First Team Project Nathan Schilling Artificial Gravity...

First Team Project Artificial Gravity Habitat

Bianca Foltan Zane Gardenhour

Logan GibbMandeep Sawhney

Nathan Schilling

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Design/ Mission Overview● Design a spinning space habitat capable of creating up to 1 Earth g of acceleration.● The space station will be capable of transporting a 4 man crew to Mars. Upon arrival part of the

space station habitat will be deployed to the Martian surface to serve as a human habitable outpost.● For logistics reasons, the Mars Ascent Vehicle would ideally be landed on the Martian surface far in

advance of this mission.● The crew will descend in a landing capsule, and take refuge in the new habitat until ready for

departure.● The landing capsule would be left behind, and the crew would return to the space station in the

MAV.

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RASCAL Competition Requirements● As this design is a precursor to the RASCAL competition, we strove to meet

RASCAL guidelines wherever possible and create a feasible space habitat.● RASCAL requirements that guided our design include:

○ Launch at least partially on block 1A/B SLS○ Mission duration of 1100 days, to Mars and back.○ Hybrid Propulsion system uses 750 kW Beginning-of-Life solar arrays○ All technology capabilities must be implementable and operational by 2029○ Designs are consistent with

■ NASA Technical Standards 3000 and 3001■ NASA Human Integration Design Handbook

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Interplanetary Transport Design● The spinning space station features 2 pressurized habitat modules: 1

Mars/Moon habitat module and 1 space habitat module.● The space habitat will be where the crew lives during their mission flights.● The Interplanetary habitat will remain dormant during the flight and be landed

on the mission target surface upon arrival.● After spin-up, portions of the space habitat module will experience up to 1

Earth g of acceleration.● The station features a hybrid solar chemical propulsion system, the details of

which are outside the scope of this project.● Power is generated by a set of rotating solar arrays.

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Delivery To Low Earth Orbit

● SLS Rocket 1B○ Capable of delivering 105

metric tons to low earth orbit

○ 8.4 meter diameter payload○ Max payload volume of 286

m^3

● Falcon 9 Rocket○ 22.8 metric tons to low

earth orbit○ 4.6 meter diameter payload

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Space Assembly● First launch will use a SLS 1B.

○ A crew of 4 to build the spacecraft○ The truss system, the rotational system, solar panels, and the propulsion

system.■ Will consist of 3 major builds:

● Main habitat will be attached to solar panels, and first section of truss● Remaining truss sections, along with propulsion system and rotational system will be

attached● Solar panels will be relocated to final position at the end of the final truss section

● Second launch will use a Falcon 9 Rocket.○ Will carry secondary Mars habitat, exterior oxygen tanks and communication

dish■ Build:

● The secondary habitat will self dock at its proper location.● The communication dish and extra oxygen tanks need to be attached to their final

location.

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Space Assembly● Up to final launch will use Falcon 9 Rockets.

○ Will deliver all necessary propellant required for missions.

● Final launch will use Falcon 9 Rocket.○ Mission astronauts and all required supplies will be sent to the fully

assembled spacecraft○ The original assembly crew will use reentry vehicle to return back to Earth

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Artificial Gravity Accommodations● The primary design constraints for a rotating space station are:

○ Mass of the space station itself (cost).○ Nausea caused by spinning.○ Lightheadedness from the gravity gradient between head and

feet.● In order to make such a station feasible, mass must be minimized

and astronauts have to be comfortable.● The governing equation is, Gapparent = ω²·r● Gapparent is the simulated gravitational acceleration at the floor.● r is the radial distance from the observer to the center of the spin

axis. ● ω is the angular velocity of the spacecraft● Radius is inversely proportional to the square of angular velocity ● Looking at the figure on the right, it is clear that maximizing the

angular velocity is paramount in reducing radius and mass

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Artificial Gravity Accommodations

● Design requirements include simulated gravity up to 1 Earth g.

