PROJECT MANAGER TREVOR JAHN THURSDAY LAB 1/28/2016

APM/SYSTEMS MIKE YOUNG

1/28/16: Storyboard

STORYBOARD

2018 – XM 1 (LEO) 14 Mg

2018 – XM 2 (CLO) 14 Mg

2023 – Base equip 90 Mg, 4 launches

STORYBOARD

2025 – XM 3 (CLO) 14 Mg

2025 – Base assembly

2025 – First crew 26 Mg

STORYBOARD

2027-2031 Base construction

2025 – 2027 ISRU ramp-up

2035 – Cycler rendezvous

BACK-UP SLIDES STORYBOARDINFO.XLS

BACK-UP SLIDES

Mass and missions per year

1. 2018 – 12 Mg, 1 mission

2. 2020 – 14 Mg, 1 mission

3. 2023 – 59 Mg, 4 missions

4. 2025 – 102.5 Mg, 4 missions

5. 2027 – 66 Mg, 2 missions

6. 2029 – 71 Mg, 2 missions

7. 2031 – 67 Mg, 2 missions

Total missions – 17 + cycler rendezvous

FROM STORYBOARDINFO.XLS

Habitat volume per year

1. 2018 – 0 m3

2. 2020 – 0 m3

3. 2023 – 32 m3

4. 2025 – 64 m3

5. 2027 – 96 m3

6. 2029 – 128 m3

7. 2031 – 144 m3

Volume and mass capacities based on analysis by compiled by Katy O’Connor

(presented 1/21/16) in consultation with group leaders, Trevor Jahn, and James

Millane

SYSTEMS BENJAMIN MISHLER

January 28th, 2016

Back of the envelope Steady-state Mass/Power/Volume calculations

MASS/POWER/VOLUME ESTIMATES CATEGORY BREAKDOWN FOR A 16 PERSON STEADY-STATE SYSTEM

Category Mass [Mg] Power [kW] Volume [m3]

Human Consumables 20.5 0.0 25.0

Expendable Supplies 6.9 0.0 29.6

Vehicles 3.0 0.0 42.0

Communications 0.0 0.5 0.0

Facilities 0.0 45.0 0.0

MASS/POWER/VOLUME ESTIMATES TOTALS PER YEAR ASSUMPTIONS

Mass [Mg] Power [kW] Volume [m3]

34.4 45.5 96.6

• Water is 80% recycled

• No food is being grown

• Rovers use Hydrogen fuel-cells

• Lander propellant cost not

included

MASS/POWER/VOLUME FULL TABLE Categories Mass (Mg) Power (kW) Volume (m^3) Assumptions

Human Consumables

Food 11 N/A 15

Water 9.5 N/A 10 With ~80% water reclamation

Expendable Supplies

Medical 0.5 N/A 1

Clothes N/A N/A 1

Personal / Maintenance 1.4 N/A 7.6

Spare Parts for habs, XMs and rovers 5 N/A 20 ~30% of hab mass and ~1.5Mg for rovers parts could be large

Vehicles

Communication Satellites x3? N/A N/A N/A Self-powered

Construction Rover 1 N/A 14 Just considering Hydrogen for the fuel cells

Exploration Rover x2 2 N/A 28 Just considering Hydrogen for the fuel cells

Science Rover N/A N/A N/A Self-powered

Communications

Moon to Lunar Orbit N/A N/A N/A Negligible power

Moon to LEO/Earth N/A 0.5 N/A

Facilities

ISRU N/A 15 N/A Rough numbers based on an ISRU designed for Mars

Hab life support/water recycling N/A 29.6 N/A Scaled up from numbers for a crew of 6

Space Suits N/A 0.4 N/A

Category Subtotal 0 45 0

Total Mass (Mg) Power (kW) Volume (m^3)

30.4 45.5 96.6

REFERENCES

Project Aldrin-Purdue:

https://engineering.purdue.edu/AAE/Academics/Courses/aae450/2015/spring/docs/Project

Aldrin-PurdueFinalReport.pdf

Fuel-Cell Mine Vehicle - Development and Testing:

https://www1.eere.energy.gov/hydrogenandfuelcells/pdfs/28890bb.pdf

NASA's Lunar Communications & Navigation Architecture:

https://www.nasa.gov/pdf/203072main_LAT2%20C-N%20to%20ESTO%20TEC%202007-

11-15%20rev2.pdf

NASA human exploration of mars design reference architecture 5.0:

https://www.nasa.gov/pdf/373665main_NASA-SP-2009-566.pdf

SYSTEMS KYLE BUSH

Launch Site Selection

1/28/2016

POSSIBLE LAUNCH LOCATIONS

• There were two major factors considered when choosing a site:

• The latitude at which the site is located

• The launch vehicles that each site will support

Space Port Location Latitude

Cape Canaveral Air Force Station United States 28.489° N

Vandenberg Air Force Base United States 34.733° N

Baikonor Cosmodrome Kazakastan (Russia) 45.955° N

Xichang Satellite Launch Center China 28.246° N

Tanegashima Space Center Japan 30.391° N

Guiana Space Center French Guiana (ESA) 5.327° N

Satish Dhawan Space Center India 13.737° N

LATITUDE AND DELTA V

• Based on a desired

orbit inclination of

about 28°

• Difference in Δv

between a latitude of

0° and 90° is 0.4633

km/s.

• Best to launch from

latitude with similar

value as desired angle

• Questions?

BACKUP SLIDE

• Same table as before but with two additional sites, the launch vehicles the site has plans to launch in the near future and the max payload to LEO of the most powerful launch vehicle.

BACKUP SLIDE

Additional Numbers: Δv required for LEO at 0°

latitude: 7.4465 km/s Δv required for LEO at 90°

latitude: 7.9098 km/s Δv difference between the

two: 0.4633 km/s

REFERENCES Chatters, E. P., IV, Eberhardt, B., & Warner, M. S. (2009). Space primer (2nd ed.). Retrieved January 24, 2016, from

http://www.au.af.mil/au/awc/space/au-18-2009/au-18_chap06.pdf

Falcon Heavy. (2016). Retrieved January 23, 2016, from http://www.spacex.com/falcon-heavy

Launch Calendar. (n.d.). Retrieved January 25, 2016, from http://www.spaceflightinsider.com/launch-schedule/

Launch Manifest. (n.d.). Retrieved January 23, 2016, from http://www.spacex.com/missions

Launch Schedule. (n.d.). Retrieved January 25, 2016, from http://spaceflightnow.com/launch-schedule/

Launchers - ISRO. (2016). Retrieved January 26, 2016, from http://www.isro.gov.in/launchers

Long March Launch Record. (2013, December 31). Retrieved January 26, 2016, from

http://www.cgwic.com/LaunchServices/LaunchRecord/LongMarch.html

Rocket Launch Schedule, Astronaut Appearances & Events. (n.d.). Retrieved January 23, 2016, from

https://www.kennedyspacecenter.com/events.aspx

Sforza, P. M. (2015). Manned spacecraft design principles. Waltham, MA: Butterworth-Heinemann.

Webb, B. (2016). Vandenberg AFB Launch Schedule. Retrieved January 23, 2016, from

http://www.spacearchive.info/vafbsked.htm

Williams, D. R. (2015, December 22). Earth Fact Sheet. Retrieved January 24, 2016, from

http://nssdc.gsfc.nasa.gov/planetary/factsheet/earthfact.html

Williams, D. R. (2015, December 22). Moon Fact Sheet. Retrieved January 24, 2016, from

http://nssdc.gsfc.nasa.gov/planetary/factsheet/moonfact.html

Writers, S. (2015, November 30). The Ins and Outs of NASA's First Launch of SLS and Orion. Retrieved January

23, 2016, from

http://www.spacedaily.com/reports/The_Ins_and_Outs_of_NASAs_First_Launch_of_SLS_and_Orion_999.html

SYSTEMS ANAIS ARNAIZ

Risk Assessment, Requirements Documentation and Tracking

1/28/2016

RISK ASSESSMENT MATRIX

1 2 3

4

5 6 7

8

9 10 12 11

Authors: Anais Arnaiz

RISK ASSESSMENT MATRIX

Example: Moonquakes have a 0.002% chance per year. Need more Research to support this. A quake could potentially damage equipment. So…. Approach is R – Research Likelihood is 1 Consequence is 2

□ New

↑ Increased

↓ Decreased

→ Same

Risk Trend

Rank Trend Risk Title Approach Likelihood Consequence

1 □ Cyclical Solar Flare Activity M 100% Astronauts exposed to radiation

2 □ Coronal Mass Ejection M 100% Exposing Astronauts to dangerous energetic particles

3 □ Magnetotail R 100% Could affect equipment. Sudden voltage peaks

4 □ Launch Failure A 8.90% Loss of Crew and payload

5 □ Missing the Rendezvous A 0.30% Loss of Crew and payload

6 □ Loss of Navigation to Transfer Vehicle A 0.01% Loss of Crew and payload

7 □ Crashing into lunar surface A 0.00% Loss of Crew and payload

8 □ Communication failure R Maybe Loss of communication

9 □ Meteorite Impact A 0.0263% chance of impact Could cause damage to our habs and equiment 

10 □ No Water in Cabeus Crater A 20% of Cabeus is  5wt%  Need a different way of getting water and fuel

11 □ Moonquakes R 0.002% per year Could cause damage to our habs and equiment

12 □ Landing in a different location A 0.001% Not in designated location

Authors: Anais Arnaiz, Nicholas Hobar

REFERENCES

• E E Bassett, J M Bremner, ”Statistics: Problems and Solutions”. Pp.19

• Robert H. Lewis, “Human Safety in the Lunar Environment” http://www.nss.org/settlement/nasa/spaceresvol4/human.html.

• Gordon Holman, “Space Weather: What impact do solar flares have on human activities?” http://hesperia.gsfc.nasa.gov/sftheory/spaceweather.htm

• I-Shih Chang, “Space Launch Vehicle Reliability”. http://www.ewp.rpi.edu/hartford/users/papers/engr/ernesto/cedenc/SMRE/Project/Space%20Shuttle%20Vehicle%20Reliability.pdf pp.3

Calculations:

• Roughly 1000 observed meteorite impacts in the last decade on the lunar surface, all ranging from the 10s of grams to several kilograms. Lunar Surface Area: 3.8E7 km^2 / 1000 impacts = 38000.

