AAE 450 SENIOR DESIGN WEEK 03 PRESENTATIONS€¦ · 3 parth shah | apm launch mission timeline o...

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AAE 450 SENIOR DESIGN WEEK 03 PRESENTATIONS 1/30/2014

SCHEDULE

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§  8:32 – Parth S. §  8:40 – Krista G. §  8:46 – Michael C. §  8:52 – Jose Miguel B. -‐  End Morning Session §  10:32 – Spenser G. §  10:38 – Ryan A. §  10:44 – Hani K. §  10:50 – Ben F. §  10:56 – Jessica C. §  Break

§  11:12 – Eric M. §  11:18 – Cameron H. §  11:24 – Erik S. §  11:30 – Andrew E. §  11:36 – Arika A. §  Break §  11:52 – Finu L. §  11:58 – Bryan F. §  12:04 – Eric F. §  12:10 – Tas Powis §  12:16 – Divinaa B. §  12:22 – Joe A.

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PARTH SHAH | APM LAUNCH MISSION TIMELINE

o  ORGANIZATIONAL ITEMS o  LAUNCH MISSION TIMELINE

1/30/2014

ORGANIZATIONAL ITEMS

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§  Vehicle Specifications Spreadsheet •  Moved to Google Drive •  Link uploaded to Wiggio •  More to come from Andrew after presentations

§  Project Name Submissions •  Minimum 1 idea per person •  Please submit by Sunday •  Some really great ideas so far!

§  Document your progress •  LaTeX tutorial being prepared by Andrew •  Start writing now!

Parth Shah | APM

VEHICLE OPERATION REGIMES

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§  Launch Vehicles •  Ground to LEO •  3 Classes: – Light: Atlas V/ Titan IV – Medium: Falcon 9 / Heavy – Heavy: SLS / Falcon XX

§  Cargo and CTV •  LEO to Lunar Orbit

§  Cargo Lander and CTV Lander •  Lunar Orbit to Lunar Surface

Parth Shah | APM

LAUNCH MISSION TIMELINE

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§  Initial estimates: •  From R & D completion to Crew Landing on the Moon •  All Low-‐Energy Transfers estimated at ~100 days •  Launch Start Date: 1/1/2020 •  Visual Representation [Gantter]

Parth Shah | APM

What # Launches # Per colony Max Transfer Time (days)

Launch Vehicle Type

Comm Sats 2 n/a 100 Light Construc-on Bots 3 2 100 Heavy

Habitats & Construc-on Mat. 6 2 100 Heavy Rovers & Science 12 4 100 Heavy

CTV & Consumables 3 1 4 Medium Resupply 12 4 100 Heavy

KRISTA GARRETT | MISSION DESIGN LUNAR LANDER 1/30/2014 o  LUNAR DESCENT ΔV AND PROPELLANT o  LUNAR ASCENT PROPELLANT o  OTHER WORK: MISSION PROFILES FOR PRESSURIZED ROVER

AND LAUNCH VEHICLE SELECTION

LUNAR LANDER DESCENT

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§  Lander from CTV orbit to Powered Descent IniEaEon: •  Hohmann transfer to 15.24 km alEtude •  Δv = 39.0 m/s

§  Propellant for descent:

*Calculated using initial mass of 45.0 Mg and an Isp of 311 s **Value scaled from Lunar Module descent propellant requirements2

Hohmann Transfer* 0.57217 Mg

+ Powered Descent** 30.77332 Mg

= Total 31.34549 Mg

Krista GarreS 1/30/2014

LUNAR LANDER ASCENT §  Lander from lunar surface to CTV orbit

•  Vertical rise phase + single-‐axis rotation + Powered Explicit Guidance

§  Propellant for ascent: •  9.51527 Mg*

§  Total propellant: •  40.86076 Mg •  Option of refueling on the moon

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*Value scaled from Lunar Module ascent propellant requirements2

Fig. 1: Ascent Profile (From Ref. 1 by Kos, Polsgrove, Sostaric, Braden, Sullivan, Lee; reproduced)


REFERENCES

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1Kos, Polsgrove, Sostaric, Braden, Sullivan, Lee, “Altair Descent and Ascent Reference Trajectory Design and Initial Dispersion Analyses,” American institute of Aeronautics and Astronautics meeting, Toronto, Canada, 2010.

2Bennett, Floyd V., “Apollo Lunar Descent and Ascent Trajectories,” NASA TM X-‐58040, 1970.


MICHAEL CREECH| MISSION DESIGN PROPELLANT BUDGET AND BLOCK CONSTRUCTION 1/30/2014

o  PROPELLANT BUDGET FROM LEO TO LUNAR ORBIT o  REGOLITH HEATING AND BLOCK FORMATION

PROPELLANT BUDGET

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§  Propellant costs per item §  Analysis of 3 propellants §  Total propellant cost from

LEO to Lunar Orbit

Michael Creech | Mission Design

Propellant Cost Item LOX/LH2 [kg] LOX/RP-‐1 [kg] LOX/methan [kg] L2 Comm Sat 4947.43 15869.32 13795.54 1 Moon Sat 4947.43 15869.32 13795.54 3 Moon Sats Total 14842.30 47607.97 41386.63 1 Tracking StaEon 659.66 2115.91 1839.41 4 Tracking StaEons Total 2638.63 8463.64 7357.62

Total Single Habitat 164914.46 528977.49 459851.45 Garage 6596.58 21159.10 18394.06 Single Unit: Water & Storage Container 99882.09 320380.51 278513.63 Shielding Support Material 230880.25 740568.48 643792.03 Food 255393.14 819195.70 712144.35 EMU suit 379.30 1216.65 1057.66 Total 8 Suits 3034.43 9733.19 8461.27

Cargo Vehicle (Dry Weight) 89053.81 285647.84 248319.78 Lunar Lander 146114.22 468674.06 407428.38

Transport (inert) 39579.47 126954.60 110364.35 Crew Capsule (inert) 14512.47 46550.02 40466.93 Crew 2308.80 7405.68 6437.92 Food 369.41 1184.91 1030.07 Water 1920.92 6161.53 5356.35

1 Heavy Pressurized Rover 32982.89 105795.50 91970.29 2 Heavy Pressurized Rovers Total 65965.79 211591.00 183940.58 1 Light Pressurized Rover 29684.60 95215.95 82773.26 4 Light Pressurized Rovers Total 118738.41 380863.79 331093.04 1 Tire (4 per heavy & light rover) 1649.14 5289.77 4598.51 24 Tires Total 39579.47 126954.60 110364.35

1 Helper Robot 16491.45 52897.75 45985.14 6 Helper Robots Total 16491.45 52897.75 45985.14

Total 1404558.02 4505242.03 3916503.27 Sources -‐Item Masses Compiled by Erik SleSehaugh

REGOLITH CONSTRUCTION

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§  Use geothermal reaction with aluminum powder §  Regolith Brick

•  20 x 10 x5 [cm] •  1.5 [kg] •  .15 [m] of NiChrome wire •  0.004224 W for heating

§  Habitat Costs •  33,500 bricks •  16,500 kg of aluminum powder •  141.504 W of power


ASSUMPTIONS

14 Michael Creech | Mission Design

AssumpEons

Fuel Isp [sec] MR ΔV [km/s] g_o [km/s^2] Propellant FracEon

LOX/LH2 448 3.413718 5.3942 9.81E-‐03 0.9

LOX/RP-‐1 311 5.863031

LOX/methane 321 5.548724

Sources -‐ΔV Calculated by Thomas Rich

EQUATIONS

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§  Ideal Rocket Equation

§  Propellant Mass


∆𝑣= 𝐼↓𝑠𝑝 𝑔↓0 ln(𝑀𝑅)

𝑀↓𝑝𝑟𝑜𝑝𝑒𝑙𝑙𝑎𝑛𝑡 = 𝑀↓𝑝/𝑙  (𝑀𝑅−1)/𝑀𝑅− 𝑀𝑅−1/λ  

CONSTRUCTION PARAMETERS

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§  Resistivity of NiChrome = 1.1e-‐6 [Ωm] §  Density of Regolith = 1.5 [g/cm^3] §  Required current for heating = 24 [A]


REFERENCES

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§  NASA – “The J-‐2X Engine: NASA’s New Upper Stage Engine” §  Space Launch Report – “SpaceX Falcon 9 Data Sheet” §  Burkhardt, Sippel, Herbertz, Klevanski – “Comparative Study of Kerosene

and Methane Propellant Engines for Resuable Liquid Booster Stages” §  NASA – “Space Launch System (SLS) Fun Facts” §  Faierson, Logan, Steward, Hunter – “Demonstration of concept for

fabrication of lunar physical assets utilizing lunar regolith simulant and a geothermite reaction”


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JOSE MIGUEL BLANCO/ MISSION DESIGN LOW THRUST TRAJECTORIES

o  BASIC RESULTS FOR FUEL CONSUMPTION IN LOW THRUST TRAJECTORIES USING ELECTRIC PROPULSION

1/30/14

LOW THRUST TRAJECTORIES FOR CARGO VEHICLE

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§  Electric propulsion •  High specific impulse(Isp):

–  Isp electric propulsion from 1 000 to 10 000 s –  Isp liquid propellant around 455 s

(http://ccar.colorado.edu/asen5050/projects/projects_2009/stansbury/)

•  Low thrust output –  Increased mission time – Special trajectory design (Hohmann and other types of direct transfers can not be implemented)

§  First approximation: relative 2 body problem (planar) •  + acceleration due to thrust •  + fuel consumption

Jose Miguel Blanco / Mission Design / Cargo and Lander

FUEL CONSUMPTION

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§  Propagate numerically in Matlab [1] •  100 Mg cargo vehicle •  1 year •  Spacecraft initially parked in LEO at an

altitude of 170 km •  Constant tangential acceleration •  Reach Moon orbit

§  Results [2] •  Isp = 5000 s m_fuel/m_total=12.96% •  Isp = 9000 s m_fuel/m_total=7.42% Liquid propellant m_fuel/m_total=71.81% (Numbers for liquid propellant consumption assuming Hohmann transfer, from Thomas Rich (MD) slides 01/23/2014)


NEXT STEPS

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§  Try different Thurst vectoring strategies §  Non constant acceleration §  Introduce the moon

•  First by using patched relative 2 body motion •  Maybe look into 3 body dynamics


[1] NUMERICAL PROPAGATION

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§  2 body relative motion propagation using MATLAB •  Initial conditions

– Circular orbit at LEO (170km above Earth surface) Vc= 7.8021 [km/s] –  Initial mass of the vehicle of 100 Mg

•  Constant tangent acceleration relative to the position vector.

