oxygen from carbon dioxide cutting the necessary amount of oxygen in half to L. Since oxygen aboard the spacecraft will be stored in the form of water, it is necessary to transform this value into the amount of water necessary. Knowing that one mol of oxygen has 22.4L and weight 32g we can estimate the weight of the oxygen calculated: Having the weight of the oxygen and knowing that 1kg of water contains kg of oxygen we can calculate the final amount of water necessary for oxygen production: However, due to the possibility of leakages or necessity of maintenance we will adjust that value by additional 35kg, an equivalent of 30 days, to 637.5kg. Total consumable payload it then: Food Kg 1806 Drinkable Water Kg 410 Oxygen Generation Water Kg Total Kg Exercise Equipment Experiences with the International Space Station has shown that during the six month mission aboard the space station the astronauts could lose as much as 15% of muscle volume and up to 25% bone mass without proper exercise equipment. Therefore it is necessary for our mission to provide similar equipment to the crew. The equipment would be modeled after the International Space Stations Cycle Ergometer (cycling), Advanced Resistive Exercise Device (weight-lifting) and a treadmill (running). Habitable Volume For long-duration space flights adequate habitable space inside the spacecraft is important to provide an element of privacy which can be instrumental in lowering the tension between Inspiration Mars 18

19 the crewmembers. Following NASA Subsystem Integration Standards (NASA-STD-3000) we can find the necessary amount of habitable volume aboard the space station. Fig. 1 - NASA -STD-3000 habitable volume guide We can conclude that the amount of habitable volume aboard the spacecraft should be 38m 3 or higher for the optimal living conditions. It will be a important condition further in the project when selecting the type of spacecraft to be used for the mission. Inspiration Mars 19

20 3.3 System Weight Estimation For further conceptual design, it is necessary to estimate the approximate weight needed at launch. The conceptual estimation of the weight of ECLSS modules is based on the International Space Station weights. System Mass [kg] Volume [m 3 ] Air Revitalization System (ARS) Carbon Dioxide Removal Assembly (CDRA) Trace Contaminant Control Subsystem (TCCS) Major Constituent Analyzer (MCA) Oxygen Generation Assembly (OGA) Temperature and Humidity Control System (THCS) Common Cabin Air Assembly (CCAA) Avionics Air Assembly (AAA) Intermodule Ventilation Fan Intermodule Ventilation Valve High Efficiency Particle Atmosphere (HEPA) Filter Water Recover and Management (WRM) and Waste Management (WM) Water Processor (WP) Process Control Water Quality Monitor (PCWQM) Urine Processor (UP) Fuel Cell Water Storage Condensate Storage Commode/Urinal Fire Detection and Suppression Smoke Detector Portable Fire Extinguisher (PFE) Other Food Drinkable water reserve Oxygen generation water reserve Cycle Ergometer Advanced Resistive Exercise Device (ARED) Treadmill Total >12.94 Inspiration Mars 20

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