• 1/38000 chance of impact

• Surface Area / (Average quakes per year * 1000 km^2 area of effect)

BACKGROUND

Risk Title Approach Likelihood Consequence

Cyclical Solar Flare ActivityM

100% More flares during 11-year solar cycle Astronauts exposed to radiation

Coronal Mass EjectionM

100% Approx. 90 sunspots at peak of activity in 2023. Exposing Astronauts to dangerous amounts of energetic particles

MagnetotailR

100% for 6 days every lunnar cycle Need more reserach. Lunar particle could affect equipment. Sudden voltage peaks

Launch FailureA

8.9% launch failure, Of the 4378 space launches conducted worldwide between 1957

and 1999, 390 launches failed (the success rate was 91.1 percent) Loss of Crew and payload

Missing the Rendezvous A 0.3% According to refernce, design for 3-sigma reliability (99.7%) Loss of Crew and payload

Loss of Navigation to Transfer Vehicle A 0.01% 3 components in each system to have less than 1% chance of failure Loss of Crew and payload

Crashing into lunar surface A 0.001% chance of not succefulling landing on the surface Loss of Crew and payload

Communication failure R The magnetotail could potentially cause issues here Loss of communication between moon and surface and between vehicles

Meteorite Impact A 0.0263% chance of impact Could cause damage to our habs and equiment 

No Water in Cabeus Crater A Conservative estimates say 20% of Cabeus floor is 5wt% "wet" Need a different way of getting water and fuel

Moonquakes R 0.002% chance per year. Roughly 6 shallow moonquakes per year. Could cause damage to our habs and equiment

Landing in a different location A 0.001% chance of not succefulling landing on the surface Need to move everything to designated location or work with landing site

BACKGROUND

http://www.aeronautics.nasa.gov/images/con

tent/sae_graphic4_lg2.jpg

SCIENCE ELLEN CZAPLINSKI

January 28, 2016

Cabeus Crater Details

Resource Abundances

25

CABEUS CRATER DETAILS PHYSICAL DETAILS OF CABEUS

Average depth of crater walls (LOLA)

4 km

Slope of crater walls (JMARS)

10˚ - 15˚

Average temperature of PSR (Lunar sourcebook)

-236.11˚ C

Slope angles of Cabeus Crater. Created in JMARS by Ellen Czaplinski.

• Negligible Sulfur • Most sulfur is found in the

mineral Troilite (FeS) • Troilite forms < 1% by

volume of lunar rock

0˚ 40˚

26

RESOURCE ABUNDANCES

Compound Fraction of Regolith

Mass of Compound

(kg)

H2O 0.03 1.14 x 1010

H2S 4.52 x 10-3 1.92 x 109

NH3 1.63 x 10-3 6.90 x 108

SO2 8.61 x 10-4 3.65 x 108

C2H4 8.42 x 10-4 3.57 x 108

CO2 5.86 x 10-4 2.48 x 108

CH3OH 4.19 x 10-4 1.77 x 108

CH4 1.76 x 10-4 7.44 x 107

OH 8.1 x 10-6 3.43 x 106

BASED OFF RESULTS FROM LCROSS IMPACT INTO CABEUS PSR

Modified from Colaprete et al. (2010). All values are from the top 10 m of regolith only.

Created in JMARS using the LROC color shaded relief and normalized reflectance layers. Made by Ellen Czaplinski.

27

REFERENCES Colarprete, A. et al. (2010). Detection of Water in the LCROSS Ejecta Plume. Science, 330, 463-468. Heiken, G. H., Vaniman, D. T., and French, B. M. (1991). Lunar Sourcebook: A user’s guide to the moon. Cambridge University Press, New York. Kozlova, E. A., Lazarev, E. N. (2010). Crater Cabeus as possible cold trap for volatiles near south pole of the moon. 41st LPSC. 1779. Kring, D. A., et al. (2012) A Global Lunar Landing Site Study to Provide the Scientific Context for Exploration of the Moon. LPE-JSC Center for Lunar Science and Exploration.

28

BACKUP SLIDES

Calculations of the resource abundances for Cabeus

Slope angles of Shackleton (left) and Amundsen (right). Created in JMARS by Ellen Czaplinski. These slopes are much steeper than Cabeus’s 10˚ - 15˚.

Compound fraction of regolith Mass of Compound

H2O 0.027 1.14E+10

H2S 0.0045225 1.92E+09

NH3 0.0016281 6.90E+08

SO2 0.0008613 3.65E+08

C2H4 0.0008424 3.57E+08

CO2 0.0005859 2.48E+08

CH3OH 0.0004185 1.77E+08

CH4 0.0001755 7.44E+07

OH 0.0000081 3.43E+06

29

BACKUP SLIDES

LROC shaded relief map created in JMARS by Ellen Czaplinski

30

SCIENCE JAKE ELLIOTT

Sites of Interest

Probe Landing Sites

31

SITES OF INTEREST JAKE ELLIOTT

Site Long (°)

Lat (°)

Elevation (m)

Illumination (%)

Distance to ISRU (km)

ISRU 314.20 E 84.67 S -3570 <5 -

Sample Return

308.64 E 85.50 S -4980 21-25 30

Mountain Peak

323.44 E 83.60 S 5100 >50 42

Landing Sites • Dependent on amount of sunlight & area needed • Range of rovers • Slope <5°

Sample Return • Interior of crater within Cabeus • ~1 kg of rock • Date to within ±20 mya • Analyze minerals to within ppb

Mountain Peak • ~80% illumination • Near constant view of Earth • Possible back-up comm. site

32

PROBE LANDING SITES JAKE ELLIOTT

Landing Sites

Site # Long (°) Lat (°)

1 313.03 E 30.71 N

2 315.26 E 1.33 N

3 312.50 E 30.44 S

4 309.66 E 61.53 S

Criteria • Along same longitude • ~30° of latitude spacing • Low regolith content

Purpose • Determine size/structure of crust, mantle, and core • Determine distribution of lunar seismic activity • More TBD

90 S

0

33

REFRENCES JAKE ELLIOTT

Freed, A.M., Interview, 1/26/2016 Kozlova, E. A., and E. N. Lazarev (2010), Crater Cabeus as Possible Cold Trap for Volatiles Near South Pole of the Moon, , 2–3. McKay, D. S., G. Heiken, A. Basu, G. Blanford, S. Simon, R. Reedy, B. M. French, and J. Papike (1991), Lunar Sourcebook, Vision and Voyages for Planetary Science in the Decade 2013-2022 Committee,

34

CABEUS JAKE ELLIOTT

Kozlova (2010)

Elevation, rings mark 5 km spacing 35

CABEUS JAKE ELLIOTT

Slope, rings mark 5 km spacing

36

PROBE LANDING SITES JAKE ELLIOTT

TEMPERATURE VARIATION FROM GLOBAL AVERAGE

Site 1 Site 2

Site 3 Site 4

37

MISSION DESIGN PAUL WITSBERGER

Hyperbolic Rendezvous with Cycler Vehicle in S1L1

Trajectory

1/27/2016

HYPERBOLIC RENDEZVOUS THE PROBLEM

Started with assumptions made by

Project Aldrin-Purdue1

• Mars and Earth are in circular,

coplanar orbits

• Neglect Mars gravity

• One synodic period is exactly 2

1/7 years

Figure 1: Inner and Outer sections of S1L1 Trajectory

39

HYPERBOLIC RENDEZVOUS PRELIMINARY ANALYSIS

Cycler delta V wrt Earth @ inf:

• 4.77 km/s

Cycler delta V wrt Earth @ periapsis:

• 6.61 km/s

• periapsis altitude: 31818 km

• flight path angle: 38.69 deg

Required delta V from circular orbit at 31818 km:

• 4.56 km/s

• 62.5 Mg total mass assuming Isp of 400 sec

• Payload mass fraction of 31.8% 40

Cycler:

6.61 km/s

Delta V:

4.56 km/s

Rendezvous

Vehicle: 3.23 km/s

REFERENCES

[1] Machuca, Pablo. “Circular-Coplanar Model of the S1L1 Cycler Trajectory.” Project

Aldrin-Purdue (2015): 568-578.

41

BACKUP SLIDES MATLAB CODE FOR PLOT ON PAGE 2

42

% Paul Witsberger

% AAE450

% 01/27/2016

% S1L1 Cycler Orbit

sun = Planet(1.327e11,695500,0);

earth = Orbit(sun, 1.496e8, 0, 0, 0, 0);

mars = Orbit(sun, 1.5206*1.496e8, 0, 0, 0, 0);

inner = Orbit(sun, 1.0483*1.496e8, 0.1609, 0, 0, 20.4);

outer = Orbit(sun, 1.3039*1.496e8, 0.2554, 0, 0, -31.1);

figHand = figure(1);

set(figHand,'Position',[400 0 1000 1000])

h1 = earth.plot3d;

hold on

h2 = mars.plot3d;

h3 = inner.plot3d;

h4 = outer.plot3d;

grid on

title('S1L1 Orbit - Witsberger')

xlabel('x (km)')

ylabel('y (km)')

axis equal

xlim([-4e8 4e8])

ylim([-4e8 4e8])

h5 = scatter(0,0,[],'yellow','filled');

view([0,0,30])

legend('Earth','Mars','Inner Orbit','Outer Orbit')

set(gca,'FontSize',15)

print(figHand, '-djpeg', '-r300', 'CyclerOrbit');

BACKUP SLIDES MATLAB CODE FOR PLOT ON PAGE 2

43

Cabeus Orbit:

1 year

Cabeus Orbit:

5 years

MISSION DESIGN MICHELLE MADALINSKI

Hyperbolic Lambert Arcs

Propellant Mass Ratio

Fuel Percentage

STK Visualization

TRANSFERS & MISSION SPECIFICATIONS 1H LAMBERT ARC

Delta V: 4.791 km/s

Propellant Mass Fraction: 2.953

Payload Percentage: 33.858%

Delta V [km/s]

Earth to LEO (inclination correction burn)

9

LEO to CLO 4.791

CLO to Surface 2.656

Total 16.447

Table 1: Total Delta V Breakdown

References: Jay Millane, Josh Ostman, Propulsion Team

LEO TO CLO

STK VISUALIZATION

MOON FRAME

Green: Hyperbolic Transfer

Blue: LAO

Magenta: Hohmann Transfer

Cyan: 5,000 km radius CLO

Yellow: Hohmann Transfer

REFERENCES

• (2012, May 1). Retrieved January 25, 2016, from

http://www.nasa.gov/mission_pages/station/expeditions/expedition30/tryanny.html

o Compared numbers for sanity check

• Propulsion Team, Jay Millane, Josh Ostman, Mike Young

BACKUP SLIDES 2 H LAMBERT ARC TRANSFER

Here is the analysis for the 2H Lambert Arc Transfer to the Moon’s

Orbit. This proved to be more expensive, in terms of delta V, than the

1H Lambert Arc Transfer. Clearly these hyperbolic options are both way

more expensive than the currently chosen Hohmann Transfer, but we

wanted to consider all options. This can at least be used as an

emergency plan to get to the moon in about 2 days if that were needed.