Gapparent = 9.81m/s²

● Human factors studies have shown that trained individuals can adapt to a spin rate of ω = 6 revolutions/min.● Using the governing equation, r = (Gapparent /ω²)(1/2) , we find that the minimum radius of the spacecraft is

approximately 25m.● The gravity gradient is a measurement of the apparent gravity at two different radial distances and can be expressed

by G1/G2=(R1–R2)/R2. ● For a station of r = 25 m, an astronaut 2m tall, and a spin rate of ω = 6 rev/min we find that G1/G2 = 8.7%, which is

within the recommended limit of most case studies.

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Starting point:Radius = 25 mAngular Velocity = 6 rev/min

Dynamics and Stability● A rigid body is stable if its spin axis corresponds to the principal axis with the smallest or largest

moment of inertia.● If the rigid body has two equivalent principal moments of inertia, then the principal axis

corresponding to the distinct moment is the only stable axis.

For the typical barbell spacecraft, Ixx < Iyy = Izz . Since one of the principal axes corresponding to a repeated moment of inertia is also the spin axis, the system is unstable.

Here Ixx < Iyy < Izz where y is the spin axis. The middle principal moment of inertia corresponds to the spin axis and so the system remains unstable.

Solution: Break up the “pod” shapes to eliminate repeated moments of inertia. Build up mass along the z axis so that Iyy > Izz . Now, Ixx < Izz < Iyy where Iyy is the spin axis. The system should remain relatively stable.

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● In order to maintain load balancing but break up the principle moments of inertia, the power system and solar arrays, are mounted on the arm opposite from the space habitat.

● To further increase mass on the starboard side and push out the center of gravity, attitude thrusters (for spin up and spin down) and their propellant tanks are mounted at the very end of the truss system.

● Depending on the final mass of the power system, and solar arrays, the length of the starboard truss can be modified to balance the moment from the rotating habitat module.

● The crewed Mars habitat/lander will be mounted in the positive z direction and be balanced in the negative z direction by the thermal radiators of the space station.

● The Mars landing habitat will be placed along the spin axis opposite of the propulsion system.

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Other Design Considerations● Limiting the crew to a single pod eliminates the need

for pressurized arms connecting different parts of the space station. By minimizing pressurized volume we also minimize mass and can use an open truss design for the arms.

● A single living pod design also minimizes the amount of time that astronauts spend moving radially in the station. While moving parallel to the spin axis, astronauts can avoid the disorienting effects of Coriolis acceleration.

● The space habitat has been oriented so that most of the astronauts movement within the habitat will occur in the direction parallel to the spin axis.

● The station uses articulating solar panels to maximize efficiency depending on trajectory/ mission profile.

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x

yz Direction of Travel

Final Layout

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Space Habitat

Fuel Tanks

Solar Arrays

Airlock

Articulating Communications

Antenna

Propulsion System

Attitude Thruster

Integrated Equipment Assembly (IEA)

Thermal Radiators

x

z Unused Docking Mechanism

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Final Layout

Mars habitat propulsion module

Mars crewed lander

Mars habitat module

Airlock

Space habitat module

Z

yx

Truss Axial Load Calculations

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0 m

Capsule mass: 3500 kg

Hab mass: 32700 kg

Truss mass: Volume*Density

11 m26.725 m37.5m

F = m·ω²·r where ω = 0.578 rad/s∑ F = FCapsule + FHab + FTrussFCapsule = 3500·37.5·ω2 = 43900NFHab = 32700·26.725·ω2 = 287000NFCapsule + FHab = 330000N

● To take axial stress only, how massive would the truss have to be? For A6061 aluminum, tensile strength (σt) = 124–290 MPa and ρ = 2700 kg/m3.

