Inspiration Mars. International Student Design Competition WUT. Dream Team

Transcription

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

21 4. Launcher selection Considering this will be the fist manned flight to red planet, we can not predict every complication we might find on our way. To reduce the risk we have to rely on trusted, tested many times, technologies and make our mission as simply as it is possible (the less elements take part in program-the less elements can let us down- the reliability rises). Having in mind those statements, we decide to use only one, heavy-lift rocket and that would be one of the group of rockets which already flown on Martians missions or those which are under development, but in the order to take people to red planet. We chose Delta IV Heavy, Atlas V, Ariane V ECA, Falcon Heavy and SLS Block II to further analyses, to see which one have best suited performance to our mission. 4.1 Potential vehicles Delta IV Heavy general I stage boosters II stage mass to LEO [kg] mass to GEO [kg] company launch cost ULA ~381M$ specific impulse propellant engines thrust [kn] 410 s LOX/LH2 1x RS-68A specific impulse propellant engines thrust/engine [kn] 410 s LOX/LH2 2x RS-68A specific impulse propellant engines thrust [kn] 462 s LOX/LH2 1x RL-10B 110 Atlas V Heavy general I stage boosters II stage mass to LEO [kg] mass to GEO [kg] company launch cost ULA ~223$ specific impulse propellant engines thrust [kn] 338 s LOX/RP-1 1xRD specific impulse propellant engines thrust [kn] 338 s LOX/RP-1 1xRD specific impulse propellant engines thrust [kn] 451 s LOX/LH2 1xRL 10A ,2 Inspiration Mars 21

22 Ariane V ECA general I stage boosters II stage Falcon Heavy general I stage boosters II stage mass to LEO [kg] mass to GEO [kg] Space Launch System block II company launch cost arianespace ~120M$ specific impulse propellant engines thrust [kn] 431 s LOX/LH2 1xvulcain specific impulse propellant engines thrust/engine [kn] 275 s solid 2xP specific impulse propellant engines thrust [kn] 446 s LOX/LH2 HM-B 64,7 mass to LEO [kg] mass to GEO [kg] company launch cost SpaceX ~110M$ specific impulse propellant engines thrust/engine [kn] 311 LOX/RP-1 9xMerlin 1D 801 specific impulse propellant 311 LOX/RP-1 specific impulse propellant 342 LOX/RP-1 engines 2x9xMerlin 1D engines 1xMerlin 1D vacuum thrust/engine [kn] 801 thrust [kn] 801 general I stage boosters II stage mass to LEO [kg] mass to GEO [kg] company launch cost NASA ~500M$ specific impulse propellant engines thrust/engine [kn] 452,3 LOX/LH2 4x RS specific impulse propellant engines thrust [kn] 259 s HTPB SRB specific impulse propellant engines thrust/engine [kn] 448 s LOX/LH2 3x J-2X Inspiration Mars 22

23 4.2 Performance analyses At first we like to see limits on masses that our vehicles are able to launch and how much it would cost us. The chart presents payload weights each rocket can put to LEO or GTO orbit and approximate costs of launch. As we can see Delta IV Heavy and Atlas V, with very high costs, do not have such impressive performance. The SLS Block II has huge costs, but also lifts very big weight of payload. However we should not need that much of payload to complete our mission, which makes using this rocket unreasonable. costs and payload payload to LEO [kg] payload to GTO [kg] costs of launch M$ delta IV Heavy atlas V 551 Ariane V falcon heavy SLS Block II 0 That makes us focus our attention on Ariane V and Falcon Heavy. Launch of Falcon Heavy is cheaper than Ariane V and lifts more payload at the same time. That makes Falcon Heavy our favorite, but we have to keep in mind this rocket have not flied as many times as Ariane V, so its reliability is not that sure as Ariane s. If we preliminary estimate our payload to at about metric tons, Delta IV, Atlas V and Ariane V will not be enough for us. From this point of view Falcon Heavy wins with other rockets. Inspiration Mars 23

24 500 specific impulse in vacuum [s] I stage Boosters II stage delta IV Heavy atlas V 551 Ariane V falcon heavy SLS Block II Comparing specific impulses of each vehicle we can see Falcon Heavy does not have such impressive performance at this point. It means the rocket uses more propellant, to generate the same thrust, that other rockets can reach with lower propellant consumption ( ). That makes our rocket little bit ineffective, giving that the rest of vehicles have specific impulse at about 100 seconds higher Thrust of I stage+boosters [kn] 0.00 delta IV Heavy atlas V 551 Ariane V falcon heavy SLS Block II Thrust of II stage [kn] delta IV Heavy atlas V 551 Ariane V falcon heavy SLS Block II Inspiration Mars 24