Below is a table with the minimum from both transfer types.

Type 1H Type 2H

Total Delta V [km/s]

5.0019 5.2767

Propellant Mass Fraction

3.0974 3.2958

Fuel Percentage [%]

67.7144 69.6586

Table 2: Hyperbolic Lambert Arcs

BACKUP SLIDES MATLAB CODE (LAMBERT ARCS)

% Michelle Madalinski

% Ref. Jay Millane

% AAE 450

% Investigating Hyperbolic Transfers

%% 2H varying TA

r1 = 6771; % km

r2 = 389403; % km

%ta = 250; % deg

tof = 2*24*3600; % sec

mu = 398600.441800; % km^3/s^2

mu_moon = 4904.8695; % km^3/s^2

Isp = 451; % LOX LH2

g0 = 9.81/1000; % km/s^2

for ta = [180:1:360]

[type, a, p, e, energy, theta1, theta2, v1, v2, y1, y2] = lambertArc2(r1, r2, ta, tof, mu); % Jay’s Code

V_D = v1;

V_A = v2;

gamma_D = y1;

gamma_A = y2;

V_earth = sqrt(mu/r1); % km/s

V_moon = sqrt(mu/r2); % km/s

deltaV1 = sqrt(V_earth^2 + V_D^2 - 2*V_earth*V_D*cosd(gamma_D)); % km/s

deltaV2 = sqrt(V_moon^2 + V_A^2 - 2*V_moon*V_A*cosd(gamma_A)); % km/s

deltaV_total = abs(deltaV1) + abs(deltaV2); % km/s

%fprintf('deltaV_total = %f km/s\n', deltaV_total)

MR = exp(deltaV_total/(Isp*g0));

hold on

figure(1)

subplot(2,1,1)

h = plot(ta, deltaV_total, 'o');

set(h, 'MarkerEdgeColor','r','MarkerFaceColor','r')

xlabel('Transfer Angle [deg]')

ylabel('\DeltaV [km/s]')

title('\DeltaV vs TA')

set(gca,'fontsize',16)

grid on;

hold on

subplot(2,1,2)

g = plot(deltaV_total, MR, 'o');

set(g, 'MarkerEdgeColor','b','MarkerFaceColor','b')

xlabel('\DeltaV [km/s]')

ylabel('Propellant Mass Fraction')

title('Propellant Mass Fraction vs \DeltaV')

set(gca,'fontsize',16)

grid on;

end

%suptitle('2H Lambert Transfer Arc')

%set(gca,'fontsize',16)

%% H varying TA r1 = 6771; % km r2 = 389403; % km %ta = 250; % deg tof = 2*24*3600; % sec mu = 398600.441800; % km^3/s^2 mu_moon = 4904.8695; % km^3/s^2 Isp = 451; % LOX LH2 g0 = 9.81/1000; % km/s^2 for ta = [0:1:180] [type, a, p, e, energy, theta1, theta2, v1, v2, y1, y2] = lambertArc2(r1, r2, ta, tof, mu); % Jay’s Code V_D = v1; V_A = v2; gamma_D = y1; gamma_A = y2; V_earth = sqrt(mu/r1); % km/s V_moon = sqrt(mu/r2); % km/s deltaV1 = sqrt(V_earth^2 + V_D^2 - 2*V_earth*V_D*cosd(gamma_D)); % km/s deltaV2 = sqrt(V_moon^2 + V_A^2 - 2*V_moon*V_A*cosd(gamma_A)); % km/s deltaV_total = abs(deltaV1) + abs(deltaV2); % km/s %fprintf('deltaV_total = %f km/s\n', deltaV_total) MR = exp(deltaV_total/(Isp*g0)); hold on figure(2) subplot(2,1,1) h = plot(ta, deltaV_total,'o'); set(h, 'MarkerEdgeColor','r','MarkerFaceColor','r') xlabel('Transfer Angle [deg]') ylabel('\DeltaV [km/s]') title('\DeltaV vs TA') set(gca,'fontsize',16) grid on; hold on subplot(2,1,2) g = plot(deltaV_total, MR, 'o'); set(g, 'MarkerEdgeColor','b','MarkerFaceColor','b') xlabel('\DeltaV [km/s]') ylabel('Propellant Mass Fraction') title('Propellant Mass Fraction vs \DeltaV') set(gca,'fontsize',16) grid on; end %suptitle('1H Lambert Transfer Arc') %set(gca,'fontsize',16)

%% Minimum clc; r1 = 6771; % km r2 = 389403; % km ta = 163; % deg tof = 2*24*3600; % sec mu = 398600.441800; % km^3/s^2 mu_moon = 4904.8695; % km^3/s^2 [type, a, p, e, energy, theta1, theta2, v1, v2, y1, y2] = lambertArc2(r1, r2, ta, tof, mu); % Jay’s Code V_D = v1; V_A = v2; gamma_D = y1; gamma_A = y2; V_earth = sqrt(mu/r1); % km/s V_moon = sqrt(mu/r2); % km/s deltaV1 = sqrt(V_earth^2 + V_D^2 - 2*V_earth*V_D*cosd(gamma_D)); % km/s deltaV2 = sqrt(V_moon^2 + V_A^2 - 2*V_moon*V_A*cosd(gamma_A)); % km/s deltaV_total = abs(deltaV1) + abs(deltaV2); % km/s fprintf('deltaV_total = %f km/s\n', deltaV_total) Isp = 451; % LOX LH2 g0 = 9.81/1000; % km/s^2 MR = exp(deltaV_total/(Isp*g0)); fprintf('Propellant Mass Fraction = %f\n', MR) FP = 100 - (MR^-1 * 100); fprintf('Fuel Percentage = %f\n', FP)

BACKUP SLIDES MATLAB CODE (MASS & FUEL PERCENTAGE) % Michelle Madalinski

% Ref. Jay Millane, Josh Ostman, Alex Burton, Propulsion Team

% AAE 450

% Mass Fraction and Mass & Fuel Percentage

%%

clc;

Isp = 451; % LOX LH2

g0 = 9.81/1000; % km/s^2

deltaV_Earth2LEO = 9; % km/s

deltaV_LEO2CLO = 4.791483; %km/s

deltaV_CLO2surface = 2.656; %km/s

deltaV_total = deltaV_Earth2LEO + deltaV_LEO2CLO + deltaV_CLO2surface; %km/s

fprintf('Total Delta V = %f km/s\n', deltaV_total)

MR_total = exp(deltaV_total/(Isp*g0));

fprintf('Total Propellant Mass Fraction = %f\n', MR_total)

PP_total = MR_total^-1 * 100;

fprintf('Total Payload Percentage = %f\n', PP_total)

FP_total = 100 - (MR_total^-1 * 100);

fprintf('Total Fuel Percentage = %f\n\n', FP_total)

fprintf('Earth to LEO Delta V = %f km/s\n', deltaV_Earth2LEO)

MR_Earth2LEO = exp(deltaV_Earth2LEO/(Isp*g0));

fprintf('Earth to LEO Propellant Mass Fraction = %f\n', MR_Earth2LEO)

PP_Earth2LEO = MR_Earth2LEO^-1 * 100;

fprintf('Earth to LEO Payload Percentage = %f\n', PP_Earth2LEO)

FP_Earth2LEO = 100 - (MR_Earth2LEO^-1 * 100);

fprintf('Earth to LEO Fuel Percentage = %f\n\n', FP_Earth2LEO)

fprintf('LEO to CLO Delta V = %f km/s\n', deltaV_LEO2CLO)

MR_LEO2CLO = exp(deltaV_LEO2CLO/(Isp*g0));

fprintf('LEO to CLO Propellant Mass Fraction = %f\n', MR_LEO2CLO)

PP_LEO2CLO = MR_LEO2CLO^-1 * 100;

fprintf('LEO to CLO Payload Percentage = %f\n', PP_LEO2CLO)

FP_LEO2CLO = 100 - (MR_LEO2CLO^-1 * 100);

fprintf('LEO to CLO Fuel Percentage = %f\n\n', FP_LEO2CLO)

fprintf('CLO to Surface Delta V = %f km/s\n', deltaV_CLO2surface)

MR_CLO2surface = exp(deltaV_CLO2surface/(Isp*g0));

fprintf('CLO to Surface Propellant Mass Fraction = %f\n', MR_CLO2surface)

PP_CLO2surface = MR_CLO2surface^-1 * 100;

fprintf('CLO to Surface Payload Percentage = %f\n', PP_CLO2surface)

FP_CLO2surface = 100 - (MR_CLO2surface^-1 * 100);

fprintf('CLO to Surface Fuel Percentage = %f\n\n', FP_CLO2surface)

• Using the delta V from Earth to LEO that the Propulsion Team calculated, I was able to obtain a total delta V for a mission to the Moon. The total delta V was then used to output a propellant mass fraction, payload percentage, and a fuel percentage that can be passed along to Mike and the Systems Team.

Delta V [km/s]

Propellant Mass

Fraction

Payload Percentage

Fuel Percentage

Earth to LEO 9 7.646 13.078 86.922

LEO to CLO 4.791 2.953 33.858 66.142

CLO to Surface

2.656 1.823 54.864 45.136

Total 16.447 41.162 2.429 97.571

Table 3: Total Mission Breakdown

• This is not the final decision, we are looking at the payload percentage for each orbit and what we can and cannot take to LEO, and how we will then get that to CLO. We are looking to follow the Storyboard Mike created and determine how much fuel we will need to get the payloads where they need to be.