•  Iterate until the distance relative to Earth is similar to that of the Moon –  To reach a distance of 387420 km (Moon distance from the Earth assumed to be 384400) an acceleration of a_theta=2.2*(10^-‐5)*g0 =2.1582e-‐07[km/s] is needed


[2] RESULTS

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§  The results obtained from the numerical propagation were: •  For a conservative value of Isp of 5000 s the fuel consumption is 12.96 Mg

•  Whereas if we used an optimistic value for the specific impulse of 9000 s, the fuel consumption would be 7.42 Mg

§  From Thomas Rich presentation in 01/23/2014: •  Total mass of the vehicle = 302.2 Mg •  Liquid propellant mass =217 Mg •  m_fuel/m_total=71.81%


ACKNOWLEDGEMENTS

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I would like to thank Thomas Rich whose numbers I used for comparison and Frank Laipert for pointing me in the right direction


GLOSSARY

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Isp = specific impulse m_total = initial mass of the vehicle including fuel m_fuel = mass of fuel needed to complete the

mission = fuel flow rate in kg per second


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SPENSER GUERIN | CONTROLS CARGO VEHICLE

o  NAVIGATION SYSTEM REQUIREMENTS o  PRELIMINARY EXTERNAL TORQUE ASSESSMENT AND MODEL

1/30/2014

NAVIGATION SYSTEM REQUIREMENTS

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§  Does not include processing power consumption.

§  Can be extended to other vehicles; different requirements depending on colony.

Spenser Guerin | Controls

SENSOR POWER [W] MASS [kg] VOLUME [] OPERATING TEMPERATURES [°C]

IMU 22 4.5 7206 -‐30→65

Star Tracker 12 3.5 7888 -‐20→50

AlEmeter 15 3 3540 N/A

GNSS Receiver 0.4 1 18 -‐40→85

PRELIMINARY TORQUE MODEL

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§  Environmental forces affect attitude control. §  Forces due to aerodynamic, magnetic, and

micro-‐meteorite forces negligible. §  Currently developing simple gravity-‐gradient

model.

Spenser Guerin | Controls

Vehicle Offset from Radial Vector

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RYAN ALLEN | CONTROLS CREW TRANSPORT VEHICLE

o  PRELIMINARY NON-GRAVITATIONAL FORCE MODEL o  IN-TRANSIT DELTA V CORRECTIONS

1/30/2014

NON-GRAVITATIONAL FORCE MODEL

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§  Solar electromagnetic radiation, reflected radiation, and non-‐propulsive mass expulsion forces included in model.

§  Model intended to combine with environmental torque calculations.

Ryan Allen | Controls

Reflected Solar Radiation parameter definition

PRELIMINARY DELTA-V CALCULATIONS

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SOURCE FORCE [N] ADDITION. ΔV [m/s]

Solar RadiaEon 8.936e-‐5 4.022e-‐04

Reflected Solar RadiaEon (Earth) 7.726e-‐6 3.477e-‐05

Reflected Solar RadiaEon (Moon) 5.733e-‐7 2.580e-‐06

Mass Expulsion (Leaks) 7.625e-‐7 3.4315e-‐06

§  Model for CTV flight from LEO to Lunar Orbit. §  Can be applied to all launch vehicles given

mass and time in transition orbit.



§  %Analysis of Nongravitational Forces and Space Environmental Torques

§  %Ryan Allen, AAE450 Spacecraft Design §  %Version 1.0, 1/28/14 §  %Version 1.1, 1/29/14 §  §  %Forces Accounted for: Solar Radiation, Reflected Solar

Radiation, §  %Non-‐propulsive leaks §  %Torques Accounted for: TBD §  §  %Input Parameters §  clear all; close all; clc;



§  %All values initially for earth for testing purposes: §  §  k_elm = 1.5; %Reflectivity constant [dim.] §  A = 13.2; %Effective spacecraft area [m^2] §  f0 = 1353; %Solar flux constant [W/m^2] §  rss = 1; %Sun-‐spacecraft distance [AU] §  f_elm = f0/(rss^2); %Energy flux [W/m^2] §  c = 299792458; %Speed of Light {m/s] §  r = 149597870700; %Sun-‐earth distance [m] §  rm = 146692378; %Sun-‐moon distance [m] §  Rp = 6378100; %Radius of planet Earth [m] §  Rpm = 1737400; %Radius of moon [m] §  rps = 384400000/2; %Planet-‐spacecraft distance [m] §  a = .11; %Albedo constant [dim] §  m_dot = 5.4*10^-‐10; %Mass flow rate [kg/s] §  R = 2077; %Gas constant for pressurization gas [J/kg-‐K] §  T0 = 300; %Stagnation temperature [K] §  k = 1.667; %Specific heat ratio [dim]



§  %Calculations §  Fsolar = k_elm*A*f_elm/c §  Frefe = pi*Rp^2*f0/(r^2) §  Frefm = pi*Rpm^2*f0/(r^2) §  F_leak = m_dot*(2*R*T0*(1+k)/k)^.5

HANI KIM | HUMAN FACTOR PRESSURIZED ROVER 1/30/2014 o  H.F. : PERSONAL ITEMS & CLOTHING & WASTE o  P.R. : MASS PER PERSON

HABITAT

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§  Personal Item •  0.4 x 0.33 x 0.2 m (1) •  Volume : 0.027 •  Mass restriction : neglect •  Estimated mass : 15kg/crew (2)

§  Clothing (research area) •  Skinsuit (3) + ordinary clothes – Mineral density loss from weight bearing bones – Painful spine elongation (up to .7 cm)

Figure 1: Personal carry-‐on dimension

WASTE & PRESSURIZED ROVER

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§  Waste (4) •  Possible uses :

–  Shielding –  Energy

Pressurized Rover Available space : 1.678 m^3

Name Volume (m^3) Mass (kg) Crew 0.079 80.000

Food 0.060 (5) 0.617 (6)

Oxygen 0.007 16.000

Water 0.004 3.524 (6)

Feces 0.008 0.109 (6)

Urine 0.004 3.869 (6)

Total 0.213 104.119

•  Note •  Value per person •  Only related to human survival (CommunicaEon + other device TBA)

Figure 2: Pressurized dimension by Eric

Feces Name Wet weight (kg/person/day) Dry weight (kg/person/day)

Feces 0.096 – 0.132 0.03

General composiEon Substance Approximate percent

Dead bacteria 14 -‐ 30 Fats 10 -‐ 20

Inorganic maSer 10 -‐ 20 Protein 2-‐3

Food residues 25 -‐ 40

APPENDIX

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(1) Based on average underseat bags in current market. (2) Long, W.A, Fischer, W.E., and Brunelle, R.J., “Framing system for aircraft passenger seat”, United States Patent, 4,375,300, Mar. 1, 1983. (3) Waldie, J.M., and Newman, D.J., “A gravity loading countermeasure skinsuit”, Acta Astronautica, vol. 68, issues 7-‐8, pp. 722-‐730. (4) Wydeven, T., and Colub, M.A., “Waste Streams In A Crewed Space Habitat”, Waste Management & research, Vol.9, 1991, pp. 91-‐101. (5) Size of 4.2 cu ft compact refrigerator (6) Carrasquillo, R., “ISS Environmental Control and Life Support System(ECLSS) Future Development for Exploration”, 2nd Annual ISS Research and Development Conference, NASA Headquarters, Denver, CO, 2013. (7) Guyton, A.C., Textbook of Medical Physiology, W.B. Saunders Company Philadelphia, PA, 1981. (8) Diem, K., and Lentner, C., Scientific Table, Basle, 1970.

APPENDIX

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Waste (4), (7), (8)

Name Wet weight (kg/person/day) Dry weight (kg/person/day)

Toilet waste

Urine 1.5 – 2.11 0.059

Feces 0.096 – 0.132 0.03

Toilet paper 0.494 x

Hygiene water

Laundry 12.5 x

Shower/hand-‐wash 5.5 x

Dish-‐wash 5.4

Others

Trash 0.816 -‐ 1.162 0.593 – 0.845

Total 26.306 – 27.298 0.682 -‐ 0.934

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BEN FISHMAN | HUMAN FACTORS PRESSURIZED ROVERS 1/30/2014 o  LIFE SUPPORT/OXYGEN

HF: PRESSURIZED ROVER (LARGE)

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§  Air mixture in atmosphere is comprised of: •  78% -‐ Nitrogen (N2) •  21% -‐ Oxygen (O2) •  ~1% -‐ Water vapor & other inert gases

§  Average human consumes 0.64 kg of O2 / day [1] •  Value changes based on energy exerted by crew

§  Sensors that take readings of rover environment will automatically release correct amount of each component

§  To supply air to rover, we will use pressurized industrial cylinders

§  Carbon dioxide scrubbers will filter and vent CO2 to space via bottom of rover [2]

Ben Fishman | Human Factors


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Type of Work Number of Crew

Amount of Air (24 hrs.)