25 As we can see on charts above Falcon Heavy may have low specific impulse, but generate much more thrust than Delta IV, Atalas V and Ariane V. Merlin engines are also under improvement works, tests and analyses, so there are pretty good chances that their specific impulses will increase. If SpaceX finally change kerosene for liquid methane fuel, as promised, performance should more impressive: costs come down and specific impulse rise (slightly-in Earth orbit missions it would not make such difference, but in context of flight to Mars it is big advantage). Another benefit of Falcon Heavy is type of propellant: kerosene or methane- does not matter which one, it is still much cheaper and accessible than liquid oxygen and hydrogen. It is also (for some people surpassingly) more environment-friendly. Kerosene and methane, during incomplete combustion, emit carbon dioxidetrue. Liquid hydrogen and oxygen (generating more energy at the same time) do not- correct. But hydrogen does not occur in nature in its pure form, so to produce one, on the way we (in most cases) emit more carbon dioxide than it would be created during incomplete combustion of kerosene or methane. In perspective of future manned missions to colonize red planet, we can keep eye on Space Launch System- it can hold bigger payload ( almost 6 times heavier than Delta IV) with less than 2 times bigger cost per launch, but in current mission- flyby Mars-it is much more than we need, and first fly tests are planned in 2017 (and there are some chances the project will be cancelled)-so we can not relay on that technology yet. 4.3 Conclusions Giving those circumstances the best vehicle to our mission profile seems to be Falcon Heavy. It launch the biggest payload, with the lowest cost. On its official website, SpaceX claims that the rocket can bring kg of payload to Martian orbit- considering the fact that our crew capsule and other necessary equipment are estimated to kg, it is more than enough for us. Most of the construction is based on Falcon 9 rocket, which flied 8 times and 8 times succeeded, also used in resupply International Space Station mission. In our opinion it makes SpaceX s technologies regarded as safe and reliable enough. Falcon Heavy is also designed in order to take people to Mars, so it is the best prepared rocket to do so- there is no need to provide additional developments for Martian trajectory. In further analyses we will base on Falcon Heavy data. Inspiration Mars 25

26 5. Spacecraft Design Following the conceptual design of the Environment Control and Life Support System and selection of the rocket we can start to design the spacecraft on which the mission will be carried out. The spacecraft needs to have ample space to store the necessary equipment, food and water as well as enough habitable space to allow two crew members to live inside for 500 days. Requirements from ECLSS analysis state that the necessary volume should be at least m 3 without food and water and should contain at least 38m 3 of habitable volume. Since the launch is selected for 2018, there is no time to design a completely new spacecraft. Therefore, keeping the volume requirement in mind we will consider the possible options. Spacecraft Habitable Volume [m 3 ] Soyuz 8.5 Shenzhou 14.0 Gemini 2.26 Apollo 6.17 Space shuttle orbiter 74 Orion 8.95 From the analysis we can notice that only Space Shuttle orbiter has the necessary amount of volume to fulfill the requirements of our mission. That leaves us with two possible solutions. Either adopt Space Shuttle orbiter as the missions spacecraft or modify one of the other spacecrafts to include bigger habitation volume. Space Shuttle Orbiter The Orbiter appears as an obvious first choice of the spacecraft. It provides more volume than required. Moreover, since its cargo bay won't be necessary for the mission, it is possible to extend the habitable volume of the spacecraft by placing a Spacelab-like module inside the cargo bay. While the exact data has proven difficult to find, it was estimated that the habitable volume could be increased by roughly 60m 3 this way. Unfortunately, Space Shuttle Orbiters are no longer operational. The mission would require to either obtain one of the formerly retired Orbiters or create a new Orbiter as well as Spacelab-like cargo bay module if necessary. Moreover, since the life support system aboard the Orbiter has only 30 person days worth of life, the spacecraft would need to be adapted for mission purposes. Modified Spacecraft While the available spacecrafts appear to have too little volume inside to provide acceptable living conditions for a deep-space mission, it is possible to extend the volume by creating a habitation module and connecting it with the rest of the spacecraft via a docking module. Inspiration Mars 26