BACKUP SLIDES STK VISUALIZATION FROM EARTH TO MOON ORBIT (EARTH FRAME)

From Earth to LEO, we believe we can insert the spacecraft in the right inclined orbit to be aligned with the Moon. Here is the Hohmann transfer that is currently being used from an orbit around the Earth to the Moon’s Orbit. This is in the Earth Frame. The previous STK model was considered from the Moon frame. After the Hohmann Transfer gets the spacecraft to the Moon’s vicinity, the delta V will be applied to send the spacecraft into a hyperbolic translunar injection. In this transfer orbit, the spacecraft will enter the sphere of influence parallel to the Moon’s velocity in a polar orbit. This allows free delta V to be obtained. From there is will enter the LAO (Lunar Arrival Orbit). This is just a general orbit around the Moon. A Hohmann Transfer will be used to get the spacecraft into the chosen 5,000 km CLO (Circular Lunar Orbit). From there, another delta V will be applied to land on the surface. Using STK, I was able to prove that the generated delta V’s and orbits are plausible. As discussed previously, this was an orbit that was found to be stable around the Moon. We are currently looking into lower circular orbits that are equally or more stable. This would allow for increased performance in communications and a smaller required delta V to descend to the surface.

Earth’s Orbit

Transfer Arc

Moon’s Orbit

MISSION DESIGN MASON BUCKMAN

1B & 2B Elliptical Lambert Arcs

Rocket Mass Ratios

Useful ΔV Estimates

01/28/2016

LAMBERT ARCS & MASS RATIOS DETERMINING AN EFFICIENT TRAJECTORY

• LEO to Lunar orbit: ΔV ≈ 4.20 km/s • Varied TOF and transfer angle • Hohmann Transfer best option for missions

NEXT STEPS GENERALIZING ΔV ESTIMATES

Excerpt from the Storyboard Information Spreadsheet by Mike Young

• Comprehensive Earth to Moon Trajectory code • Credits to Jay Millane for code

• Systems and Propulsion need realistic numbers • ΔV = 4.791 km/s for LEO – Circular Lunar Orbit

• ΔV = 7.447 km/s for LEO – Lunar Surface

BACKUP SLIDES 2B LAMBERT ARCS

The ΔV requirements for 1B and 2B lambert arcs

were calculated using a function file created by

Jay Millane. The analysis showed that, not

surprisingly, the most efficient possible path had a

transfer angle of 180 degrees, and a time of flight

close to the minimum energy time of flight. This

corresponds to a general Hohmann transfer. The

ΔV required was calculated from LEO to a

5000km Lunar orbit. The Earth and Moon were

assumed to be coplanar with circular orbits for this

initial estimate. The table below summarizes the

most efficient results obtained from the ΔV

analysis.

Most Efficient ΔV 4.20 km/s

Most Efficient TOF 5.10 Days

Most Efficient Mass Ratio

2.80

BACKUP SLIDES MORE ON ΔV ESTIMATES FOR SPECIFIC MISSIONS

Mission Destination Tentative # Missions

Approximate ΔV [km/s]

Power Equipment Lunar Surface 1 7.447

Communication Satellites

Polar Lunar Orbit 1 4.971

Rover Missions Lunar Surface 2 14.894

Habitat Missions Lunar Surface 8+ 59.580

Using the ΔV’s obtained in our code, we were able to begin to piece together a general total ΔV requirement for the first portion of the overall mission. These numbers are valid for any mission involving going to a polar lunar orbit, or landing on the lunar surface. This past week, the base location was chosen to be in Cabeus Crater which is situated near the south pole on the Moon. Any lunar surface mission is assumed to be landing near the south pole which is what was taken into consideration while obtaining the ΔV numbers in the code. Rough Estimate for Early Mission ΔV Requirements: 86.890 km/s The systems and propulsion teams mentioned that they were looking for rough ΔV estimates so that they could piece together mass capabilities of each mission.

BACKUP SLIDES MATLAB CODE USED FOR OBTAINING PLOTS %Mason Buckman %1&2B Lambert Arcs with reference to function file created by Jay Millane clear; clc; Isp = 451; %Isp for LOX LH2 in a vacuum mu_E = 398600.4418; %Earth's gravitational parameter mu_M = 4904.8695; %Moon's gravitational parameter r2 = 389403; %Arrival in lunar vicinity r1 = 6771; %Low Earth orbit tof = 5.1*24*3600; %Time of flight v_E = sqrt(mu_E/r1); %Velocity in LEO v_M = sqrt(mu_E/r2); %Velocity at lunar vicinity arrival g = 9.81/1000; %Gravitational acceleration of the Earth % --- Type 1B varying transfer angle --- % %Plotting Mass Ratio and dv vs. TA for (tof = (5.1:1:10.1)*24*3600); for (ta = 1:5:175) [type, a, p, e, energy, theta1, theta2, v1, v2, y1, y2] = lambertArc2(r1, r2, ta, tof, mu_E); dv1 = sqrt(v_E^2 + v1^2 - 2*v_E*v1*cosd(y1)); dv2 = sqrt(v_M^2 + v2^2 - 2*v_M*v2*cosd(y2)); dv = abs(dv1) + abs(dv2); figure(1); hold on; plot(ta,dv,'go'); grid on; xlabel('Transfer Angle [Deg.]'); ylabel('\DeltaV [km/s]'); title('Type 1B \DeltaV vs. Transfer Angle'); figure(2); MR = exp(dv/(g*Isp)); hold on; plot(dv,MR,'bo'); grid on; xlabel('\DeltaV [km/s]'); ylabel('Mass Ratio'); title('Comparison of \DeltaV and Mass Ratio'); end end

%Plotting 2B dv vs. TA for ta = 181:5:355 [type, a, p, e, energy, theta1, theta2, v1, v2, y1, y2] = lambertArc2(r1, r2, ta, tof, mu_E); dv1 = sqrt(v_E^2 + v1^2 - 2*v_E*v1*cosd(y1)); dv2 = sqrt(v_M^2 + v2^2 - 2*v_M*v2*cosd(y2)); dv = abs(dv1) + abs(dv2); figure(3); hold on; plot(ta,dv,'ro'); grid on; xlabel('Transfer Angle [Deg.]'); ylabel('Delta V [km/s]'); title('Type 2B \DeltaV vs. Transfer Angle'); legend('TOF of 5.1 Days') end

This code uses a function that was written to calculate lambert arcs. The initial radius, final radius, transfer angle, time of flight, and gravitational parameter of the central body were placed into the function. From there, ΔV was calculated using vector relationships. The mass ratios were obtained using the rocket equation.

MISSION DESIGN ALEXANDER BURTON

Inclination Changes Around the Moon

1/28/2016

DEFINING THE PROBLEM SUMMARY

1. Proper inclination on arrival at the Moon can be

achieved cheaply by controlling exactly how we

approach it.

2. Need estimated of ΔV cost of plane changes in Lunar

orbit for sizing purposes.

Analysis Overview

Objective Minimize ΔV for a 15o plane change around the Moon.

Variables Intermediate elliptical orbits. Circular orbit radius.

RESULTS

Orbital Radius ΔV Cost

1740 km (Lunar Surface)

0.438 km/s

5000 km 0.258 km/s

• Considered ellipses with

eccentricities up to 0.9.

• A transfer ellipse makes

plane changes more

expensive up until about

45o

• All plane changes

should be made in the

original orbit to minimize

ΔV costs.

BACKUP SLIDE CLEAN 15O PLOT

BACKUP SLIDE 45O ΔV PLOT

BACKUP SLIDE 45O TOF PLOT

BACKUP SLIDE 90O ΔV PLOT

BACKUP SLIDE 90O TOF PLOT

PROPULSION DAYLE ALEXANDER

• Cryogenic Boiloff and Electrolysis Information

• Potential Fuel Depot Concept Design Ideas and Tank Sizing

MASS AND SIZING CALCULATIONS

DAYS REQUIRED TO MAKE H2 AND O2

FLUID DAYS REQUIRED (DAYS)

LH2 60

LOX 41

VOLUME REQUIRED FOR HOLDING TANKS

FLUID VOLUME (m^3)

LH2 817

LOX 280

MASS REQUIRED FOR 1 LANDER LAUNCH TO LUNAR ORBIT FROM THE MOON

FLUID MASS (Mg)

LH2 58

LOX 319

H2O 1.04

• Calculations for 1 launch from the lunar surface into lunar orbit with the 20Mg lander (full of cargo) and RL10B-2 engine

• Assuming 1000kg of LH2

can be made in 1 day by electrolysis of H2O

• Cryogenic boil-off rate of

3%/month

ISRU/FUEL DEPOT CONCEPT • Concept for combined ISRU and fuel depot for LOX and fuel • Requires supply of liquid water for electrolysis • Requires condensers to put fuel and oxidizer into liquid forms • Holding tanks require a surrounding water tank to protect from radiation

and an insulating layer underneath

CONDENSERS

FUEL HOSE TO LANDER

LOX HOSE TO LANDER

BATTERY

O2 FILL LINE

RELIEF VALVES

H2 FILL LINE

LIQUID WATER TANK

CATHODE ANODE

WATER TANK

INSULATION LAYER

REFERENCES • “RL10 Engine”, Aerojet Rocketdyne Launch Vehicle,

http://www.rocket.com/rl10-engine [retrieved 26 January 2016]

• “Cryogenic Fluid Management”, NASA Ames Technology Capabilities and Facilities, http://www.nasa.gov/centers/ames/research/technology-onepagers/cryogenic-fluid-management.html [retrieved 26 January 2016]

• “Hydrogen Production: Electrolysis”, energy.gov Office of Energy Efficiency & Renewable Energy, http://energy.gov/eere/fuelcells/hydrogen-production-electrolysis [retrieved 27 January 2016]

• “Liquid Oxygen and Liquid Hydrogen Storage”, NASA Space Shuttle, http://www.nasa.gov/mission_pages/shuttle/launch/LOX-LH2-storage.html, [retrieved 27 January 2016]

BACKUP SLIDE 1 MATLAB CODE FOR TANK SIZING INFORMATION

% -MODEL FOR FUEL DEPOT TANK SIZING-

% AUTHOR: DAYLE ALEXANDER

% LAST UPDATED: 1/27/2016

% ASSUMPTIONS: -FUEL AND OXIDIZER TANKS WILL BE FILLED IN THE SAME DAY AS

% THE LAUNCH AND TANKS ARE PRE CHILLED (NO MASS LOSS)