Mass of Tanks* (kg)

Volume of Tanks (m3)

AcEve 2 433 L 10.7 0.37

Intense 2 570 L 11.8 0.44

AcEve 4 (emergency)

866 L 22.2 0.62

Intense 4 (emergency)

1140 L 25 0.84

*Includes mass of elements at ~ 200 PSI [3] •  Air supply fits into alloSed

volume required by rover team

•  Spacesuits oxygen will be supplied by burning powdered sodium chlorate (future topic) [4]


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§  %% AAE 450 §  %% Oxygen Calculations for Habitat §  %% Written by Ben Fishman §  §  clear all; close all; clc §  %% Pressurzed Rover Calculations §  §  % 78% nitrogen §  % 21% oxygen §  % 1 % water vapor §  §  %%%%%% Insert Number of Expected Passengers %%%%%% §  num_people = 4; §  %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% §  §  §  %%%%%%%%% Values for Liters of Air/min %%%%%%%%%%% §  rest = 6; §  inter = 28; §  hard = 50; §  %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% §  §  §  %%%%%%%% By weight %%%%%%%%%%%% §  mass_o2_rest = .64 * num_people; §  mass_o2_inter = .84 * num_people; §  mass_o2_hard = 7.2 * num_people; §  §  conversion = 1.251/1.429; %this is done by density §  Ben Fishman | Human Factors


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§  mass_n2_rest = .64 * num_people * conversion * 1.78; §  mass_n2_inter = .84 * num_people * conversion * 1.78; §  mass_n2_hard = 7.2 * num_people * conversion * 1.78; §  %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

§  %%%%%% Total Calculations by Weight %%%%%%%% §  tot_air_rest = mass_o2_rest + mass_n2_rest; §  tot_air_inter = mass_o2_inter + mass_n2_inter; §  tot_air_hard = mass_o2_hard + mass_n2_hard; §  %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% §  §  %%%%%%% Liter Calculations %%%%%%%%%%%%%%%% §  o2_liter_rest = mass_o2_rest * 142.9 §  o2_liter_inter = mass_o2_inter * 142.9 §  n2_liter_rest = mass_n2_rest * 125.1 §  n2_liter_inter = mass_n2_inter * 125.1 §  §  %%%%%%%%%%%5%%%% OUTPUT %%%%%%%%%%%%%%%%%%%%%%%% §  fprintf('Rest: Need %2.1f kg of O2 and %2.1f kg of N2 to have a total of %2.1f kg of air per day for %1.0f people

\n',mass_o2_rest,mass_n2_rest,tot_air_rest,num_people) §  fprintf('Intermediate: Need %2.1f kg of O2 and %2.1f kg of N2 to have a total of %2.1f kg of air per day for %1.0f people

\n',mass_o2_inter,mass_n2_inter,tot_air_inter,num_people) §  fprintf('Hard: Need %2.1f kg of O2 and %2.1f kg of N2 to have a total of %2.1f kg of air per day for %1.0f people

\n',mass_o2_hard,mass_n2_hard,tot_air_hard,num_people) §  %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%



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1.  Hooge,O. “How much Oxygen for a person to survive in an air-‐tight enclosure?” http://members.shaw.ca/tfrisen/how_much_oxygen_for_a_person.htm

2.  Carbon dioxide scrubber. Nov, 29, 2013. http://en.wikipedia.org/wiki/Carbon_dioxide_scrubber

3.  Airgas. http://www.airgas.com/content/details.aspx?id=7000000000234 4.  Freudenrich, Craig. “How Space Suits Work.”

http://science.howstuffworks.com/space-‐suit1.htm


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JESSICA CALLINAN | HUMAN FACTORS CTV SUPPLY AND CLOTHING

o  HUMAN FACTORS GROUP LEADER o  CREW TRANSPORT VEHICLE FOOD AND WATER SUPPLY o  CREW SIZING AND FORCES o  CREW CLOTHING

1/30/14

CREW TRANSPORT VEHICLE | CREW

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§  For a one week food and water supply for 8 crew members[1]

§  Dragon capsule[5]

•  Currently designed to seat 7 with standing room for 3

§  Forces crew can

withstand[7] •  Launch: ~3 g’s •  Re – entry: ~7-‐8 g’s

Jessica Callinan | Human Factors

CTV Supply Mass (kg) Volume (m3)

Food 112 0.4667

Water 582.3708 2.4265

Total 694.3708 2.8932

Dragon Capsule Mass (kg) Volume (m3)

Launch 6000 25

Return (empty) 3000 11

CREW SIZING| CLOTHING

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§  Height of crew[3]

§  Weight of crew[2]:

§  Clothing[4] [6]


Height (m)

Minimum 1.4859

Maximum 1.9304

Weight (kg)

Female 80

Male 95

Total (colony) 700

Clothing Mass (kg) Volume (m3)

Individual 12.1274 4.0452

Total (colony) 97.0197 32.36

TABLE OF CLOTHING MASSES AND AMOUNTS

49 Jessica Callinan | Human Factors

Clothing Type Mass (kg) Volume (m3)

Gym Shoes (2) 1.814368 -‐

Walking/comfortable shoes (2) 1.814368 -‐

T-‐shirts (8) 1.530873 -‐

Work shirts (2) 1.15 -‐

Pants (2) 1.360776 -‐

Shorts (2) 0.907184 -‐

Exercise shorts (6) 2.041164 -‐

Underwear (10) 0.3118445 -‐

Socks (10) 0.226796 -‐

Bras (2) 0.170097 -‐

Sweater (2) 0.8 -‐

Individual Total 12.1274705 4.0452

Colony Total (8) 97.019764 32.362

MATLAB CODE: CTV FOOD AND WATER-1

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§  % Jessica Callinan adapted from Taylor Schultz §  % AAE 450 Spacecraft Design §  % Last Updated: 1/29/2014 §  §  % These numbers are based off the first launch SpaceX had to the ISS §  §  % SpaceX delivered 660 kg of cargo - half of this was food. This means that §  % there was approximately 330 kg of food sent up. It was stated this equaled §  %160 meals. §  §  §  % Food Weight Requirements §  num_people = 8; % Number of people in each colony §  §  weight_food = 2; % kg of food per person per day §  §  weight_colony_food = num_people * weight_food; % Weight of food per day per

colony §  §  weight_week_food = weight_colony_food * 7 % Weight of food per week per

colony §  §  % Water Weight Requirements §  water_person = 1000 * 3.78541; % kg of water per person per year (changed

gal to kg)


MATLAB CODE: CTV FOOD AND WATER-2

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§  water_week_person = water_person/52; % kg of water per person per week § 

§  weight_week_water = water_week_person * 8 % kg of water per colony per week § 

§  total_food_and_water = weight_week_food + weight_week_water %Total mass needed for delivery

§ 

§  % Volume Requirements

§  b = 6000/330; % Scaling factor from Dragon delivery

§ 

§  tot_vol = 25; % Dragon has 25 m^3 of total launch payload volume § 

§  vol_used = tot_vol/b; % Total volume of payload area used to deliver food

§ 

§  delivery = 330; % Amount delivered on mission (used for scaling)

§ 

§  vol_food = (weight_week_food / delivery) * vol_used % Volume needed for food delivery

§ 

§  vol_water = (weight_week_water / delivery) * vol_used % Volume needed for water delivery § 

§  total_vol = vol_food + vol_water % Total volume needed for delivery


MATLAB CODE: HABITAT CLOTHING - 1

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§  %Jessica Callinan §  % AAE 450 Spacecraft Design §  % Last Updated: 1/29/2014 §  §  %Clothing Requirements §  §  %These estimates are based off an article from Space.com where Dr. §  %Aspelund, a textiles, fashion merchandising, and design professor from the §  %University of Rhode Island is quoted saying that for a 30 year voyage, §  %each person would require 100 cubic feet of clothing. I have contacted Dr. §  %Aspelund and pending response these values are subject to change. §  §  %Clothing Volume Requirements §  §  %Individual Clothing Volume Requirements §  §  clothing_person_30year_ft = 1000; %ft^3 per 30 person per 30 years §  §  clothing_person_30year_m = clothing_person_30year_ft * 0.0283168; %Converting from

ft^3 to m^3 §  §  clothing_person_year = clothing_person_30year_m / 30; %Clothing required per year

(m^3) §  §  clothing_volume_person_mission = clothing_person_year * (30/7) %Clothing required

per 4 2/7 year mission (m^3)


53

§  %Colony Clothing Volume Requirements §  §  clothing__volume_colony_mission = clothing_volume_person_mission * 8 %Clothing

required per colony per 4 2/7 year mission (m^3) §  §  §  %Individual Clothing Mass Requirements §  §  %Gym shoes - 2 lbs §  %Walking/comfortable shoes - 2 lbs §  %Average t-shirt - 6.75 oz, one for every 3 days of exercise and 1 under §  %every work shirt every 7 days §  %Average work shirt - 575 g, one for every 7 days §  %Average pants - 1.5 lbs, one for every 10 days §  %Avearge shorts - 1 lbs, one for every 10 days §  %Average exercise shorts - 0.75 lbs, one for every 3 days of exercise §  %Underwear ~ 1.1 oz, one pair for every two days §  %Socks - 0.8 oz, on pair for every two days §  %Bras ~ 3 oz, one for every week §  %Sweaters - 400g, as needed, bring 2 §  §  %This analysis will assume as much clohting as a typical astronaut would §  %take. At the moment we are assuming we will have washing capabilities and §  %will take two sets of clothes that can be interchanged during wash cycles. §  %These are assuming a 10 day cycle of clothes, and the calculation will §  %show a full 20 weeks of clothing.



54

§  gym_shoes = 2 * 0.453592 * 2; %lbs to kg, with 2 pairs of shoes §  §  walk_shoes = 2 * 0.453592 * 2; %lbs to kg, with 2 pairs of shoes §  §  tshirt = 6.75 * 0.0283495 * 8; %ounces to kg, with 4 t-shirts §  §  work_shirt = 575 * 0.001 * 2; %g to kg, with 2 work shirts §  §  pants = 1.5 * 0.453592 * 2; %lbs to kg, with 2 pairs of pants §  §  shorts = 1 * 0.453592 * 2; %lbs to kg, with 2 pairs of shorts §  §  exer_shorts = .75 * 0.453592 * 6; %lbs to kg, with 6 pairs of shorts §  §  underwear = 1.1 * 0.0283495 * 10; %ounces to kg, with 10 pairs of underwear §  §  socks = 0.8 * 0.0283495 * 10; %ounces to kg, with 10 pairs of socks §  §  bras = 3 * 0.0283495 * 2; %ounces to kg, with 2 bras §  §  sweaters = 400 * 0.001 * 2; %g to kg, with 2 sweaters §  §  clothing_mass_person_mission = gym_shoes + walk_shoes + tshirt + work_shirt... §  + pants + shorts + exer_shorts + underwear + socks + bras + sweaters §  %total mass of each persons clothing §  §  %Colony Clothing Mass Requirements §  §  clothing_mass_colony_mission = clothing_mass_person_mission * 8 %total mass of colony's

clothing


REFERENCES

55

1.  Carrasquillo, Robyn. “ISS ECLSS Future Development for Exploration.” July 2013. [http://www.nasa.gov/sites/default/files/files/issrdc_2013-‐07-‐17-‐1600_carrasquillo2013.pdf]

2.  Center for Disease Control and Prevention. “Body Measurements.” 2010. [http://www.cdc.gov/nchs/fastats/bodymeas.htm]

3.  NASA. “Astronaut Requirements.” January 2004 [http://www.nasa.gov/audience/forstudents/postsecondary/features/F_Astronaut_Requirements.html]

4.  Moskowitz, Clara. “Packing for an Interstellar Space Voyage: What to Bring?” September 2012. [http://www.space.com/17763-‐interstellar-‐spaceflight-‐clothing-‐packing.html]

5.  Spacex. “Dragon.” 2013. [http://www.spacex.com/dragon] 6.  Spector, Dina. “Here’s What Astronauts Pack When They Go To Space.”