● Summing the mass of the truss members as a function of radius (r), we find MTruss= 80·π·r2·ρ + 40·π·2.487·r2·ρ;

● If one truss panel takes ¼ of the axial load, then the reaction forces at the attachment points are:

FAx + FBx = (FHab+ FCapsule+ MTruss·ω²·*11)/4 ● Let σt = 124MPa, and solve for r.

r = 0.0103m Solve for MTruss= 161.5 kg

● The truss could be relatively small if only looking at axial stress.

A

B

FAx

FBx

x

Airlocks● Based on main ISS EVA airlock● Space Suits Stored in airlock

structure● One astronaut exits at a time

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Floor Design● Two floors, the furthest normalized

to the target acceleration● “Bottom Floor” RB = 29.325 m● “Top Floor” RT = 26.725 m● Gravity Gradient:

○ (RB-RT)/RT= 9.72%

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Floor design● Spherically curved floor to ensure

astronauts stay parallel to the acceleration vector at all times

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Bottom Floor Layout● Waste Management● Personal Hygiene● Exercise Equipment● Airlock exit

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Storage

Storage

Personal Hygiene

Waste Management

Exercise

Exercise

Airlock

Ascent to Top Floor

Top Floor Layout● Sleeping Compartments● Food Preparation● Dining/Common Space● Work Stations

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Common Area Table

Sleep Compartments

Sleep Compartments

Food Preparation

Descent to Bottom

Floor

Work Station

Work Station

Work Station

Work Station

Habitat Summary● Habitable Volume per astronaut:

20.2 m3/person● Total Storage volume: 15.6 m3

○ Food Volume: 9.3 m3

● One workspace per astronaut● Sleeping compartments double as

radiation shelters● Work stations/tables have chairs

that can be folded out of the way for use in microgravity

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Conops: Morning● Astronauts will typically wake up at 6:00 AM GST. From there they will have

about 30 minutes of post sleep. Which gives them time to review their day, read news, speak with friends family, and other morning rituals.

● The next hour is for the astronauts to enjoy their breakfast and continue with other post sleep routines giving them a little free time.

● 7:30 AM everyday is the morning Daily Planning Conference (DPC). Astronauts will sync with Earth to get daily work schedule and update on mission progress.

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Conops 7:45 AM - 6:15 PM● The astronauts will do their 2.5 hours of exercise during this time. They will do

both cardio and weight intensive exercises.● 15 minutes of daily medical evaluation. Being in artificial gravity can take a toll

on the body, and different people may respond differently over the course of the mission. Having daily evaluations allows the astronauts to stay ahead of any issues, should they arise.

● They will have 6.25 hours of work a day. This would be cleaning the spacecraft, doing experiments, repairs, etc.

● Typically after their last exercise astronauts will have time to bathe and do other hygienic upkeep.

● In groups of two, there will be a 1 hour lunch break during this time. 25

Conops: Evening● At 6:15 PM astronauts will begin their evening conference preparation. This

time will be used for them to plan their next day given what they accomplished today.

● At 7:00 PM the evening DPC will begin. The crew will discuss what they had accomplished for the day discuss everything they prepared for the next day.

● After the conference with earth, the crew will have until 9:30PM as free time. Here they can have dinner together, talk with family, and enjoy the views.

● 9:30 PM the crew will go to sleep. This will give them 8.5 hours of rest.

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Crew Schedule

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Mission Operations● All orbital maneuvering, excluding station-keeping, would be done by the

hybrid solar-electric/chemical propulsion stage● Solar-electric propulsion is used for earth-mars (or other) interplanetary

trajectories● Chemical propulsion is impulsive and used for maneuvering between orbits in

the Earth-Moon system or between orbits in the Moon-Phobos/Deimos system.