% -CRYOGENIC BURNOFF RATE IN LEO IS THE SAME AS THE SURFACE OF

% THE MOON

% -IT TAKES 1 DAY TO MAKE 1000KG OF H2

% -100% EFFICIENCY

% KNOWN VARIABLES

RL10_mlh2=58; % Mass of H2 needed to launch 10Mg lander in Mg

RL10_of=5.5; % O/F ratio for the RL10 engine

RL10_mlo2=RL10_mlh2*RL10_of; % Mass of O2 needed to launch in Mg

boiloff=0.1; % Boiloff rate for cryogens in space in %/day

rho_lox=1141; % Density of LOX in kg/m^3

rho_lh2=71; % Density of LH2 in kg/m^3

rho_gox=1.35; % Density of GOX in kg/m^3

rho_gh2=0.085; % Density of GH2 in kg/m^3

mm_o2=0.016; % Molar mass of oxygen in kg/mol

mm_h2=0.001; % Molar mass of hydrogen in kg/mol

mm_h2o=0.018; % Molar mass of H2O in kg/mol

m_h2_daily=1000; % Mass of GH2 able to be made in 1 day

m_o2_daily=m_h2_daily/mm_h2/2*mm_o2; % Mass of GO2 able to be made in 1 day

% EQUATIONS

v_lox=(RL10_mlo2*1000)/rho_lox % Min volume of LOX needed in m^3

v_lh2=(RL10_mlh2*1000)/rho_lh2 % Min volume of LH2 needed in m^3

v_gox=(RL10_mlo2*1000)/rho_gox; % Min volume of GOX needed in m^3

v_gh2=(RL10_mlh2*1000)/rho_gh2; % Min volume of GH2 needed in m^3

mol_o2=RL10_mlo2/mm_o2; % Moles of O2 required

mol_h2=RL10_mlh2/mm_h2; % Moles of H2 required

mol_water_o2=mol_o2*2; % Moles of H2o to get required O2

mol_water_h2=mol_h2; % Moles of H2O to get required H2

m_water_o2=(mol_water_o2*mm_h2o)/1000; % Mass of H2O to get required O2 in Mg

m_water_h2=(mol_water_h2*mm_h2o)/1000 % Mass of H2O to get required H2 in Mg

% MODEL FOR DAYS TO MAKE REQUIRED H2 days=linspace(1,1000,1000); mlh2=[]; mlh2(1)=m_h2_daily; i=1; while(mlh2(i)<RL10_mlh2*1000) i=i+1; mlh2(i)=mlh2(i-1)+m_h2_daily-mlh2(i-1)*(boiloff/100); end days_h2=i % MODEL FOR DAYS TO MAKE REQUIRED O2 mlo2=[]; mlo2(1)=m_o2_daily; i=1; while(mlo2(i)<RL10_mlo2*1000) i=i+1; mlo2(i)=mlo2(i-1)+m_o2_daily-mlo2(i-1)*(boiloff/100); end days_o2=i % RESULTS % VOLUME REQUIRED OF LOX TANK % v_lox = % 279.5793 % % VOLUME REQUIRED OF LH2 TANK % v_lh2 = % 816.9014 % % MASS OF WATER REQUIRED % m_water_h2 = % 1.0440 % % DAYS IT TAKES TO MAKE REQUIRED LH2 % days_h2 = % 60 % % DAYS IT TAKES TO MAKE REQUIRED LOX % days_o2 = % 41

BACKUP SLIDE 2

PROPULSION BROCK MILLER

10Mg and 20Mg Cargo Lander Propulsive Specifications

CARGO LANDERS WITHOUT TUG

-SLS Launches the cargo lander to LEO

-Cargo lander goes from LEO to Lunar Surface following MD trajectory

-Trajectory does not account for inclination change

Engine Initial Mass (Mg) Final Mass (Mg) Inert Mass (Mg) Prop. Mass (Mg) PMF Isp (Seconds)

RL10B-2 77 15.09 5.09 61.91 0.92 464

RS-25 77 14.47 4.47 62.53 0.93 452.3

J-2 77 12.78 2.78 64.22 0.96 421

Merlin D (Vac) 77 8.77 -1.23 68.23 1.02 348

10Mg Cargo Lander using the SLS Block 1A Launch Vehicle (No Tug)

Engine Initial Mass (Mg) Final Mass (Mg) Inert Mass (Mg) Prop. Mass (Mg) PMF Isp (Seconds)

RL10B-2 105 20.58 0.58 84.42 0.99 464

20Mg Cargo Lander using the SLS Block 1B Launch Vehicle (No Tug)

CARGO LANDERS WITH TUG

-Launch vehicle launches the cargo lander to LEO without payload

-Cargo lander goes from LEO to Cis Lunar Orbit following MD trajectory

-Payload is obtained from Tug in CLO, Lander descends to Lunar Surface

-Trajectory does not account for inclination change

Engine Initial Mass (Mg) Final Mass (Mg) Inert Mass (Mg) Prop. Mass Used (Mg) Excess Mass (Mg) Isp (Seconds)

RL10B-2 53 15.97 5.3 47.03 0.67 464

RS-25 53 15.46 5.3 47.54 0.16 452.3

10Mg Cargo Lander using the Falcon Heavy (With Tug)

Engine Initial Mass (Mg) Final Mass (Mg) Inert Mass (Mg) Prop. Mass Used (Mg) Excess Mass (Mg) Isp (Seconds)

RL10B-2 105 31.75 10.5 93.25 1.25 464

RS-25 105 30.73 10.5 94.27 0.23 452.3

20Mg Cargo Lander using the SLS Block 1B Launch Vehicle (With Tug)

REFERENCES

[1] Aerojet Rocketdyne. “RL10 Engine.” http://www.rocket.com/rl10-engine

[2] Aerojet Rocketdyne. “RS-25 Engine.” http://www.rocket.com/rs-25-engine

[3] Encyclopedia Astronautica. “J-2.” http://www.astronautix.com/engines/j2.htm

[4] SpaceX. “Falcon 9.” http://www.spacex.com/falcon9

BACKUP SLIDES

BACKUP SLIDES

% Code Created by Brock Miller clear all; close all; clc %Givens g0 = 9.8; %m/s %Inputs mpaymg = input('What is the Payload Mass in Mg? '); mpay = mpaymg * 1e6; Isp = input('What is the Isp of your LRE in seconds? '); menuvar = menu('What Launch Vehicle will be used?','Falcon Heavy','SLSB1A','SLSB1B'); if menuvar == 1 minitial = 53e6; %g elseif menuvar == 2 minitial = 77e6; %g else minitial = 105e6; %g end Delta_V = 7410; %m/s MR = exp(Delta_V/(g0*Isp)); %Initial Mass over Final Mass mfinal = (minitial/MR); minert = (mfinal - mpay); mprop = (minitial - mfinal); pmf = mprop/(mprop+minert); fprintf('\n----Lander Specs---- \n') fprintf('Initial Mass %4.2f Mg\n',minitial/1e6) fprintf('Final Mass %4.2f Mg\n',mfinal/1e6) fprintf('Inert Mass %4.2f Mg\n',minert/1e6) fprintf('Propellant Mass %4.2f Mg\n', mprop/1e6) fprintf('Propellant Mass Fraction %4.2f \n',pmf)

BACKUP SLIDES

% Code Created by Brock Miller clear all; close all; clc %Givens g0 = 9.8; %m/s Delta_V1 = 4750; %m/s Delta_V2 = 2656; %m/s mpaymg = input('What is the Payload Mass in Mg from Tug? '); mpay = mpaymg * 1e6; Isp = input('What is the Isp of your LRE in seconds? '); menuvar = menu('What Launch Vehicle will be used?','Falcon Heavy','SLSB1A','SLSB1B'); if menuvar == 1 mini = 53e6; %g elseif menuvar == 2 mini = 77e6; %g else mini = 105e6; %g End %LEO to CLO MR1 = exp(Delta_V1/(g0*Isp)); %Initial Mass over Final Mass mfinal1 = (mini/MR1); minert1 = mini/10; mprop1 = (mini - mfinal1);

%CLO to Surface With Payload mprop2 = mfinal1 - minert1; mini2 = mfinal1+mpay; MR2 = exp(Delta_V2/(g0*Isp));%Initial Mass Leg 2 over Final Mass Leg 2 mfinal2 = (mini2/MR2); mexcess = mfinal2 - mpay - minert1; mtotprop = mprop1+mprop2-mexcess; fprintf('\n----Lander Specs---- \n') fprintf('\n----LEO to CLO----\n') fprintf('Initial Mass %4.2f Mg\n',mini/1e6) fprintf('Final Mass %4.2f Mg\n',mfinal1/1e6) fprintf('Inert Mass %4.2f Mg\n',minert1/1e6) fprintf('Propellant Mass Used %4.2f Mg\n', mprop1/1e6) fprintf('Propellant Mass Remaining %4.2f Mg \n',mprop2/1e6) fprintf('\n----CLO to Surface----\n') fprintf('Initial Mass %4.2f Mg\n',mini2/1e6) fprintf('Final Mass %4.2f Mg\n',mfinal2/1e6) fprintf('Inert Mass %4.2f Mg\n',minert1/1e6) fprintf('Total Propellant Used %4.2f Mg\n', mtotprop/1e6) fprintf('Excess Mass %4.2f Mg\n', mexcess/1e6)

PROPULSION MATT SCHURMAN

Evaluating an Electric Transfer Vehicle

1/21/2016

ELECTRIC PROPULSION

ARCHITECTURE -Goal for the transfer vehicle is to move a 20-30 Mg payload from LEO to GEO

-The payload will then separate from the electric transfer vehicle and perform a chemical burn to the moon

-The electric transfer vehicle will spiral back to LEO to be mated with the next cargo mission

ELECTRIC TRANSFER VEHICLE

Specifications of the Electric Transfer Vehicle:

-109 NSTAR engines

-Providing a total 10 N of thrust (.092 N each)

-Propellant used is Xenon

-Uses 250 kW of power total

-For moving a 30 Mg payload:

-Propellant required: 7,351 kg

-Transit Time (Round trip): ~1 year

ELECTRIC TRANSFER VEHICLE

Physical Dimensions and Feasibility

-Engine diameter is 30 cm

-Approximately 148 engines could fit on a vehicle the same diameter as the core of a Falcon Heavy

-60% efficient – so the vehicle would need to dissipate ~100 kW

-Would need to be refueled in LEO between missions (ISS)

-7,351 kg of Xenon costs $8.8m

-Changes overall architecture to Falcon Heavies and Falcon 9’s instead of large numbers of SLS launches. In theory a ~75% reduction in cost.