September 2013. [http://www.businessinsider.com/what-‐astronauts-‐pack-‐when-‐they-‐go-‐to-‐space-‐2013-‐9]

7.  Wikipedia. “g-‐force.” January 2014. [en.wikipedia/wiki/G-‐force]


56

ERIC MENKE | COMMUNICATION HEAVY ROVER SHIELDING

o  REDUCTION OF ROVER VOLUME/SHIELDING MASS o  SIZE USING REGOLITH VS. EXTRACTED IRON OXIDE o  LASER COMMUNICATION

1/30/2014

HEAVY ROVER EARLY CONCEPTS

57

§  Shaved off ~4-‐5 Mg §  Al wrapped in

carbon fiber

Eric Menke | Communication

SHIELDING THICKNESS & COMM.

Eric Menke | Communication

*assuming 3 cm thick Al walls

Unrefined regolith fill

Extracted iron oxide fill

Mass 4.634 Mg

Volume 10.188 m3

Mass 3.360 Mg

Volume 4.671 m3

Laser Communication

Download Upload Ground Terminal

Laser on Vehicle

622 Mbps 20 Mbps 40 W 0.5 W

REFERENCES

59 Eric Menke | Communication

McKay, D.S., Carter, J.L., Boles, W.W., Allen, C.C., and Allton, J.H., "JSC-‐1: A New Lunar Soil Simulant," Engineering, Construction, and Operations in Space IV American Society of Civil Engineers, 1994, pp. 857-‐866.

Washington, D., "Laser Demonstration Reveals Bright Future for Space Communication ,"

NASA's Goddard Space Flight Center, December 2013.[http://www.nasa.gov/content/ goddard/laser-‐demonstration-‐reveals-‐bright-‐future-‐for-‐space-‐communication/ #.Uumq9Hfdh8E. Accessed 1/29/13.]

60

CAMERON HORTON | AERODYNAMICS SAMPLE CARRIER RE-ENTRY

o  SCIENCE SAMPLE CARRIER RE-ENTRY POD DESIGN

JANUARY 30TH, 2014

SAMPLE SIZE APPROXIMATION

61

§  Samples will be primarily of Anorthosite (An75) •  [1]Density ≈ 2.67 g/cm3 (Mg/m^3)

§  100 kg sample = 0.03745 m3 §  125 kg sample = 0.04682 m3

§  Anorthosite sample is approximately 56% the size of Stardust[2]

[1] http://adsabs.harvard.edu/full/1979LPI....10..978P [2] http://www.jpl.nasa.gov/news/press_kits/stardust-‐return.pdf

SCIENCE SAMPLE CARRIER POD

62

§  Anorthosite max sample size is 56% the size of Stardust re-‐entry pod

§  Scaled up, carrier total volume ≈ 0.103 m3 (0.6m height, 1m diameter)

§  Carrier material volume (total volume – sample volume) ≈ 0.05618 m3

§  Total mass ≈ 0.202 Mg •  (0.125 Mg for samples + 0.046 Mg for apertures/controls + 0.036 Mg for shielding

Apollo/CEV pod rendering

Gallileo pod rendering

MSL,MERS pod rendering

All models created in CATIA

HEAT SHIELDING

63

§  [1]PICA – phenolic impregnated carbon ablator (similar to shielding used on Apollo capsules)

§  [2]Withstands heating as high as 2,760 degrees Celsius (5000 degrees Fahrenheit)

§  Replaced TPS (thermal protection system) §  [3]Low density ≈ 300 kg/m3

[1] http://www.nasa.gov/centers/ames/research/msl_heatshield.html [2] http://www.nasa.gov/centers/ames/news/releases/2008/08_43AR.html [3] http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20100026827_2010028571.pdf

BACKUP SLIDES - RISK IDENTIFICATION

64

§  Risks •  Critical Launch Failure – loss of launch vehicle •  Launch Window failure – weather •  Miscalculation of payload sizing •  Miscommunication between supplier and integrator

•  ……

References: hSp://www.nasaspaceflight.com/2013/11/sls-‐us-‐proposals-‐increasing-‐payload-‐desEnaEon-‐opEons/ hSp://www.nasaspaceflight.com/2012/11/nasa-‐payload-‐fairings-‐opEons-‐mulE-‐mission-‐sls-‐capability/

65

ERIK SLETTEHAUGH| STRUCTURES LAUNCH VEHICLE & SATS

o  PAYLOAD MASS PROPERTIES o  PAYLOAD FAIRING DIMENSIONS o  NUMBER OF LAUNCHES AND TIMELINE

1/30/2014

MASS PROPERTIES - PAYLOAD

66 Erik Slettehaugh | Launch Vehicle & Sats

Vehicle Group Items Volume (m^3) Mass (Mg)

Launch Vehicle & Com Sat4 Sats, 4 Tracking

Stations 27.63 8.40

Habitat & Sample Return VehicleHabs, Storage,

Shielding Support, Food, Water

873.51 230.64

Cargo Vehicle & Moon Lander Cargo Vehicle, Lander

225.00 71.30

Crew Transport Vehicle Transporter, Crew Capsule

75.00 13.29

Pressurized Rovers 2 Heavy & 4 Light Rovers, 24 Tires

147.65 68.00

Helper & Construction Robots 6 Helper Robots 51.60 7.20Single Colony 1325.16 384.03

3 Colonies & SATs 3991.11 1158.09

Estimated Totals

Total Values

Single Colony

PAYLOAD DIMENSIONS & LAUNCHES


Height = 19.1 m

Diameter = 8.4 m

Volume = 1058.48 m^3

Logistical Launch Timeline

Payload Fairing for SLS

Launches by Mass

# of Launches

2 6 3 12 6 6 3 3 6 47

Payload Sats

Helper Robots

& Cargo Vehicle

Tracking Stations

& Building Material

Habs, Shielding Support

Rovers Food & Water

Transport, Crew

Capsule, & Lander

Sample Return

Resupply Total Launches

Launch Vehicle

Mass Payload to TLI (Tans-lunar Injection) (Mg)

Mass Payload to LEO (Mg)

Mass (Mg)

Volume (m^3)

# of Launches by Mass

SLS 4-engine 38.10 129.73 1158.09 3991.11 30.40

Required PayloadLaunch Capability

BACKUP SLIDES - MASS PROPERTIES


Estimated Vehicle Group Contact Date Last Updated Notes Item Volume (m^3) Mass (kg)

Launch Vehicle & Com Sat Ian Bennett 1/28/2014 L2 COM Sat 3.91 1500.00 Ian Bennett 1/28/2014 1 Moon Sat 3.91 1500.00 Ian Bennett 1/28/2014 2 Moon Sat 3.91 1500.00 Ian Bennett 1/28/2014 3 Moon Sat 3.91 1500.00 Ian Bennett 1/29/2014 4 per colony 4 Tracking Stations 4.00 800.00 Ian Bennett 1/29/2014 1 Tracking Station 1.00 200.00 Sat Total 15.63 6000.00 Total 27.63 8400.00

Values for a Single Colony

Habitat & Sample Return Vehicle

1/23/2014

All going to be shipped in 1 or 2 units

Sleeping 70.00 ? 1/23/2014 Personal/Work 20.00 ? 1/23/2014 Showering/Toilet 8.00 ? 1/23/2014 Eating/Meeting 7.50 ? 1/23/2014 Food Storage/Preparation 7.50 ? 1/23/2014 Food Growing 2.50 ? 1/23/2014 Lab 63.00 ? 1/23/2014 Recreation/Exercise Room 90.00 ? 1/23/2014 Water filtration ? ? 1/23/2014 Air Lock 8.00 20.00

Andrew Emans 1/28/2014 This is not based on the above info Total Single Habitat (8ppl) 454.00 50000.00

Andrew Emans 1/28/2014 Garage 40.00 2000.00

1/28/2014 Single Unit: Water & Storage Container? 30.28 30283.00

Andrew Emans 1/28/2014 Shielding Support Material 9.00 70000.00

1/28/2014 Volume is for one container only Food (1 Colony) 322.63 77432.00

1/28/2014 EMU suit (1 suit) 2.20 115.00 1/28/2014 Total 8 Suits 17.60 920.00 Total 873.51 230635.00



Arika Armstrong 1/28/20145m diameter, 13m

long Cargo Vehicle (Dry Weight) 200.00 27000.00

Scott Sylvester 1/29/2014 Includeds: Inert, fuel

Lunar Lander 25.00 44300.00

Total 225.00 71300.00

Scott Sylvester 1/29/2014 Transporter (inert) 50.00 7500.00Scott Sylvester 1/29/2014 Crew Capsule (inert) 25.00 4400.00Scott Sylvester 1/29/2014 Crew NA 700.00Scott Sylvester 1/29/2014 Food NA 112.00Scott Sylvester 1/29/2014 Water NA 582.40

Total 75.00 13294.40

Saagar Unadkat 1/28/2014 1 Heavy Pressurized rover 17.83 10000.00Saagar Unadkat 1/28/2014 1 Heavy Pressurized rover 17.83 10000.00Saagar Unadkat 1/28/2014 1 Light Pressurized rover 10.00 9000.00Saagar Unadkat 1/28/2014 1 Light Pressurized rover 10.00 9000.00Saagar Unadkat 1/28/2014 1 Light Pressurized rover 10.00 9000.00Saagar Unadkat 1/28/2014 1 Light Pressurized rover 10.00 9000.00