● In orbit assembly will be done in a circular orbit about Earth approximately a 400 km high - this is the nominal orbital altitude the spacecraft will stay at

● All missions will accordingly start from this orbit

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Station-keeping● The station will go into orbit in 2029, around the time of the new solar maximum in 2031● Using [4], for a 400km altitude near solar maximum, the orbit will decay 500m/s per

year.● Using 500m/s per year for all years (conservative assumption), using hydrazine attitude

thrusters with an Isp=240sec [5], and using (propellent fraction=1-e-ΔV/Isp*g) indicates per year the amount of propellent burned should be 0.15% of the total station mass

● The total station mass should be in the 100s of mT range, meaning at maximum the station should require around 200kg of hydrazine per year

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Resupply Schedule● The station will be supplied 4 times a year, similar to ISS, so each supply mission would carry 3

months worth of supplies● 2 astronauts will be on-station at a time, along with a year's worth of redundant supplies (food,

water, propellant)● For 3 months (91.3days) the crew would require 113kg of food, 76.25kg of water, and 200/4=50kg of

propellant.● This yields 239kg of consumables per resupply mission● Initial supply to the station would bring a year’s and 3 month’s worth of supplies for a total load of

988kg, which can be brought along with the astronauts on the initial Falcon 9 launch. ● Crews will be rotated out every 3 months while in LEO

○ Arrive with every resupply mission○ Crews will also be rotated out after a mission (but not during missions)

● 2-day overlap between crews yields an extra 2 days worth of supplies or 7.3kg (2.5kg of food, 4.8kg of water)

● So a resupply mission would carry a total of 246.3kg of supplies30

Moon Mission● Moon has slightly elliptical orbit with earth (e=0.0549, a=384,000km) so can consider

spacecraft-moon rendezvous at either perigee (363,000km) or apogee (405,000km).● Perigee

○ Starting at 400km altitude circular earth orbit, and using a Hohmann transfer, ΔV for trans-lunar injection is 3.08km/s

○ Hyperbolic excess velocity when entering the moon’s sphere of influence is 0.2km/s○ ΔV to slow down into a 100km altitude lunar orbit is 0.685km/s, resulting in a total ΔV of 3.76km/s○ TIme of flight is about 5 days

● Apogee○ Starting at 400km altitude circular earth orbit, and using a Hohmann transfer, ΔV for trans-lunar injection is

3.09km/s○ Hyperbolic excess velocity when entering the moon’s sphere of influence is 0.180km/s○ ΔV to slow down into a 100km altitude lunar orbit is 0.684km/s, resulting in a total ΔV of 3.77km/s○ Time of flight is about 13 days.

● ΔV for descent and propulsive landing of the lander is 2.71km/s

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Moon Mission Scheduling● Perigee mission has smaller delta-v and shorter

time of flight, so it will be used.● Need moon distance to be less than or equal to

363,000km at approach. ● From [3] in 2030 see table to right● Dates are dates of lunar closest approach minus

5 days (time of flight)

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Month Day

March 16

April 13

May 12

June 9

July 8

October 23

November 20

December 19

Moon Mission Profile● A mission to the moon would be for a 100-day stay on the moon● Would feature a 4 person crew

○ 2 to stay aboard the spacecraft ○ 2 to complete mission objectives on the moon

● For a 100-day stay, a two-person crew would require 124kg of food and 600kg of water (200kg potable, 400 grey water) for 724kg of consumables.

● Two persons in the station would also require 724kg consumables for 100 days

● All 4 persons would require 10 days worth of consumables (72.4kg) for the 5-day journey to the moon and the 5-day transit back.

● Adding in 30% of contingency food (30 days) yields 941.2kg of consumables. ● If the hab has a mass of 32.7 MT, and the hab requires 3286kg of life support

equipment, this yields a payload mass of 36927kg. 33

Tentative Propellent Mass Calculations● Landing ΔV=2.71km/s ● Assent ΔV=2.334km/s● ml=36927.2kg● Isp=~350sec in vacuum

○ Conservative estimate but moon does have a small amount of atmosphere● Use 1=0.09 and 2=0.07