EXTRA SLIDES

EXTRA SLIDES

PROPULSION ANDREW CULL

Nuclear Thermal Rocket

General LH2/LOX propellant mass function

NUCLEAR THERMAL ROCKET

Advantages

• ISP = 900 sec

• High thrust output

(980 kN)

• Hydrogen based

Disadvantages

• Not currently developed

• High development costs

• High risk for spreading

radioactive material

NERVA Diagram drawn by Hakusho Chin

PROPELLANT MASS ESTIMATION

General Propellant Mass Function for LH2/LOX For IMF = .25 Propellant Mass is 58.17 Mg For O/F = 6 LH2 Mass = 8.31 Mg LH2 Vol = 117.06m^3 LOX Mass = 49.86 Mg LOX Vol = 43.69m^3

REFERENCES

Sutton, George P., and Oscar Biblarz. Rocket Propulsion Elements. 8th ed. New York:

Wiley, 2010. Print.

"Liquid Oxygen." Wikipedia. Wikimedia Foundation, n.d. Web. 27 Jan. 2016.

<https://en.wikipedia.org/wiki/Liquid_oxygen>.

"Liquid Hydrogen." Wikipedia. Wikimedia Foundation, n.d. Web. 27 Jan. 2016.

<https://en.wikipedia.org/wiki/Liquid_hydrogen>.

BACK UP SLIDE 01

% AAE 450

% Andrew Cull

% LH2/LOX Calculations for Single Stage

% Constants

dV = 4000; %m/s

ISP = 450; %s %ISP from Centaur RL10 engine

g0 = 9.81; %m/s^2

mpay = 20000; %Kg

%EXECUTION

c = ISP*g0;

MR = exp(-dV/(g0*ISP));

% finert = minert/(mprop + minert) %Once we know the finert, we can narrow

% the mass of the propellant to a single value instead of a range

finert = linspace(0,MR,101);

PMF = 1 - MR;

mprop = (mpay*(exp(dV/c)-1).*(1-finert))./(1-finert.*exp(dV/c));

% Plot

plot(finert,mprop/1000)

hold on;

plot(MR,0,'rx','MARKERSIZE',16)

title('Propellant Mass versus Inert Mass Fraction','FontSize',16)

xlabel('Inert Mass Fraction','FontSize',16)

ylabel('Propellant Mass [Mg]','Fontsize',16)

legend('Propellant Curve','Mass Ratio','LOCATION','NORTHWEST')

% Exact Propellant Estimation

% x = input('Do you want to do a exact estimation? [1/0]');

% if x == 1

% Minertknown = input('Inert Mass [kg] = ');

% dVreq = input('Required Delta V [m/s] = ');

% mpayreq = input('Payload Mass [kg] = ');

% ISPengine = input('ISP for Engine [s] = ');

% c = ISP*g0;

% MR = exp(-dVreq/c);

% PMF = 1 - MR;

% syms mpropreq;

% mprop = double(solve(PMF == mpropreq/(mpropreq+Minertknown)));

% fprintf('Propellant Mass Required = %.2f [kg]\n',mprop)

% else

%

% end

MATLAB CODE

BACK UP SLIDE 02 Density of LOX = 1141 kg/m^3 Density of LH2 = 70.99 kg/m^3

POWER AND THERMAL NICK RAMSER

SOLAR CELL COMPARISONS • Monocrystalline, polycrystalline

• Older technology

• Efficiencies between 14% and 22%

• Multi-junction

• Forefront of technology

• Commercially available with 29%

efficiency

• Over 40% efficiency achieved in lab

CONCLUSIONS

• Solar cells will not be a

good choice as a main

power source for the lunar

base

• Even worse as a power

source on Mars

• Possible uses

• Auxiliary power

• Vehicles

• Satellites

• Cyclers

Power (kW) per

Solar Cell

Type

Mass

(kg)

Volume

(m^3)

Area

(m^2)

Cost

($)

Mono 0.32 63 0.06 0.08

Poly 0.53 100 0.21 0.38

Junction 0.57 34 0.48 0.48

Performance characteristics

EXTRA PLOTS

CODE % solarCellComp.m

%

% Authors: Nick Ramser

%

% Description:

% This script is meant to compare the mass, power,

volume, and costs

% of competing solar cell technologies. The values used

here are mainly

% estimations based on specification sheets of

commercially available cells

% and academic research.

%

% References:

% “29.5% NeXt Triple Junction (XTJ) Solar Cells,” May

2010.

% http://www.spectrolab.com/DataSheets/cells/PV XTJ

Cell 5-20-10.pdf

%

% Bagzhou, Y., “EE 446/646 Photovoltaic Devices III.”

% http://www.egr.unlv.edu/~eebag/Photovoltaic Devices

III.pdf

%

% “International Space Station Electric Power System

(EPS):”

% http://www.boeing.com/assets/pdf/defense-

space/space/spacestation/systems

% /docs/ISS Electric Power System.pdf

%

% Luque, A., and Hegedus, S., Handbook of photovoltaic

science and

% engineering, London: John Wiley & Sons, 2006.

clear all

close all

clc

% Characteristics of monocrystalline silicon cells

mono_eff = 24.7; % efficiency

mono_v = 0.615; % volts (V)

mono_i = 0.8; % current (A)

mono_p = 63; % power per area (W/m^2)

mono_t = .001; % thickness (m)

mono_m = 0.2; % mass per area (kg/m^2)

mono_c = 0.75; % $/W

% Characteristics of polycrystalline silicon cells

poly_eff = 20.3; % efficiency

poly_v = 0.615; % volts (V)

poly_i = 8.35; % current (A)

poly_p = 211; % power per area (W/m^2)

poly_t = 0.002; % thickness (m)

poly_m = .4; % mass per area (kg/m^2)

poly_c = .55; % cost ($/W)

% Characteristics of Triple Junction Cells

mj_eff = 40.7; % efficiency

mj_v = 2.6; % volts (V)

mj_i = 1.81; % current (A)

mj_p = 476; % power per area (W/m^2)

mj_t = .014; % thickness (m)

mj_m = 84; % (mg/cm^2)

mj_m = mj_m * (100^2/1000^2); % kg/m^2

mj_c = 1; % cost ($/W)

% Property vectors

eff = [mono_eff poly_eff mj_eff] ./ 100;

v = [mono_v poly_v mj_v];

i = [mono_i poly_i mj_i];

p = [mono_p poly_p mj_p];

t = [mono_t poly_t mj_t];

m = [mono_m poly_m mj_m];

c = [mono_c poly_c mj_c];

CODE % Parameters

n = 10; % data points

p_example = 500e3;

p_range = linspace(100e3,1000e3,n)'; % Desire power

(required) range (W)

% Computations

p_produced = p_range * eff.^(-1); % power to be

produced (W)

areas = zeros(n, 3);

volumes = zeros(n, 3); % volume of cells

required (m^3)

masses = zeros(n, 3); % mass of cells

required (kg)

costs = zeros(n, 3); % cost ($)

for i = 1:n

areas(i,:) = (p_produced(i,:) ./ p)

volumes(i,:) = areas(i,:) .* t;

masses(i,:) = areas(i,:) .* m;

costs(i,:) = (p_produced(i,:) .* c);

end

p_ex_prod = p_example .* eff;

power_mass = p ./ (m .* 1000)

power_vol = p ./ (t .* 1000)

ex_area = p ./ 1000

ex_cost = p ./ (c .* 1000)

p_range = p_range ./ 1000; % Convert to kW

% Plots

title_sz = 30;

label_sz = 20;

tick_sz = 15;

p_label = 'Power (kW)';

c_label = 'Cost ($)';

m_label = 'Mass (kg)';

v_label = 'Volume (m^3)';

a_label = 'Area (m^2)';

% volume v. power

pv_tit = 'Volume vs. Power';

easy_plot(p_range, volumes, p_label, v_label, pv_tit,

label_sz, ...

title_sz, tick_sz, 1, 'southeast');

% mass v. power

mv_tit = 'Mass vs. Power';

easy_plot(p_range, masses, p_label, m_label, mv_tit,

label_sz, ...

title_sz, tick_sz, 2, 'northwest');

CODE % cost v. mass

norm = 0;

for i = 1:n

if (masses(i,1) <= masses(n,3))

norm = i;

end

end

mc_tit = 'Cost vs. Mass';

easy_plot(masses, costs, m_label, c_label, mc_tit,

label_sz, ...

title_sz, tick_sz, 3, 'southeast');

% cost v. power

cp_tit = 'Cost vs. Power';

easy_plot(p_range, costs, p_label, c_label, cp_tit,

label_sz, ...

title_sz, tick_sz, 4, 'southeast');

% area v. power

pa_tit = 'Area vs. Power';

easy_plot(p_range, areas, p_label, a_label, pa_tit,

label_sz, ...

title_sz, tick_sz, 5, 'northwest');

REFERENCES

“29.5% NeXt Triple Junction (XTJ) Solar Cells,” May 2010. http://www.spectrolab.com/DataSheets/cells/PV XTJ Cell 5-20-10.pdf Bagzhou, Y., “EE 446/646 Photovoltaic Devices III.” http://www.egr.unlv.edu/~eebag/Photovoltaic Devices III.pdf “International Space Station Electric Power System (EPS):” http://www.boeing.com/assets/pdf/defense-space/space/spacestation/systems/docs/ISS Electric Power System.pdf Luque, A., and Hegedus, S., Handbook of photovoltaic science and engineering, London: John Wiley & Sons, 2006.

POWER AND THERMAL RACHEL LUCAS

January 28, 2016

Power Estimates, Experimental Reactor Technology

POWER ESTIMATES

Power Requirement Estimates

• The ISS produces 84 kW to support 6 people

• A maximum of 16 astronauts living on the Moon at once gives 224 kW

• Approximate 200 kW extra power for experiments, mining resources, ….