Saagar Unadkat 1/28/2014 1 Tire (4per heavy&light rover) 3.00 500.00

Saagar Unadkat 1/28/2014 Total 24 tires 72.00 12000.00

Total 147.65 68000.00

Cargo Vehicle & Moon Lander

Crew Transport Vehicle

Pressurized Rovers



Arika Armstrong 1/29/2014 1 Helper Robot 8.60 1200.00 Arika Armstrong 1/29/2014 1 Helper Robot 8.60 1200.00 Arika Armstrong 1/29/2014 1 Helper Robot 8.60 1200.00 Arika Armstrong 1/29/2014 1 Helper Robot 8.60 1200.00 Arika Armstrong 1/29/2014 1 Helper Robot 8.60 1200.00 Arika Armstrong 1/29/2014 1 Helper Robot 8.60 1200.00

Total 51.60 7200.00Single Colony 1325.163 384029.400

3 Colonies & SATs 3991.114 1158088.2Total Values

Helper & Construction Robots

Launch Vehicle

Mass Payload to TLI (Tans-lunar

Interjection) (kg)

Mass Payload to LEO (kg)

Radius (m) (Could be 5m) Height (m)

Volume (m^3) (Rough Estimates) (Assume cylinder) (About 9 School

buses)

Required Payload Mass (kg)

Required Payload Volume (m^3)

# of Launches by mass

# of Launches by volume

SLS 4-engine 38100.00 129727.00 4.20 19.10 1058.48 1158090.00 3991.11 30.40 3.770612274SLS 2-engine 39700.00 4.20 19.10 1058.48 1158090.00 3991.11 29.17 3.770612274SLS 1-engine 38500.00 4.20 19.10 1058.48 1158090.00 3991.11 30.08 3.770612274

BACKUP SLIDES - RISK IDENTIFICATION

71

§  Risks •  Critical Launch Failure – loss of launch vehicle •  Launch Window failure – weather •  Miscalculation of payload sizing •  Miscommunication between supplier and integrator

•  ……

Erik Slettehaugh | Launch Vehicle & Sats

References: hSp://www.nasaspaceflight.com/2013/11/sls-‐us-‐proposals-‐increasing-‐payload-‐desEnaEon-‐opEons/ hSp://www.nasaspaceflight.com/2012/11/nasa-‐payload-‐fairings-‐opEons-‐mulE-‐mission-‐sls-‐capability/

DATE

72

ANDREW EMANS | STRUCTURES LUNAR RESOURCES

o  HABITAT SHIELDING o  REGOLITH AS A CONSTRUCTION MATERIAL

30 JAN 2014

HABITAT SHIELDING

73

§  Total shield mass needed for a standing 13m/13m/3m structure varies from 673Mg – 749Mg.

§  Burying a 15m/30m/5m carbon fiber structure with a 10 degree slope ramp to the surface reduces the amount of material needed from Earth to:

Mass: 112 Mg Volume: 70 m^3 §  The structure must be capable of supporting at least

3.2 kPa . §  Less dense material is preferable since dense

materials can amplify the effects of radiation.

LUNAR MATERIAL §  Lunar glass (LG)

•  LG is relatively easy to produce on the Moon and allows for the collection of gases (H2, N2, C, He, Ar, S) and metals (iron, aluminum, titanium).

•  Volume for thread manufacturing ~2 m^3. •  Fiberglass can be used in new 3D printers to create building blocks,

spare parts, tools, etc. §  Lunar Bricks

•  Sintering releases oxygen. •  Automated lunar brick maker has already been built ~ 3 m^3.

§  Lunarcrete •  Stronger than Earth concrete due to the porosity of regolith but

requires more water to cure for the same reason. •  Resin can be applied to the surface of lunar dust to create stiff

pathways and hold cave walls together. Using resins in the habitat caves would save ~40 Mg of mass from Earth per shelter.

74

BACK-UP: REFERENCES

75

Allen, C.C., Graf, J.C., McKay, D.S., “Sintering Bricks on the Moon,” Engineering, Construction, and Operations in Space IV, American Society of Civil Engineers, 1994, pp. 1220-‐1229. [http://ares.jsc.nasa.gov/HumanExplore/Exploration/EXLibrary/DOCS/EIC049.HTML. Accessed 1/20/14.]

Armstrong, T. W., Parnell, T.A., Watts, J.W., “Radiation Effects and Protection for Moon and Mars

Missions,” NASA. [http://science.nasa.gov/media/medialibrary/1998/05/11/msad28apr98_1a_resources/cosmic.pdf. Accessed 1/25/14.]

Blacic, J.D., “Mechanica Properties of Lunar Materials Under Anhydrous, Hard Vacuum Conditions:

Applications of Lunar Glass Structural Components,” Lunar Bases and Space Activities of the 21st Century, Washington DC, 1984. [http://library.lanl.gov/cgi-‐bin/getfile?00261768.pdf. Accessed 1/21/14.]

Ethridge, E.E., Tucker, D.S., “Processing Glass Fiber from Moon/Mars Resources,” NASA. [http://

science1.nasa.gov/media/medialibrary/1998/05/11/msad28apr98_1a_resources/fiber.pdf. Accessed 1/25/14.]

Lin, T.D., Love, H., and Stark, “Physical Properties of Concrete Made with Apollo 16 Lunar Soil Sample”

Space Manufacturing 6, Proceedings of Eighth Princeton/AIAA/SSI Conference, 1987, pp. 361-‐366. [http://www.nss.org/settlement/moon/library/LB2-‐515-‐ConcreteMadeFromApollo16sample.pdf. Accessed 1/20/14.]

76

ARIKA ARMSTRONG/STRUCTURES CARGO SHIP SIZING

o  CARGO SHIP SIZING & UPDATES o  MATERIAL INVESTIGATIONS o  ISSUES WITH REGOLITH o  MACHINERY SIZING

01/30/14

CARGO SHIP DESIGN

77

§  “Disposable” one-‐way ship from Earth to moon for cargo

§  Lander enclosed within or attached outside §  5 m diameter x 13 m long, Aluminum walls §  Dry mass: 21.23 Mg §  Pressurized volume: 198 m³ §  Cargo mass: 20.3 Mg

Arika Armstrong | Structures

COMBINED CARGO/LANDER DESIGN

78

§  Reuse lander in habitat •  Garage •  Rooms •  Storage

§  Control systems/thermal control included §  Dry mass: 35 Mg §  Pressurized volume: 150 m³

Arika Armstrong | Structures

79

FINU LUKOSE | PROPULSION COMMUNICATION SATELLITES; ROVER MOTORS

o  COMMUNICATION SATELLITE PROPULSION o  ROVER MOTOR SIZES

01/30/2014

COMMUNICATION SATELLITES

80

§  Communications satellites of comparable size propulsion systems (i.e. MRO, LRO) •  Medium sized thrusters, tank volume, mass •  Assumption of using hydrazine monopropellant; common monopropellant for satellites

•  Look into greater detail on electric propulsion

Finu Lukose | Propulsion

Parameter Value

Propellant Tank Volume 1.2 m3

Propellant Mass 1200 kg Power (Valve, controls) ~0.5kW

ELECTRIC MOTORS SIZING

81

§  Electric motors of Manned Lunar Rover, scaling based on existing motors

§  Similar system to LRV envisioned •  Per-‐wheel electric drive unit


Parameter Value

Volume (Space for motor) 4 m3

Mass (of motor system) 80 kg [per wheel]

Power 15kW 4 wheel scheme: 3.8 kW 6 wheel scheme: 2.5 kW

LUNAR ROVING VEHICLE

82

§  Unshielded and light for transport of two passengers

§  Mass of vehicle ~210 kg §  Capable of additional payload ~490 kg §  Each wheel DC series wound motor ~190 W at

10,000 rpm •  Maneuvering by DC motor ~75 W

§  Existing system used as possible scale


WORKS CITED

83

§  Houghton, Martin B.; Tooley, Craig R.; Saylor, Richard S. (July, 2007). Mission Design and Considerations for NASA’s Lunar Reconnaissance Orbiter.

§  Garulli, Andrea; Giannitrapani, Antonio; Leomanni, Mirko; Scortecci, Fabrizio. (November 2011). Autonomous Lower Earth Orbit Station Keeping With Electric Propulsion.

§  VIA Motors. (2014). The New E-‐REV Powertrain from VIA Motors.

§  Viotti, Michelle. (Jan, 2014). Mars Reconnaissance Orbiter – Spacecraft Parts: Propulsion. Retrieved from http://web.archive.org/web/20060331051038/http://mars.jpl.nasa.gov/mro/mission/sc_propulsion.html

Finu Lukose| Propulsion

84

PROPULSION SCIENCE SAMPLE RETURN

o  BRYAN FOSTER o  ELECTROMAGNETIC LAUNCHERS

1/29/2014

RAILGUN CONCEPT

85 Bryan Foster | Science Sample Return

Mass (kg)

Volume (m^3)

Force (kN)

Current (kA)

Voltage (V)

Power (MW)

Energy (MJ)

Launches

250 .0821 1187.94 1399.75 374.26 523.88

261.94

1

125 .0410 593.97

1032.12 275.97 284.83

142.41

2

25 .0082 118.79

518.81

138.72 71.97

35.98

5

10 .0033 47.51

357.29

95.53 34.13 17.06 25

5 .0016 23.75

273.43

73.11 19.99

9.99

50

Lunar Escape Velocity 2375.89 m/s

CURRENT TECHNOLOGY

86

§  Navy prototype: 3.2 kg, 10.4 MJ, 2500 m/s §  Fire 6-‐10 rounds §  Apollo Missions: 20-‐110 kg of samples §  Orion module capacity: 100 kg

Bryan Foster | Science Sample Return

ASSUMPTIONS

87

§  Semi-‐infinite conductive rails §  Length of 10 meters §  Rail radius of 0.02 meters §  Sample density will be an average density of

the 3 main lunar rock types §  Launch will take half a second §  Railgun dimensions will depend on the

chosen sample size


DERIVATION OF THE FORCE EQUATION

88

§  𝐹=𝐼𝐿𝑥𝐵 §  Lorentz force equation §  𝐵(𝑠)= 𝜇↓0 𝐼/2𝜋𝑠  §  Biot-‐Savart law §  𝐵(𝑠)= 𝜇↓0 𝐼/2𝜋 ( 1/𝑠 + 1/𝑑−𝑠 ) §  𝐵↓𝑎𝑣𝑔 = 1/2𝑑 ∫𝑟↑𝑑−𝑟▒𝐵(𝑠)𝑑𝑠  = 𝜇↓0 𝐼/4𝜋𝑑 ∫𝑟↑𝑑

−𝑟▒( 1/𝑠 + 1/𝑑−𝑠 )𝑑𝑠 = 𝜇↓0 𝐼/2𝜋𝑑 ln 𝑑−𝑟/𝑟   §  𝐹=𝐼𝑑𝐵↓𝑎𝑣𝑔 = 𝜇↓0 𝐼↑2 /2𝜋 ln 𝑑−𝑟/𝑟  


MATLAB CODE

89

§  %Bryan Foster §  %AAE 450 §  %Rail Gun sizing code §  §  clc §  clear §  close all §  §  %% §  %Escape Velocity §  mmoon = 7.34767309*10.^22; %kg §  rmoon = 1737.4; %km §  G = 6.67384*10.^-‐11; %m^3/kg s^2 §  ve = ((2*mmoon*G)/(rmoon*1000)).^(1/2) %m/s §  %% §  %Force §  m = [250, 125, 25, 10, 5]; §  for i = 1:5 §  F(i) = m(i).*ve./.5 %N §  end


MATLAB CODE CONT.