27 While this approach is technically possible for all of the available spacecraft options excluding Space Shuttle Orbiter, after careful analysis it appears to be an insufficient requirement for some of them. Apollo and Gemini spacecrafts performed their last flight in 1972 and 1966 respectively. That means that the selection of those spacecrafts as our carriers would require revitalization and modernization of the spacecraft program. Such approach is prohibitive in both cost and time required. Orion Multi-Purpose Crew Vehicle is a spacecraft currently still under development. While successful tests in the form of a multi-hour unmanned flight were performed, the first unmanned mission of the vehicle is scheduled for That means that due to our launch time window of 2018, Orion spacecraft is inapplicable for our project. Both Soyuz and Shenzhou spacecraft pass all of the above considerations. While Shenzhou is a more modern design and contains more habitable space inside, Soyuz has been in operation since 1967 and has proven to be the safest option of any spacecraft to-date. Soyuz spacecraft will be assumed as the vehicle of choice. Following the selection of spacecraft we have to decide on the method of extending the habitable volume inside. Soyuz spacecraft contains the active hybrid docking port (SSVP-M) which allows it to dock to the International Space Station Zvezda Module. Utilizing the port we can attach a habitation module similar to the Zvezda Module. In order to construct such a module we can use the Multi-Purpose Logistics Module formerly used aboard the Space Shuttle Orbiter as basis. Multi-Purpose Logistics Module is a large pressurized container of cylindrical shape with a docking port on either end. The module has the external dimensions of 6.6m length, 4.57m width and 31m 3 of habitable volume. After connecting the habitation module with Soyuz spacecraft and implementing the ECLSS system with the necessary payload the data of the spacecraft will be Weight [kg] Habitation Maximum Volume [m 3 ] Diameter [m] Length [m] Soyuz Module ECLSS & Crew Total Comparing the data with the Falcon Heavy requirements of 13200kg of payload, 5m diameter and 13.1m length it appears that our spacecraft is too large. However the Soyuz module contains many unnecessary components which will be implemented in the custom Multi-Purpose Logistics Module. Because of this, there is the possibility to redesign the Inspiration Mars 27

28 Soyuz spacecraft to our needs. It also will reduce payload mass to the level that we are capable to put on the Mars orbit using Falcon Heavy launcher. The habitation volume inside the modified spacecraft, while not satisfying the optimal recommendation of the NASA-STD-3000 stays on an acceptable level. Spacecraft selection Both methods are viable means of achieving our goal of a Mars fly-by mission. However, two of the design criterions in the project are cost and complexity of the design. With the costs of $1.7 billion to create Space Shuttle Endeavor and an approximate of $450 million the need to revitalize, redesign and launch the mission in a Space Shuttle appears to be cost prohibitive. Moreover, the necessity to repurpose the Orbiter with custom ECLSS systems as well as addition of Spacelab-like habitation module inside the cargo module adds to the complexity of this solution. Because of this, the preferred method is the modification of an existing Soyuz spacecraft. Inspiration Mars 28

29 6. Power Management System For powering spacecraft we need reliable, light and low volume. In next chapters we will consider three powering systems: 1. Solar Panels 2. Hydrogen Fuel Cell 3. Battery System Schema Inspiration Mars 29

30 6.1 Power Consumption Maximum power consumption is displayed in table: Power Consumption CDRA 0,86 kw TCCS 0,18 kw MCA 0,088 kw OGA 1,47 kw CCAA 0,468 kw AAA 0,083 kw IMV 0,055 kw IMV 0,006 kw Smoke detector 0,002 kw WP 0,3 kw PCWQM 0,03 kw UP 0,091 kw Total 3,633 kw Maximum total consumed power: 3,633 kw. 6.2 Battery Pack To ensure that during flight in shadow the systems could be working and in emergency situation the Hydrogen Fuel Cell system could be started the minimum capacity of battery system should be higher than daily consumption. Daily consumption: P t total consumed power Battery pack minimum capacity: 130,788 [kwh] Battery systems vary in capacity and lifecycles. For longtime mission toward Mars we need reliable source of power. The best solution on market are Li-ION batteries. They are tested through years, have quite good power density(~150 Wh/kg) and are safe in use. Inspiration Mars 30

31 According to table we will use SAFT VES 180 cell: Battery pack completed using VES 180 cells will have properties: Inspiration Mars 31

32 6.3 Solar panels To calculate minimum area of solar panels we need to know efficiency. Power output General power output formula: η sc efficiency of solar panel delivered by manufacturer P in cell power input[w] P out cell power output [W] A c cell area vector [m 2 ] H input power density vector [W/m 2 ] α angle between H vector and A c vector Air mass coefficient Comparing efficiency of solar panels on earth and in space we can easily observe that the air in earth atmosphere influent the efficiency of solar panels. Although efficiency on earth Is higher than in space, maximum power output we can get in space is bigger because of fact that solar light have bigger energy before entering atmosphere. Solar cells efficiency comparison Manufacturer Efficiency [%] Spectrolab 30 Emcore 29.5 Azurspace 29.5 CESI 35 The best efficiency have solar cells provided by CESI: 35% Inspiration Mars 32