• Gives total power of approximately 424 kW

Type of Reactor Approximate Power Generation

Pressurized Water 1000 MW

Boiling Water 1000 MW

Pressurized Heavy Water 800 MW

Advanced Gas-Cooled 1100 MW

Light Water Graphite- Moderated 1000 MW

Fast Neutron Reactor 300-800 MW

Radioisotope Thermoelectric Generator 25-3000 W

EXPERIMENTAL REACTOR TECHNOLOGY

Experimental Reactor Technology

• Current experimental technology capable of delivering power in the kW range

• SAFE-400 has the highest electric power output per megagram

• 5 models would need to be used in order to produce the needed amount of power

• Total weight = 2.71 Mg

Reactor Electric Power (kWe)

Thermal Power (kWt)

Mass (Mg) Electric Power per Megagram (kWe/Mg)

SAIRS 111 407.3 2.98 37.25

HOMER-15 3 15 0.214 14.02

SAFE-400 100 400 0.541 184.84

SP-100 20 600 ~3 6.67

HPCMR 112 1600 ~3.2 35

BACKUP SLIDES

BACKUP SLIDES

References [1] “Facts and Figures,” NASA, International Space Station, November 3, 2014. [http://www.nasa.gov/mission_pages/station/main/onthestation/facts_and_figures.html#.VMj9aGPN58E. [2] El-Genk, Mohamed S., Tournier Jean-Michel P., “’SAIRS’ – Scalable Amtec Integrated Reactor Space Power System," Progress in Nuclear Energy, Vol. 45, No. 1, pp. 25~59, 2004 [3] Postona, David I., Kapernicka, Richard J., Guffeeb, Ray M., Reid, Robert S., Lipinski, Ronald J., Wright, Steven A., Talandis, Regina A., “Design of a Heatpipe-Cooled Mars-Surface Fission Reactor," AIP Conference Procedings, 608, 2002, pp. 1096-1106. [4] Postona, David I., Kapernicka, Richard J., Guffeeb, Ray M., “Design and Analysis of the SAFE-400 Space Fission Reactor," AIP Conference Procedings, 608, 2002, pp. 578-588.

POWER AND THERMAL WERONIKA JUSZCZAK

Thermal Systems

Heat Transfer Analysis

THERMAL SHIELD • Multilayer Insulation (MLI)

• Heat transfer by a combination of conduction and net radiation

• Net radiation (difference between emitted and absorbed) for n layers at specified emissivity

• Does not consider heat expelled by electronic systems and astronauts

• Material and number of layers can be changed

Temperature on Moon [K]

Minimum Total q (40 Layers) [Watts/m2]

Maximum Total q (1 Layer) [Watts/m2]

96 (lunar night)

0.0388

1.55

294 (solar flux)

-1.97x10-4 -0.00790

MODEL OF HEAT TRANSFER

BACKUP SLIDES MATLAB CODE

%% Estimating Heat Transfer per unit Area across a Surface

%% Weronika Juszczak

% Finding Surface Temperature of Walls

n = 1:40; % Layers of material

L = 0.0127 .* n; % thickness of material [m]

k = .0001; % k of material (aluminum) [Watts/m*K]

h1= 5; % conduction coefficient

h2= 5;

Th = 293; % Temperature of habitat [K]

Tm = 96; % Temperature outside [K]

Rcv1 = 1/h1; % Resistance of convection between inside hab and inside wall

Rcn = L/k;

Rcv2 = 1/h2; % Resistance of convection between outside wall and moon atmosphere

Rtot = Rcv1 + Rcn + Rcv2; % Total Resistance

q1 = (Th - Tm) ./ (Rtot); % estimated total heat transfer used to find temperature of walls

Ts1 = Th - Rcv1*q1; % Surface Temperature of Inside Wall [K]

Ts2 = Tm - Rcv2*q1; % Surface Temperature of Outside Wall [K]

%% Conduction and Radiation Heat Transfer

tdiff = Ts1 - Ts2; % [K]

qc = k.*tdiff./ L; % heat transfer per area for convection heat transfer [Watts/m^2]

%% Radiation Heat Transfer:

e = 0.04; % emissivity

boltz = 1.38064852e-23; % Boltzman constant

qr = (e./((n+1).*(2-e))).*boltz.*(Ts2.^4-Tm.^4); % heat transfer per unit area for radiation heat transfer [Watts/m^2]

figure;

plot(n, qr, n, qc,n,qr+qc,'--')

legend('qr','qc','qtot')

title('Heat Transfer per Unit Area','FontSize',20)

xlabel('Number of MLI layers (n)')

ylabel('Heat Flux Density [Watts/m^2]')

REFERENCES

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20100034929.pdf http://www.nss.org/settlement/moon/library/LB2-701-ThermalControlSystem.pdf

CONTROLS TEAM ZARIN BARI

Control System of the Command Module (XM2/3):

- Reaction Wheels versus CMGs

- Desaturation Techniques

REACTION WHEEL OR CMG?

Reaction Wheel

Smaller vehicles communication

satellites (1 Mg -5 Mg)

*Based off Rockwell Collin’s RSI 68 RW

dimensions

MAX. TORQUE ON COMMAND MODULE = 275.733 NEWTON-METERS

Control Moment Gyroscope

Larger vehicles command module

(XM2/3 w/ mass 14 Mg)

*Double-Gimbal CMG from L-3

dimensions

Control System Reaction Wheel

Mass [Mg] 0.0077 / RW

Power [W] 220 / RW

0.0121 / RW

Max. Torque [Nm] 0.215 / RW

Control System CMG

Mass [Mg] 0.272

Power [W] 276

1.62

Max. Torque [Nm] 258

DESATURATION DESATURATION OF THE SYSTEM NEEDS TO BE DONE WITH A MOMENTUM CHANGE

Saturated Desaturated

Storage capability for CMG: 4880 Nms Nominal Speed: 6,600 rpm Number of CMG: 1 per XM Need to be desaturated about 3 times a day

REFERENCES Aerospace, V. (2014). Reaction Wheel VRW-1. Retrieved from http://www.vectronic-

aerospace.com/space.php?p=reactionwheel

Communications, L. (n.d.). CMG — Control Moment Gyro Space & Navigation Double-

Gimbal CMG. Retrieved from http://www2.l-

3com.com/spacenav/pdf/datasheets/L3SpNav_CMG_sellsht_9-14rev1.pdf

Frost, R. (2013, December 10). Saturation - Explained. Retrieved from

https://www.quora.com/What-does-it-mean-that-a-reaction-wheel-can-store-external-

torques-as-angular-momentum-in-wheels-over-time-until-saturation

Memi, E., & Deason-Sharp, T. (n.d.). Boeing Motion Control Subsystem. Retrieved from

http://www.boeing.com/assets/pdf/defense-space/space/spacestation/systems/docs/ISS

Motion Control System.pdf

Votel, R., & Sinclair, D. (n.d.). Comparison of Control Moment Gyros and Reaction Wheels

for Small Earth-Observing Satellites. Retrieved from

http://digitalcommons.usu.edu/cgi/viewcontent.cgi?article=1080&context=smallsat

BACKUP SLIDE 1

BACKUP SLIDE 2 Code used to calculated mass, power, volume values for the reaction wheel and CMG

%% ZARIN BARI

clear;

clc;

% values of the BA 330 (XM2 is analyzed here)

m = 14000; % mass of XM2 in kg

% CMG calculations (L-3)

l = 51 * 0.0254; %convert from inches to meters

w = 54 * 0.0254;

h = 48.5 * 0.0254;

vol_cmg = pi*((l/2)^2)*h

flywheel_speed = 6600; %rpm

freq = flywheel_speed / 60; %frequency of the CMG (Hz)

ang_vel = 2*pi*freq; %rad/s

m_flywheel = 99.7903; %kg

I = 0.5*m_flywheel*(l/2)^2; %moment of inertia approx by kg-m^2

K = 0.5*I*ang_vel*2 %kinetic energy (J)

max_torque = 258; %Nm

ang_mom = 4880 %4760; %Nms

power_cmg = ((20*max_torque)/ ang_mom^0.6) + (4.51*ang_mom^0.47) %watts

% Reaction Wheel calculations (Rockwell Collins)

diameter = 0.31; %m

height = 0.16; %m

vol_rw = pi*((diameter/2)^2)*height

ang_mom_RW = 15; %rockwell collins value [Nms]

max_torque_RW = 0.215; %Nm

power_RW = 1000*max_torque_RW + (4.51*h^0.47) %watts

CONTROLS CHAD OETTING

Command Module (XM2/3)

• Reaction Control System Thrusters

REACTION CONTROL THRUSTERS Thruster Thrust (N) Mass (kg/ea)

Propellant Mass (Mg) at 100m/s budget

22N Bipropellant Thruster [1] 22 0.68 0.47

200N Bipropellant Thruster [1] 200 1.9 0.52

R4D Thruster [2] 490 3.63 0.45

Uses: • Attitude control

• Stabilization • Orientation

• Maneuvering

TOTAL SYSTEM MASS

Thruster Total Thrusters Total System Mass (Mg)

22N Bipropellant 20 0.48

200N Bipropellant 20 0.56

R4D 20 0.52

Arbitrary total number of thrusters currently

Total system mass based on assumed DV budget of 100 m/s

BACKUP SLIDE

22N Bipropellant [1] Isp: 300s Fuel: MMH Oxidizer: N2O4 Used on: under development

200N Bipropellant [1] Isp: 270s Fuel: MMH Oxidizer: MON-3 Used on: ESA’s ATV, NASA’s Orion

R4D [2] Isp: 312s Fuel: MMH Oxidizer: N2O4 Used on: Apollo Service and Lunar Modules

Resources: [1] Airbus Defense and Space Systems. “Space Propulsion, Chemical Bi-Propellant Thruster Family,” http://www.space-propulsion.com/brochures/bipropellant- thrusters/bipropellant-thrusters.pdf [retrieved January 2016] [2] Encyclopedia Astronautica. “R-4D,” http://www.astronautix.com/engines/r4d.htm [retrieved January 2016]

BACKUP SLIDE

STRUCTURES AUSTIN BLACK

Water Radiation Shield and Storage

PROBLEM DETERMINING EFFECTIVENESS AND AMOUNT OF WATER REQUIRED

Water Shield

Thickness: 0.501 meters

Volume: 20.69 m3 (5,465 gallons)

Mass: 20.652 Mg

Aluminum Support

Thickness = .05 m

Mass =10.582 Mg

Need layer between water and

aluminum to avoid corrosion.