90

§  %% §  %Amps §  mu0 = 4.*pi.*10.^-‐7; %V s/A m §  rho = 3046; %kg/m^3 §  r = .02; %m §  for j = 1:5 §  Vol(j) = m(j)/rho; §  d(j) = (Vol(j)).^(1/3); §  I(j) = ((F(j).*2.*pi)./(4.*pi.*(10.^(-‐7)).*log((d(j)-‐r)./r))).^(1/2) %Amperes §  end §  %% §  %Voltage §  rescopper = 1.68.*10.^-‐8; %Ohms m §  L = 20; %m §  A = (r.^2)*pi; %m^2 §  Rc = rescopper.*(L./A) §  for k = 1:5 §  Volt(k) = Rc.*I(k) %Volts §  end §  %% §  %Power §  for c = 1:5 §  Power(c) = Volt(c).*I(c) %Watts §  Energy(c) = Power(c).*.5 %Joules §  end


ERIC A FLORES | PROPULSION CREW TRANSPORT VEHICLE

o  PRELIMINARY ENGINE ANALYSIS o  PRELIMINARY FUEL AND OXIDIZER TANK ANALYSIS

1/30/14

ENGINE ANALYSIS Like the Old Days?

92

Engine Fuel Oxidizer

Lunar Lander

Descent Stage

TR-‐201 (TRW) Aerozine-‐50 (50% Hydrazine and 50%

UDMH)

Nitrogen Tetroxide (N2O4)

Lunar Descent Module Engine (LDME)

Aerozine-‐50 (50% Hydrazine and 50%

UDMH)


Ascent Stage

RS-‐ 18 (In Development -‐ Rocketdyne) Liquid Methane (CH4) Liquid Oxygen (O2)

Lunas Ascent Module Engine (LAME)


UDMH)


Transport 2nd Stage J-‐2 (Rocketdyne) Liquid Hydrogen Liquid Oxygen (O2)

Same Fuel and

Oxidizer!

PROPELLANT TANKS

93

8.4 m

8.4 m 8.4 m

Fuel

Oxidizer

Fuel

Vehicle Stage Total Mass Propellant (includes Mass Engine)

[Mg]

Total Space Volume [m^3]

Lunar Lander

Descent Stage

31.3455 30.197

Ascent Stage 9.515 9.167

Transport 2nd Stage 97.913 1165.94686

Total Lunar Lander Propellant Volume

[m^3] 39.364

Total Tranport Propellant Volume

[m^3] 1165.94686

ATTACHMENTS

94

§  http://www.braeunig.us/space/comb-‐OH.htm (Mixture Ratio of L.Hydrogen and L.Oxygen)

§  http://www.braeunig.us/space/comb-‐NA.htm (Mixture ratio of Aerodine-‐50 and Nitrogen Tetroxide)

§  https://engineering.purdue.edu/AAE/Research/Propulsion/Info/rockets/liquids (Good Engine Tables)

§  https://engineering.purdue.edu/AAE/Research/Propulsion/Info/rockets/liquids (Delta V for LEO to LLO)

§  http://airandspace.si.edu/explore-‐and-‐learn/topics/apollo/apollo-‐program/spacecraft/saturn_v.cfm (SLS Stages Info)

§  http://www.braeunig.us/space/propel.htm (good analysis)

ATTACHMENTS

95

Vehicle Stage Engine Fuel Oxidizer Fuel Density (kg/m^3)

Oxidizer Density (kg/m^3)

Isp (sec)

Lunar Lander

Descent Stage

TR-‐201 (TRW) Aerozine-‐50 (50% Hydrazine and 50%

UDMH)


903 1450 303

Lunar Descent Module Engine

(LDME)


UDMH)


903 1450 311

Ascent Stage Lunas Ascent Module Engine

(LAME)


UDMH)


903 1450 311

Transport 2nd Stage J-‐2 (Rocketdyne) Liquid Hydrogen Liquid Oxygen (O2) 70.85 1141 421

ATTACHMENTS

96

Mass Engine (kg)

Total Mass Propellant

(includes Mass Engine)[kg]

Mixture RaEo

Mass of Fuel (kg)

Mass of Oxidizer (kg)

Volume of Fuel (m^3)

Volume of Oxidizer (m^3)

Total Space Volume (m^3)

136.078

179 31345.490 1.9 20536.700 10808.790 22.743 7.454 30.197

82 9515.270 1.9 6234.142414 3281.128 6.904 2.263 9.167

1788.100 97912.829 5 81594.024 16318.805 1151.644664 14.30219534 1165.946859

Delta-‐V (LEO to LLO)** [m/s] 4040

g0 [kg*m/s^2] 9.81

Total Mass CTV Dry (Includes Lunar Propellant) [kg] 58995

CTV Total Mass (Complete) [kg] 156907.8

ATTACHMENTS

97

Delta-‐V (LEO to LLO)** [m/s] 4040

g0 [kg*m/s^2] 9.81

Total Mass CTV Dry (Includes Lunar Propellant) [kg] 58995

CTV Total Mass (Complete) [kg] 156907.8

98

ANDREW POWIS | POWER/THERMAL HEAVY ROVER POWER SYSTEMS

o  VEHICLE TEAM LEADER: PRESSURIZED ROVERS

o  FUEL CELLS AND BATTERIES

1/30/2014

HEAVY ROVER LAYOUT AND MISSION PROFILE

99 Andrew Powis | Power/Thermal

System Mass (Mg) Volume (m3) Power (kW)

Power 3.5 6.5 0

Shielding/Structures 20 5.3 0

Life Support 0.2 1 0.5

Drive Train 4 14 16 (max)

Mechanisms 0.5 1 1

Tolerance Factors 1.15 1.15 1.3

Total 32.4 32.0 17.5

Shielding (regolith)

Cabin

Eric Menke & Andrew Powis Krista GarreS

POWER SYSTEMS


Proton Exchange Membrane Fuel Cell (PEMFC)

Reactant/Product Storage Mass (Mg)[1,2] Storage Volume (m3)[1,2]

Oxygen 1.18 1.50

Hydrogen 1.78 3.92

Water 0.493 0.493

Fuel Cell Power Source

Lithium-‐Ion Bahery Reactant/Product Storage Mass (Mg)[4] Storage Volume (m3)[4]

BaSery 6.58 2.66

Total Power for Mission 476.9 x 2 kWh

FUEL CELL TYPES

101

§  PAFC: Phosphoric acid (electrolyte) fuel cell. §  PEMFC: Proton exchange membrane fuel cell. §  SOFC: Solid oxide fuel cell. §  MCFC: Molten carbonate (electrolyte) fuel cell. §  AFC: Alkaline (electrolyte) fuel cell. §  DMFC: Direct methanol fuel cell.

Andrew Powis | Power/Thermal

FUEL CELL POWER RANGES


Based on the above table and details from [1], the appropriate fuel cell types for consideraEon are the PAFC, PEMFC and SOFC types. Below is a comparaEve analysis of the qualitaEve properEes of each of these fuel cell types.

FUEL CELL COMPARISON


Cell Type Pros Cons PAFC •  Developed and understood

technology. •  Highly reliable and robust.

•  High temperatures required. •  Considerable size/mass penalty for

hydrogen reforming system. •  Hydrogen storage.

PEMFC •  Solid and immobile electrolyte.

•  Simple and robust design. •  Low operaEng temperature. •  The new preference at NASA. •  Quick start up Eme.

•  Hydrogen storage. •  Difficult to manage water build up

in electrolyte.

SOFC •  High reacEon rates (due to temperature).

•  Very high operaEng temperature (. •  Complex cooling system. •  Lengthy start up Eme.

Table 1: Advantages and Disadvantages of the three appropriate fuel cell types indicated from Figure 1 [1].

FUEL CELL CHOICE

104

§  Despite the power requirements required for the Heavy Rover, the best fuel cell option appears to be the proton exchange membrane fuel cell PEMFC, due to the simplicity and robustness of this design. The infrastructure supporting this power, primarily consists of fuel storage (in particular hydrogen), whereas other cell types require alternative systems to reform and manage reactants.

§  The PEMFC also enjoys a wide range of operating powers, allowing the output power to be scaled depending on the application and required output at any given time. Therefore this technology will be compatible for use in the Light Rover, reducing the maintenance, knowledge and tools required to design, construct and maintain the rovers on Earth and the Moon.

§  The primary issue with the PEMFC (and indeed all of the above fuel cell types) is the storage of reactants. Alternatives for storing hydrogen are explored below.


HYDROGEN STORAGE


Storage Method Storage Efficiency Notes Gas •  Simplest method.

•  Indefinite storage Eme. •  No purity limits on the Hydrogen. •  Disadvantage in weight penalty.

Liquid •  Compact and comparaEvely light compared to gas storage. •  Store hydrogen as LH2 and skim gas as it evaporates. •  Not an indefinite storage Eme. •  Requires heavy cryogenics.

Metal Hydride •  Safe due to the endothermic nature of the hydrogen release reacEon.

•  Good volumetric efficiency. •  Long refill Emes. •  Exhaust gas is non recoverable.

Carbon Nano-‐Tubes Unknown •  SEll undergoing research.

Methanol •  Similar to metal hydride storage. •  Exhaust gas is non-‐recoverable.