○ First stage is beefier● Yields λ1=0.72 and λ2=0.44● Total vehicle mass (with hab) is 117mT ● 85mT for just ascent stage● Descent will require 64mT of propellant and ascent 42mT of propellant ● The 81mT of vehicle structural and propellant mass could be brought to orbit with 4

Falcon 9 launches34

Spin-Up/Spin-Down Timeline

● Spin-Up occurs once final supply is finished○ Avoid docking with spinning station

● Spin-Up/Spin-Down process lasts 48 hours○ Prevent nausea and disorientation of astronauts○ Keep forces due to acceleration of station to spinning speed low

● Spin-Down occurs once in Martian Orbit○ Station stays spun down until departure

● Return Spin-Up/Spin-Down is identical35

Interplanetary Missions● For missions to the Martian surface, the surface of Phobos, or the surface of

Deimos, the craft will carry a 4-person crew, two of whom will land on the surface, two of whom will stay in the station.

● Martian missions would be conjunction class missions - total mission duration would be 1100 earth days

● Would need a total of 7364kg of food and water for the entire 1100-day, 4-person, mission.

● For a 500 day stay on the surface of an interplanetary object, two people would need 620kg of food, and 1560kg of water (520kg potable, 1040 grey) for a consumable total of 2180kg.

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Interplanetary Missions - Primary burn● In orbit assembly done at approximately a 400km high circular orbit● All interplanetary missions begin with an orbital altitude change of 340,000km

(high earth/low lunar circular orbit)○ Considering just the Earth and Moon system (excluding N-body effects) and approximating the

moon’s orbit as circular with an orbital radius of 384399km, the force of gravity on the spacecraft due to the earth balances the force of gravity due to the the moon at an orbital altitude of 339645km, which is approximates to 340,000km.

● Therefore, using a Hohmann transfer between the two circular orbits, the resulting ΔV=4.04km/s.

● Time of flight for maneuver is approximately 2 Earth Days

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Solar Electric Propulsion - First Pass● Note that due to the manner in which our solar panels

are pointed, our vehicle can only thrust radially. ● ΔV to escape from a 340,000km high circular orbit

around earth is 1.1km/s● Approximating the orbits of Earth and Mars as circular at

~1AU and ~1.5AU respectively their orbital velocities around the sun are 30km/s and 24km/s respectively.

● At bare minimum, a solar electric propulsion system must achieve a ΔV of 30-24=6km/s

● At Mars, in order to slow down to a circular orbit with a period of 5-sol (r=205,000km) ΔV=0.46km/s

● This yields a total ΔV=6.6km/s● Need mass of Mars EDL craft

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Transfer to Low Martian Orbit ● Need to transfer from high martian 5-Sol orbit to Low Martian orbit

○ Altitude-250km○ Orbital altitude of Mars Reconnaissance Orbiter ○ Allows for lower EDL ΔV

● For a hohmann transfer, apoapsis burn is 0.37km/s, periapsis burn is 1.4km/s for a total ΔV=1.7km/s. Time of flight is for transfer is ~6 Earth days

● Performed by chemical part of hybrid stage

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Martian EDL● The spacecraft velocity in a 250km

martian orbit is 3.4km/s● EDL lander needs to be able to provide

ΔV=3.4km/s● Hab structural mass is 33mT, life

support equipment mass is 3286kg, food and water consumed is 2180kg.

● Add 10% more food and water for a 50-day buffer yields 2398kg of consumables

● Total mass brought to the surface is 38mT.