Water

Habs

Aluminum

WATER CONTAINMENT

Indoor Water Storage

“Collapsible Water Reservoirs”

3-4 per hab

V ~ 100 liters (~100 kg, 0.1 m3)

t = 12 cm (50% attenuation for 1500 keV

gamma radiation)

Outdoor Water Storage

Storage Capacity: 10 Mg

1.05 Mg per launch of ferrying vehicle

3.2 Mg per year consumed by crew of 16

Volume = 10,020.04 m3

HOUSING WATER FOR CREW, ISRU APPLICATIONS, AND ROCKET FUEL

CWR

Lunar Hab

t

ADDITIONAL SLIDES

rho = .998; %water density

a = 50; %shielding mass/area ratio

d = 7; %hab diameter

t = (a/rho)/100 %water shield thickness

V = ((1/12)*pi*((d+t)^3))-((1/12)*pi*(d^3)) %Shield Volume

m = V*rho*1000 %Shield Mass

gallons = V*264.172 %Volume in gallons

SA = 2*pi*(d/2)^2 %Dome Surface Area

V_hemi = (2/3)*pi*(3.55^3)-(2/3)*pi*(3.5^3) %Dome Volume

m_hemi = V_hemi*2780 %Dome Mass

10 Mg = 10,000 kg / 0.998 kg/m3 = 10,020.04 m3

REFERENCES

Barghouty, A. F., and S. A. Thibeault. "The Exploration Atmospheres Working Group's Report on Space Radiation Shielding Materials." Nasa.gov. NASA, Sept. 2006. Web. 26 Jan. 2016.

McAlister, Daniel R. "Gamma Ray Attenuation Properties of Common Shielding Materials." www.eichrom.com. PG Research Foundation, Inc., 3 Jan. 2013. Web. 24 Jan. 2016.

Project Aldrin - Purdue

Rask, Jon, Wenonah Vercoutere, Barbara J. Navarro, and Al Krause. "Space Faring. The Radiation Challenge." Nasa.gov. NASA. Web. 27 Jan. 2016.

"Shielding of Ionizing Radiation - Nuclear Power." Nuclear Power. Web. 27 Jan. 2016.

"Space Craft Liquid and Gas Containers." Space Craft Liquid and Gas Containers. Web. 26 Jan. 2016.

STRUCTURES ADRIAN PANSINI

Preliminary Landing Gear Strut Analysis

LANDING GEAR PRIMARY STRUT ANALYSIS

• Modeled as a composite cylinder

• Stresses found due to a load of 4087500 N

• Assume four landing struts equally holding 10 Mg lander on Moon surface

• Load applied 12 degrees from vertical

• Analysis shows that Carbon Fiber composite can withstand initial test load

Load P = 4087.50 kN

Fails around 600 MPa

LANDING GEAR

• Interpolated experimental data to estimate effects of cold temperatures

on composite material

• At -175 degrees Celsius, stiffness is lowered by 80%

• Acceptable landing loads decrease as the landing site gets colder

TEMPERATURE EFFECTS

BACKUP SLIDES

Experimental values from Ref. [3], interpolated for low temperature from Science Team Presentation last week.

BACKUP SLIDES TEST LOAD APPLICATION

• Test load applied as though landing on a 12 ⁰

incline

Load P

12⁰ angle

BACKUP SLIDES REFERENCES

• [1] Landing gear initial design/ landing conditions

• https://www.hq.nasa.gov/alsj/tnD6850LMLandingGearSubsytem.pdf

• [2] Carbon Fiber mechanical properties

– http://www.performance-composites.com/carbonfibre/mechanicalproperties_2.asp

• [3] Temperature Effects

• http://web.ornl.gov/~webworks/cppr/y2001/rpt/117096.pdf

• [4] Assumed lander weight

• Project Storyboard – Systems Team

STRUCTURES AMIT SONI

Surface Hab Radiation Shielding Concepts Using Lunar Regolith

131

DOME CONCEPT

Essential Points:

• Radiation and micrometeorite

protection.

• 0.15m of Regolith enough to

reduce micrometeorite impact

risk to 0.01% over 2-5 yr stay.

• Estimated Mass at Launch

~10.6 Mg/dome. **

• Estimated Total Mass w/

Regolith ~ 294.5 Mg/dome.

FOR INDIVIDUAL HAB MODULES

0.01m 0.01m

0.50m

Vectran™ HT

Lunar Regolith

* CATIA model by Amit Soni ** Assumed D=7.6m X L=12m Cylindrical Hab

132

BURYING HABS

• Cover Hab with 0.5m thick

layer of Regolith.

• Estimated Regolith Mass ~

1101.3 Mg/Hab. **

• Compared to Dome

• Pros:

o No prior shielding

structure needed.

o Not geometry

dependent.

• Cons:

o Large mass for hab

to support.

* CATIA model by Amit Soni ** Assumed D=7.6m X L=12m Cylindrical Hab

133

DOME BACK-UP SLIDE

Vectran™ HT density: 1.4 g/cm3 (from Project Aldrin-Purdue Radiation Dome) Lunar Regolith Density: 1.5 g/cm3

Vectran™ thickness: 0.01m = 1cm on both ends of dome. Vectran™ volume calculated from CATIA: 7.575 m3 Total mass: 7.575 m3 * 1400kg/m3 = 10.605 Mg Assumptions: Vectran™ walls self standing (no support structures internally and externally added for this analysis), Lunar Regolith of uniform consistency.

Vectran™ HT density: 1.4 g/cm3

0.5 m regolith thickness from “Radiation Background Slide” by Kate Fowee 0.15 min. m thickness of Regolith for micrometeorite shielding from Rockwell Study (Lewis 1992).

134

BURYING HAB BACKUP SLIDE

Lunar Regolith Density: 1.5 g/cm3

Assumed D=7.6m X L=12m Cylindrical Hab.

Did not consider mass of standard shielding/insulation on hab. Regolith volume calculated from CATIA: 734.239 m3 Regolith mass: 734.239m3 * 1500kg/m3 = 1101.3 Mg. Assumptions: Lunar Regolith of uniform consistency, uniform thickness of Regolith around hab (except around bottom half)

135

REFERENCES

Kurray America Inc. (2006). Vectran™[Brochure]. Fort Mill, SC:N.P.

Lewis, R.H. (1992). Human Safety in the Lunar Environment. Retrieved from

http://www.nss.org/settlement/nasa/spaceceresvol4/human.html

Rapp,D. (2006). Radiation Effects and Shielding Requirements in Human Missions to the

Moon and Mars. The International Journal of Mars Science and Exploration, Mars 2, 46-

71.

Wilson, J.W., Miller, J., Konradi, A., & Cucinotta, F.A. (Eds). (1997). Proceedings from

NASA: Shielding Strategies for Human Space Exploration. Hampton, VA: NASA Langley.

136

HUMAN FACTORS RACHAEL HESS

Aeroponic Crop Production

AEROPONICS

• Growth of plants in nutrient rich mist

• Plants selected: Broccoli, Carrots,

Green Beans, Leaf Lettuce, Pears,

Potatoes, Soybeans, and Sweet Corn

–Based on nutrients and growth

viability

• Humidity must remain between 70%

and 80%

• 14 hours of light and 10 hours of dark

for all plants

• Red and Blue light needed

Figure 1: Schematic of vertical aeroponic system, adapted from Tower Garden, Ref. 1

ROOM LAYOUT AND SPECIFICTIONS

• 3x3m room

• 8 growing systems

• A total of 288 plants

• One crop per tower

• Total water needed: 151.42 L

• Volume of growing space:

16.54 m3

• Approximate power needed

with lights on: 382 W (256 W

when lights are off)

• Total Mass: 733.15 kg

Figure 2: layout of aeroponic systems

RESOURCES

[1] “Facts about future growing’s aeroponic tower garden technology,”

http://www.futuregrowing.com/info.html [retrieved 19 Jan 2016]

[2] Kliss, M. and MacElroy, R., "Salad Machine: A Vegetable Production Unit for Long

Duration Space Missions," SAE Technical Paper 901280, 1990

[3] “Nutrition Information for Raw Fruits, Vegetables, and Fish,”

http://www.fda.gov/Food/IngredientsPackagingLabeling/LabelingNutrition/ucm063367.htm

[retrieved 20 Jan 2016]

[4] Clawson, J., Hoehn, A., Stodieck, L., Todd, P. et al., "Re-examining Aeroponics for

Spaceflight Plant Growth," SAE Technical Paper 2000-01-2507, 2000

BACK UP SLIDES NUTRIENT INFORMATION ABOUT AEROPONIC COMPATIBLE CROPS

BACK UP SLIDES POWER AND MASS CALCULATIONS

HUMAN FACTORS KELLY KRAMER

Volume of habs

Floorplan of hab system

January 28,2016

AREA AND VOLUME OF HABS Room Floor area (m2) Ceiling height (m) Utilized area volume (m3)

Bedroom (2 person) 11.44 3.05 34.88

Bathroom (per 2 people)

3.01 3.05 9.17

Common area (per 4 people)

20.57 3.05 62.74

Recreation area 390.19 6.00 2341.14

Medical station 7.51 3.05 22.88

Food storage/preparation

90.25 3.05 275.08

Garage (per 2 exploration rovers)

37.21 4.00 148.84

Laboratory/work station

37.21 4.00 148.84

Total Volume = 3043.647 m3

HAB SYSTEM FLOOR PLAN 6.1m 6.1m

6.1

m 6

.1m

2.74m

2.44m

3m

3m 5.01m 1.49m

2.89m 1.52m

2.89m

25.603m

15.24m

BACKUP SLIDE 1

0

5

10

15

20

0 2 4 6 8 10 12 14

Hab

itab

le v

olu

me

per

cre

wm

emb

er (

m3

)

Mission Duration (months)

Habitable volume vs. Mission duration

Tolerable limit Performance limit Optimal

Based on Figure 6-2 from Human Spaceflight: Mission Analysis and Design Reference: Larson, W. J., and Pranke, L. K., “Human Spaceflight: Mission Analysis and Design”, Analyzing Space and Surface Elements, McGraw-Hill Companies, Inc.New York, 1963, pp149.

Kitchen dimensions in food prep/storage area, based on figure from houseplanshelper.com Reference: “Kitchen Triangle Dimensions”, Kitchen Dimensions, House Plans Helper, 2016.

BACKUP SLIDE 2

In accordance with the International Residential Code (IRC), the minimum side-to-side clearance from the centerline of a toilet to the nearest wall is 15 inches. The smallest dimension can 30 inches. There must also be 21 inches in front of the toilet. Minimum shower area is 30 inches by 30 inches. Reference: International Code Council. International Residential Code, “Section R307 Toilet, bath, and shower spaces”, Chapter 3: Building Planning