33 Solar Irradiance in space Power density depends on distance of the object from the Sun. General formula for power density in space: R sun radius of sun D object distance from sun center H sun Power density on sun surface defined by Stefan Boltzmann s blackbody equation General formula Putting together in general formula: Minimal solar power output For simultaneous spacecraft powering and charging battery pack we need to calculate minimum output power higher than consumed power. We consider charging full battery pack and ensuring power for current demands. The solar cells should have minimum area: 36 m Hydrogen fuel cell In hydrogen fuel cells we convert chemical energy into electricity. In that process we use hydrogen and oxygen to perform chemical reaction in which result is electrical power. In matter of fact that there are few manufacturers of hydrogen fuel cells we chose solution provided by Ballard: FCgen 1300 This hydrogen fuel cell can give us 3,7 kw of energy from 42l/min of hydrogen. To make this system scaled down we need to use liquid hydrogen and oxygen. The amounts of hydrogen and oxygen needed for reaction: Hydrogen Fuel Cell Conversion rate 1l of liquid hydrogen to gas 845 l Hydrogen Consumption 2520 l/h Liquid Hydrogen Day consumption 71, l/day Cubic meter duration 13, days Oxygen consumption l/h Conversion rate 1l of liquid oxygen to gas 861 l Oxygen vs hydrogen conversion 8 Liquid Oxygen Day consumption 583, l Inspiration Mars 33

34 We can see that hydrogen fuel cell oxygen consumption is too high to consider using it on longtime mission. So finally the power management system schema: Inspiration Mars 34

35 48 hour simulation of 20 hours in shadow Battery capacity versus solar power input 48h simulation solar power input battery capacity Inspiration Mars 35

36 6. Cost Total cost of this mission is hard to define at this level, because we do not have access to the actual data from executive companies, but we based on information found on internet and estimated what is following: Falcon Heavy Soyuz ISS module Total We find data for a few another components but the input of them is much smaller than the listed ones. 7. Mission Science Program One of the concerns of a deep-space manned mission is the radiation. On Low Earth Orbit the Earth magnetic field protects the astronauts aboard the International Space Station. This inhibit the available means of conducting research on long-term effects of exposure to space radiation. This research could be conducted on board of our mission. The screw would take saliva samples in regular periods. The samples would then be dried and stored at 'room' temperature preventing degradation of the genetic material. When the mission finishes, these samples would be retrieved from the spacecraft and analyzed for any DNA damage acquired due to radiation and microgravity. Another experiment can involve diagnostic sonography during a deep-space mission. Vanderbilt Center of Space Physiology and Medicine predicts that due to its low cost, versatility and low operational needs diagnostic sonography could be used to investigate cardiac contractility and physiologic changes due to space. Changes in bone density can also be monitored using diagnostic sonometry. Inspiration Mars 36

37 8. Summary Our analysis shows that manned Mars flight is technically possible. We try use systems that exist. This is only way to create all system within the prescribed time. We achieved this goal. We allow only modification in the context of ready-made solutions. This achievement allows to reduce costs significantly. Estimated costs are probably undervalued because we don't date for ale cases, especially for national agencies products. Unfortunately we don't have enough instrumentalities to find better data because group of students can't be a side during a negotiations. In conclusion our paper shows key points of manned mission design. Document is the very first step to design the real mission. Paper is useful tool to show problems and solutions of manned flight to Mars for public opinion. Inspiration Mars 37

38 9. References 1. M. Griffin, J. French, Space Vehicle Design, AIAA, G. Swinerd, How Spacecraft Fly, Copernicus Books, V. A. Chobotov, Orbital Mechanics, Third Edition, AIAA, H. Curtis, Orbital Mechanics for Engineering Students, Elsevier, P. Fortescue, J. Stark, Spacecraft Systems Engineering, Third Edition, Wiley, W. Larson, J. Wertz, Space Mission Analysis and Design, Microcosm Press, Wieland, P. (1994). Designing for Human Presence in Space: An Introduction to Environmental Control and Lift Support Systems. NASA Reference Publication 10. National Aeronautics and Space Administration (NASA). (2008). International Space Station: Environmental Control and Life Support System. Retrieved from Carrasquillo, R. (2013). ISS Environmental Control and Life Support System (ECLSS) Future Development for Exploration. 2nd Annual ISS Research and Development Conference. Retrieved from Advanced Life Support Research and Technology Development Metric - Initial Draft. (n.d.) Retrieved from Bacal, K. (n.d.). Section III.3.2 Waste Management. Retrieved from me/media/section%20iii.3.2%20waste%20management.doc 14. National Aeronautics and Space Administration (NASA). (July 1995). Man-Systems Integration Standards Volume I. Inspiration Mars 38