Alkali Metal Hydrides •  Similar to metal hydride storage. •  Necessary to dispose of poisonous exhaust.

Table 2: Storage efficiency and comments regarding hydrogen storage methods for fuel cell consumpEon [1].

CHOICE OF HYDROGEN STORAGE

106

§  For preliminary design of the heavy rover power system the gas storage method was chosen. This was due to the robust and inert nature of the storage method. Ideally this method will also allow reversal of the fuel cell process (electrolysis) to recover pure oxygen and hydrogen, effectively allowing the fuel cell to be treated as a battery.

§  The proposed storage method is a tank with an aluminium inner liner, around which is wrapped a composite of aramid fibre and epoxy resin. The material has high ductility, giving it good burst behaviour and a storage efficiency of 3.1%. Furthermore this storage design has a volume efficiency of 0.071 𝑚↑3  𝑝𝑒𝑟 𝑘𝑔 𝑜𝑓 𝐻↓2 

§  At this stage however, alternative storage options will not be ruled out, in the case that the storage efficiency of the hydrogen is insufficient.


OXYGEN STORAGE

107

§  Oxygen storage efficiency was gauged from commercially available oxygen tanks from Luxfer [2]. Each tank has an inner liner made from aluminium alloy and strengthened with carbon-‐fiber. The storage efficiency for oxygen at

3000 𝑝𝑠𝑖 is 37.2% and the volume or storage is,

§  0.00342 𝑚↑3  𝑝𝑒𝑟 𝑘𝑔 𝑜𝑓 𝑂↓2 


BATTERY PERFORMANCE


Property Value Energy Density Specific Energy Operalng Life deep cycles Temperature range to

Lithium-‐ion bahery performance. Space rated lithium-‐ion baSeries have been used in various applicaEons for decades. One of the most recent examples of such an applicaEon is the baSery used with the Mars Science Laboratory on Curiosity. The performance of this baSery was obtained from the manufacturers website [4] and key parameters are tabulated below.

Table 3: ProperEes of the Mars Science Laboratory lithium ion baSeries.

REFERENCES

109

§  [1] Larminie, J., Dicks, A., Fuel Cell Systems Explaine, John Wiley & Sons, West Sussex, 2003.

§  [2] “L6X composite cylinder specifications”, Luxfer Gas Cylinders, Boston,

MA. [http://www.luxfercylinders.com/products/medical-‐cylinders/195-‐l6x-‐carbon-‐composite-‐full-‐wrap-‐medical-‐cylinders?units=1. Accessed 1/29/2014]

§  [3] “Densities and Molecular Weights of Some Common Gases”, The

Engineering ToolBox, [http://www.engineeringtoolbox.com/gas-‐density-‐d_158.html. Accessed 1/29/2014]

§  [4] “Lithion, Inc. – World Class and Beyond”, Yardney Technical Products,

Inc. [http://www.yardney.com/Lithion/lithion.html. Accessed 1/27/2014]


MISSION PROFILE

110

§  % Code written by Krista Garrett to demonstrate speed and power profiles §  % for the pressurized rover on the moon. §  % AAE 450 §  % Written January 24, 2014. §  §  %Modified to represent Heavy Rover mission profile §  %Andrew Powis January 28, 2014. §  §  clear; clc; close all; §  §  % Calculate the maximum speed of the rover on flat terrain by assuming the maximum

power is consumed when the % pressurized rover is traveling up an incline of 30 degrees a a speed of 10 km/hr

§  §  % Set the friction coefficient for the lunar surface §  friction_coeff_lunar_dirt = 0.7; % unitless (coefficient) §  §  % Set the mass of the pressurized rover in kg §  mass_rover = 30500; % kg §  §  % Set the acceleration due to lunar gravity §  lunar_gravity = 1.622; % m/s^2


MISSION PROFILE

111

§  §  % % % §  % The equations for Power used in this code: §  % power = speed*mass*gravity*(friction_coeff_lunar_dirt*cos(incline_angle) +

sin(incline_angle)) §  % % % §  §  % Find power needed to get the rover up an incline of 30 degrees at a speed of 10

km/hr (10000 m/hr) §  §  velocity_up_30_incline = 10; % km/hr §  §  % Need to convert this to m/s §  velocity_up_30_incline_meters_per_second = velocity_up_30_incline*1000/(60*60); % m/

s §  §  % Calculate the maximum power §  max_power =

velocity_up_30_incline_meters_per_second*mass_rover*lunar_gravity*(friction_coeff_lunar_dirt*cosd(30)+sind(30))

§  §  % Assign the range of the rover: §  range_rover = 50000; % m


MISSION PROFILE

112

§  % Create a matrix for the distances: §  step_size = 1; §  distance_traveled = [0:step_size:range_rover]; % MAY NEED TO CHANGE THE STEP SIZE §  §  % The distance for each leg is given in meters §  §  % Set the initial incline and elevation §  incline(1) = 0; elevation(1) = 0; §  §  j = 2; % Counter §  §  % % % --------------------------------------------------------------------- §  % This while statements can be copied and pasted to create the desired §  % terrain for the mission. §  % % % --------------------------------------------------------------------- §  §  while distance_traveled(j) <= 5000; §  incline(j) = 11.3; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j+1; §  end


MISSION PROFILE

113

§  while distance_traveled(j) <= 10000; §  incline(j) = -11.3; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j+1; §  end §  §  while distance_traveled(j) <= 20000; §  incline(j) = 0; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j + 1; §  end §  §  while distance_traveled(j) <= 24000; §  incline(j) = 30; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j + 1; §  end

§  while j <= 28000 §  incline(j) = -30; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j + 1; §  end § 


MISSION PROFILE

114

§  while j <= 30000 §  incline(j) = 30; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j + 1; §  end §  §  while j <= 34000 §  incline(j) = 0; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j + 1; §  end §  §  while j <= 36000 §  incline(j) = -30; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j + 1; §  End §  while j <= 40000 §  incline(j) = 0; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j + 1; §  end §  §  while j <= length(distance_traveled) §  incline(j) = 5.71; % degrees §  elevation(j) = elevation(j-1) + tand(incline(j))*step_size; % m §  j = j + 1; §  end


MISSION PROFILE

115

§  % % % --------------------------------------------------------------------- §  % Calculate the velocity achieved by the rover during each phase §  % % % --------------------------------------------------------------------- §  for k = [1:1:length(distance_traveled)] §  speed_achieved_with_max_power(k) = max_power/

(mass_rover*lunar_gravity*(friction_coeff_lunar_dirt*cosd(incline(k)) + sind(incline(k))));

§  speed_in_kilometers_per_hour(k) = speed_achieved_with_max_power(k)*(60*60)/1000; % km/hr

§  end §  §  max_power_array = max_power*ones([1,length(distance_traveled)]); §  §  % % % --------------------------------------------------------------------- §  % Assume a constant speed of 10 km/hr and calculate the power usage. §  % % % --------------------------------------------------------------------- §  §  constant_speed = 10000/(60*60); % 10 km/hr converted to m/s §  constant_speed_array_km_per_hour = 10*ones([1,length(distance_traveled)]); §  §  for m = [1:1:length(distance_traveled)] §  power(m) =

constant_speed*mass_rover*lunar_gravity*((friction_coeff_lunar_dirt*cosd(incline(m)) + sind(incline(m))));

§  end §  §  time_step = step_size/constant_speed; %s §  total_power = sum(power.*time_step) %kWh


MISSION PROFILE

116

§  % % % --------------------------------------------------------------------- §  % PLOTS §  % % % --------------------------------------------------------------------- §  §  % To plot distance in kilometers, convert from meters to kilometers. §  distance_traveled_kilometers = distance_traveled./1000; % km §  §  % Plots when the rover uses maximum power §  figure(1); §  §  subplot(3,1,1); plot(distance_traveled_kilometers, elevation); §  grid on; xlabel('Distance (km)'); ylabel('Elevation (m)'); title('Elevation

Profile'); §  §  subplot(3,1,2); plot(distance_traveled_kilometers, speed_in_kilometers_per_hour); §  grid on; xlabel('Distance (km)'); ylabel('Speed (km/hr)'); title('Speed of Rover'); §  §  subplot(3,1,3); plot(distance_traveled_kilometers, max_power_array); §  grid on; xlabel('Distance (km)'); ylabel('Power (W)'); title('Power Usage'); §  §  % Plots when the rover travels at a constant speed §  figure(2); § 


MISSION PROFILE

117

§  subplot(3,1,1); plot(distance_traveled_kilometers, elevation,'LineWidth',2); §  grid on; xlabel('Distance (km)','fontsize',14); ylabel('Elevation (m)','fontsize',

14); title('Elevation Profile','fontsize',16,'FontWeight','bold'); §  §  subplot(3,1,2); plot(distance_traveled_kilometers,

constant_speed_array_km_per_hour,'color','green','LineWidth',2); §  grid on; xlabel('Distance (km)','fontsize',14); ylabel('Speed (km/hr)','fontsize',

14); title('Speed of Rover','fontsize',16,'FontWeight','bold'); §  §  subplot(3,1,3); plot(distance_traveled_kilometers, power,'color','red','LineWidth',

2); §  grid on; xlabel('Distance (km)','fontsize',14); ylabel('Power (W)','fontsize',14);

title('Power Usage','fontsize',16,'FontWeight','bold'); §  ylim([0, 12*10^4])


PEMFC CODE

118

§  %AAE450 - Spacecraft Design §  %Andrew Powis §  §  %For a given power requirement (Wh) this script determines the required §  %storage weight and volume of reactants for a Proton Exchange Membrane Fuel §  %cell. §  §  %Refer to provided Fuel Cell Notes for references. §  §  clear; §  clc; §  close all; §  §  %%%%%%%%%%%%%%%%%%%%%%%INPUTS%%%%%%%%%%%%%%%%%%%%%%%%%% §  §  %Required energy (Ws) §  Pe = 1.717e+09; §  §  §  %%%%%%%%%%%%%%%%%%%%%%%%CONSTANTS%%%%%%%%%%%%%%%%%%%%%%% §  §  %Faraday's Constant (C/mol) §  F = 96485.3399;