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Mars Lander Mass● Need ΔV=3.4km/s and payload mass is 38mT● Cannot realistically use parachutes - 1mT limit on current systems● Must use propulsive landing scheme● Isp=325sec

○ Not as good vacuum Isp○ Better than sea level Isp ○ Accounts for presence as Martian atmosphere

● Using Isp=325sec, and =0.07, yields λ=0.27.● Stage will need 93mT of propellant for a total vehicle mass (minus the payload) of 103mT.● Lander will need to be brought with Martian hab and the rest of the vehicle to Mars (assembly in

LEO)● Ascent stage would be brought over and landed on Mars beforehand● Crew will land separately

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Solar Electric Propulsion-Second Pass● Needed accurate lander mass● Use governing equation to the right (assuming

thrusting only in the radial direction)● Need parameters of solar electric propulsion

stage○ Presume an exhaust velocity of 10,000 (Isp for ion systems

is about 1000sec)○ Use 750kW, Beginning Of Life solar arrays on Earth - solar

flux decreases as 1/r^2 as the stage gets farther and farther from the sun meaning power decreases

○ Assume 30% of the solar power goes into jet power for the ion propulsion system

○ Assume payload of system is 250mT (martian decent stage is ~100mT, rest of station is ~150mT)

○ Assume λ=0.05, =0.01, so the stage, with the payload, is 5000mT leaving 4700mT of propellant left over

● Numerically integrate for 180 days 42

Solar Electric Propulsion-Second Pass Results● Resulting time of flight is ~53 days (leave 04/24/2013

at 6am UTC), which is a substantial reduction over chemical systems

● 7.4mT of fuel burned in transit● However, solution is fairly stochastic - departure

windows are measured in hours instead of days● Solution assumes Earth and Mars are co-planar

○ Mars is actually ~2° inclined with respect to earth● For diagram to the right

○ Earth is red x○ Mars is green x○ Blue is trajectory○ Ordinate corresponds to j vector and is measured in km○ Abscissa corresponds to i vector and is measured in km

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Phobos/Deimos Mission● Approximate Deimos orbit as circular with r = 23460km

○ From 5-sol parking orbit, hohmann transfer ΔV=0.71kms/○ Time of flight is 6.5 Sol

● Approximate Phobos orbit as circular with r = 9377km○ From 5-sol parking orbit, hohmann transfer ΔV=1.14km/s○ Time of flight is 18 Sol

● In both cases, the moon’s gravity field is much weaker than Mars’ so the Martian lander is sufficient

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Life Support● CO2 scrubbing● Oxygen recycling● Atmosphere design ● Water recycling ● Water Storage● Food Storage● O2 and N2 makeup supplies ● Thermal control ● Air circulation

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CO2 Removal● Modeled after the Carbon Dioxide

Removal Assembly (CDRA) system aboard the ISS

● 4 beds each comprised of○ Desiccant bed to condition the air○ CO2 adsorbent bed to trap CO2

● Pump returns clean air to the cabin● Concentrated CO2 is routed to a carbon

reduction system which removes the Oxygen

● Remaining Carbon is discharged overboard

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Oxygen Recycling system● Primary source of Oxygen is water electrolysis● Excess hydrogen generated from electrolysis is combined with CO2 from

CDRA to produce water and methane● Water generated would go to supplement the water used in electrolysis,

excess methane would be vented into space

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Atmosphere● 101.3 kPa (21% O2) nominal atmosphere with the capacity to go to 70.3 kPa

(26% O2) to prep for landing or EVAs● 30 days of contingency O2 for a crew of 4 comes adds an additional mass of

52 kg occupying 1.58 m^3 stored outside of the habitat● Humidity is maintained between 25 and 70 percent

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Water Recycling● Water from the sink, shower, and toilet is collected and distilled by boiling the

“gray water”● Steam generated is pumped through a filter to purify the water and catch

airborne contaminants ● Sink, shower and toilet gray water has 90% of water recovered● Urine processing assembly can recover 70-80% of water content● Use Zvezda system on ISS which processes water vapor from the

atmosphere for use in electrolysis

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Water Storage● Need potable water for food and drink, as well as hygiene water for bathing,

laundry● 2 L/crew member per day of potable water and 4 L/crew member per day of

hygiene water● Initially have 2000 kg (3 m3) of potable water, 4000 kg (6 m3) of hygiene water

Given 80% water content recovered, 30 day backup storage of water after 1100 days