PEMFC CODE

119

§  %Atomic Mass of O2 §  mO2 = 2 * 15.9994; §  §  %Atomic Mass of H2 §  mH2 = 2 * 1.00794; §  §  %Atomic Mass of water §  mH2O = 18.015; §  §  %%%%%%%%%%%%%%%%%%%%%%VARIABLES%%%%%%%%%%%%%%%%%%%%%%%% §  §  §  %Storage efficiency for H2 by weight. [1] §  meffH2 = 0.031; §  §  %Storage efficiency for O2 by weight [2]. §  meffO2 = 0.37196; §  §  %Storage efficiency for H2 by volume, m^3/kg [1]. §  veffH2 = 0.070967742; §  §  %Storage efficiency of O2 per volume, m^3/kg [2]. §  veffO2 = 0.003417077;


PEMFC CODE

120

§  %Output voltage (V) inherent to PEMFC, improve accuracy with a better estimate. §  Vc = 0.65; §  §  %Number of cells §  ncells = 10; §  §  %%%%%%%%%%%%%%%%%%%CALCULATIONS%%%%%%%%%%%%%%%%%%%%%%%%%%% §  §  %Seconds of operation §  %Secs = Hrs * 3600; §  §  %Output Current (Amps) §  I = Pe / (Vc * ncells); §  §  §  %O2 usage (moles/s) §  O2_moles = Pe / (4*Vc*F); §  §  %O2 usage (kg/s) §  O2_kg = O2_moles * mO2 * 10^-3; §  §  §  %H2 usage (moles/s) §  H2_moles = Pe / (2*Vc*F); §  §  %H2 usage (kg/s) §  H2_kg = H2_moles * mH2 * 10^-3;


PEMFC CODE

121

§  %Water produced (moles/s) §  H2O_moles = Pe / (2*Vc*F); §  §  %Water produced (kg/s) §  H2O_kg = H2O_moles * mH2O * 10^-3; §  §  §  %Heating rate (W) §  heat = Pe * (1.25 / Vc - 1); §  §  §  %Oxygen required for entire mission. §  O2 = O2_kg; §  §  %Mass of oxygen and storage equiptment. §  O2_storage_kg = O2 / meffO2 §  §  %Volume of oxygen storage. §  O2_storage_m3 = O2*veffO2 §  §  §  %Hydrogen required for entire mission §  H2 = H2_kg §  §  %Mass of oxygen and storage equiptment. §  H2_storage_kg = H2 / meffH2


PEMFC CODE

122

§  %Volume of hydrogen storage. §  H2_storage_m3 = H2 * veffH2 §  §  §  %Water produced over entire mission. §  H2O = H2O_kg §  §  %Water storage volume. §  H2O_storage_m3 = H2O_kg/1000 §  §  Total_reactant_mass = H2_storage_kg+O2_storage_kg §  Total_storage_vol = H2O_storage_m3+H2_storage_m3+O2_storage_m3


PEMFC CODE

123

§  %AAE450 - Spacecraft Design §  %Andrew Powis §  §  %For a given power requirement (Wh) this script determines the required §  %storage weight and volume of a lithium-ion battery. §  §  clear §  clc §  close all; §  §  %%%%%%%%%%%%%%%%%%%%%%VARIABLES%%%%%%%%%%%%%%%%%%%%%%%% §  §  %Required power (kWh) §  Pe = 477; §  §  %Hours of operation §  Hrs = 20; §  §  %Lithium-Ion Battery capacity, based on Curiosity Rover battery (Wh/kg) §  cap = 145; §  §  %Lithium-Ion Battery size, based on Curiosity Rover battery (kWh/L) §  vcap = 358;


PEMFC CODE

124

§  %%%%%%%%%%%%%%%%%%%CALCULATIONS%%%%%%%%%%%%%%%%%%%%%%%%%%% §  §  %Battery mass §  mbat = (Pe*1000)/cap §  §  %Battery volume §  vbat = Pe/vcap §  §  %Battery density §  dbat = mbat/vbat; § 


125

DIVINAA BURDER|POWER/THERMAL POWER SYSTEMS/RE-ENTRY

o  LUNAR FISSION SURFACE POWER SYSTEMS o  RADIOISOTOPE THERMOELECTRIC GENERATORS (RTG) o  RE-ENTRY CAPSULE SHIELDING

January 30,2014

FISSION REACTORS & RTGS Fission Surface Power Systems Ø  How it Works

•  Splitting uranium atoms in a reactor to generate heat that is converted into electrical power

Ø  Advantages •  Can withstand harsh environments •  Lifetime of 8 years •  Affordable

Ø  Specifications •  Mass: 6 tons •  Power: 40 kW •  Reactor Volume: 0.15 m3

Ø  Applications •  Power habitats/equipment

–  Located at a safe distance (~250 m from habitat

–  Aluminum shielding to protect from radiation

–  Require at least 2 per habitat •  Charge rovers

–  5-‐10 days

Radioisotope thermoelectric generators (RTG)s

Ø  Specifications •  Mass: 40-‐80 kg •  Power: 250-‐500 W •  Volume: 0.018-‐2 m3

Ø  Applications •  Back-‐up power for habitat

–  Used in conjunction with solar panels

•  Used by rovers for charging at waypoints

126 Divinaa Burder | Power & Thermal 1.  FSP Image based on : NASA Science News hSp://science1.nasa.gov/science-‐news/science-‐at-‐nasa/2009/15may_sErling/ 2.  RTG Image based on : hSp://www.myspacemuseum.com/alsep01b.htm

RE-ENTRY CAPSULE SHIELDING

127

§  Based on SpaceX Dragon Capsule §  Shielding

•  PICA-‐X (Phenolic Impregnated Carbon Ablator) tiles •  Re-‐usable (less than 1 cm burns off during re-‐entry) •  Thickness: 8 cm /tile •  Mass: 1 kg / tile •  Tile Dimensions:25.4 cm x 9.7 cm x 8cm •  Withstands temperatures higher than 1600 °C (3000 ° F) •  Lightweight, “slightly more dense than balsa wood”[2]

Divinaa Burder | Power &Thermal

FISSION REACTOR COMPARED TO ISS Ø  Primary Components

•  Heat Source •  Power Conversion •  Heat Rejection •  Power Conditioning & Distribution

(PCAD) Ø  Structure Description

•  Stainless steel based, UO2 fueled pumped, NaK fission reactor coupled to free piston stirling converters

Ø  Process •  Power transferred from Reactor to Power

Conversion, and then to Heat Rejection. •  Electrical power generated by power

conversion processed through PCAD to user loads

Ø  Radiation •  No radiological risk •  Provide shielding for FSP •  Similar to habitat shielding •  Can control reactivity •  Average gamma dose in local region above

shield = 5MRad

Ø  Primary Components •  Solar panels

•  75-‐90 kW generated •  Wingspan of 240 ft •  Solar panels

•  Each power channel supplies 11 kW •  Nickel-‐hydrogen batteries •  Can route power to user loads •  Primary & Secondary Power Distribution •  25 kW used for science applications

128

REFERENCES

129

[1] Poston, D.I., Kapernick, R.J, "Reference Reactor Module Design for NASA's Lunar Fission Surface Power System," Proceedings of Nuclear and Emerging Technologies for Space 2009, Atlanta, GA, June 14-‐19, 2009. [2] NASA

•  Radioisotope Power systems, http://solarsystem.nasa.gov/rps/types.cfm •  NASA Developing Fission Surface Power Technology

http://www.nasa.gov/home/hqnews/2008/sep/HQ_08-‐227_Moon_Power.html •  NASA/DOE Team Moving Forward on Fission Surface Power Technology

http://www.nasa.gov/centers/glenn/news/pressrel/2009/09-‐036_fission.html [3] SpaceX

•  Pica Heat Shielding, http://www.spacex.com/news/2013/04/04/pica-‐heat-‐shield •  Dragon Capsule: http://www.spacex.com/dragon •  Dragon Capsule infographic : http://www.space.com/12033-‐spacex-‐dragon-‐space-‐capsule-‐infographic.html

[4] Universe today •  Space X Dragon capsule: www.universetoday.com/94217/spacexs-‐dragon-‐now-‐with-‐seating-‐for-‐seven/

•  Fission Reactors:

http://www.universetoday.com/17937/nasa-‐looks-‐at-‐fission-‐reactors-‐for-‐power-‐on-‐the-‐moon/

130

JOSEPH AVELLANO | POW./THERM. CARGO VEHICLE POWER

o  LOW EARTH ORBIT o  SOLAR PANELS o  RECHARGEABLE BATTERIES

1/30/2014

LEO POWER OBSTACLES

131

§  Multi-‐junction solar cells •  26% -‐ 44% eff.

§  Sun’s Radiation •  Max: 1413 𝑊/𝑚↑2   •  Min: 1312 𝑊/𝑚↑2  

§  Low Earth Orbit •  Altitude: 160 km – 2000 km

•  Period: 88 min – 127 min

§  Dark Period •  ~38.88% of total period •  ~34 – 49 min

CARGO VEHICLE POWER NEEDS

132

§  Solar Panels •  At 40% eff. : 546.4 •  3 square meters •  Mass = ~45 kg •  Volume = ~.0762 •  Power = ~1639.2 Watts

§  Lithium Ion Batteries

•  Mass = ~3.2 kg •  Volume = ~.003 •  Power = ~1.5 kW

Navigalon system/sensors

Power required

IMU 22 waSs

Star Tracker 12 waSs

Other Systems Power required

GPS ~12 waSs

CommunicaEons ~100 waSs

Thermal Control ~1000 waSs

§  Total Power = ~1144 Watts

133

ANDREW COX | PROJECT MANAGER THIS, THAT, AND THE OTHER

o  NEW (AND IMPROVED) SPREADSHEET

1/30/2013

SPREADSHEET USAGE

134

§  Thank you to those who have contributed! •  Structures team with initial guestimates •  Comm Sat, CTV, Hab, Cargo Transport

§  Need more numbers from: •  Moon Landers, Const. Bots, Helper Bots, SSC

§  Almost nobody has power numbers – we need those ASAP so P&T can work!

Andrew Cox | PM

IMPROVED VEHICLE SPECS SHEET

135

§  Moved to Google Drive •  Wiggio kept duplicating the file without preserving modifications

•  GD is faster and has more features •  Link is posted on Wiggio where the spreadsheet used to be

§  Explanation and Feature Walkthrough •  Link

Andrew Cox | PM

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