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Food Storage● Dehydrated food to reduce weight● 0.63kg per person per day -> 1500 kg for a 4 person crew for 1100 days with

10% contingency food● Water waste will be purified using the water purification system onboard● Solid waste will be released into the vacuum of space

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Fire Suppression● Smoke detection sensors are placed in the ventilation system● CO2 extinguisher for each floor are provided● Crew is given oxygen masks to use when suppressing fire● In case of fire, air flow will be turned off to reduce the spreading of smoke

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Thermal Control of Habitat● Ammonia is circulated around the external part of the station through

radiators, cooling the habitat atmosphere and equipment. Ammonia will be used because it meets all of NASA’s thermal performance and safety requirements

● A body mounted radiator will be used for heat rejection, as well as passive multilayer insulation across the habitat’s surface

● Fans are used to circulate the air throughout the spacecraft to reduce the possibility of hot and cold air zones

● Hatch heaters and blankets

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Life Support Masses● CO2 scrubbing: 141 kg● O2 generation: 244 kg● O2 storage: 124 kg● N2 storage: 303 kg● Thermal control: 1126 kg● Water recycling: 552 kg● Waste collection: 184 kg

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● Potable water: 2000 kg● Hygiene water: 4000 kg● Dehydrated food: 1500 kg● Food storage: 612 kg● Clothing: 317 kg

Total mass: 11,103 kg

Design/Build/Test/Evaluate● Look into mitigation of the effects that dust from a planetary surface would

have on performance of an airlock.● Test the layout of an airlock for EVAs on the surface of Mars in the Neutral

Buoyancy Tank.● Compare multiple designs to see what is good in microgravity (like a moon)

v. free space v. Mars surface.

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References1. Anderson et al, “Life Support Baseline Values and Assumptions Document”

NASA, March 20152. “Advanced Life Support Requirements Document” Crew and Thermal

Systems Division, NASA, February 20033. Braeunig, R. A., Rocket Propellants, from http://braeunig.us/space/propel.htm4. Coker and Knox, “A 1-D Model of the 4 bed Molecular Sieve of the Carbon

Dioxide Removal Assembly” Comsol Conference, Boston, 20155. “Falcon 9 Launch Vehicle Payload User’s Guide revision 2” SpaceX

Cooperation, October 21 20156. Griffin, Smitherman and Howe, “Internal Layout for a Cis-Lunar Habitat,” AIAA

SPACE Forum, September 201356

References Continued7. Hall “Artificial Gravity in Theory and Practice” University of Michigan, Ann

Arbor, Michigan, July 20168. Moon Distances for Washington DC, USA - District of Columbia.from

https://www.timeanddate.com/astronomy/moon/distance.html?year=2029&n=263

9. Joosten, “Preliminary Assessment of Artificial Gravity Impacts to Deep-Space Vehicle Design” NASA, February 2007

10. May, S. (2015, June 08). Eating in Space, from https://www.nasa.gov/audience/foreducators/stem-on-station/ditl_eating

11. Ref: Alan C. Tribble, The Space Environment Princeton University Press, 1995

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References Continued12. Sherif and Knox, “International Space Station Carbon Dioxide Removal

Assembly” January 200513. Simon et all, “NASA’s Advanced Exploration Systems Mars Transit Habitat

Refinement Point of Departure Design” 2018 Rascal Competition Document14. ‘Space Launch System Mission Planner’s Guide” NASA, April 12, 2017

Render Images

15. https://images5.alphacoders.com/336/336625.jpg16. https://68.media.tumblr.com/7e45eeb73818324f554b8c02c92795c9/tumblr_nl

kssf4EAH1qd479ro1_500.jpg17. https://www.nasa.gov/mission_pages/msl/images/index.html

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References Continued 18. https://www.nasa.gov/multimedia/imagegallery/image_feature_83.html19. https://www.nasa.gov/content/satellite-view-of-the-americas-on-earth-day

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