1

PROJECT ARTEMIS | AAE 450

WEEK 05 PRESENTATIONS 2/13/2014

SCHEDULE

2

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

Presentation Schedule

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.

3

PARTH SHAH | APM

INITIAL RISK ASSESSMENT

o SINGLE LAUNCH LV RELIABILITY

o GETTING INTO LEO

o LUNAR ORBIT INSERTION

2/13/2014

MISSION SURVIVABILITY FOR KEY EVENTS

4

Mission Requirements: • >90% colonists’ mission success

• >95% returning crew home safe

Currently analyzing Phases 1-3, 7 • Phase 3 = Phase 7

Most data from launch history reports, NASA logs

Phase 4 and 6 has human element

Analyzing Phase 5 in detail with Event Tree • Defining success and catastrophic events

Parth Shah | APM

Events/Phases Probability of Success

1. Single Launch Capability of LV ~ 0.5 – 0.98

2. Flight from Earth to LEO ~ 0.990(0.658)

3. LEO to Low Lunar Orbit (LLO)[2] ~ 0.990

4. Landing on the Moon ---

5. Living on the Moon ---

6. Launch from Moon ---

7. LLO to LEO[2] ~ 0.990

8. Earth Re-entry + Landing ---

Launch Vehicle Successes

(kLV)

Attempts

(nLV)

No. Launches

Planned

Single Launch LV

Reliability (pLV)

Single Launch

LV Failure

Overall

Reliability

Atlas V (401) 18 19 2 0.9048 0.0952 0.9909

Falcon 9 8 12 8 0.6429 0.3571 0.9997

SLS (Based on Saturn V) 12 13 11 0.8667 0.1333 1.0000

Falcon XX 0 0 11 0.5000 0.5000 0.9995

PHASE 1 Bayesian 1st order

Probability of Success

Failure to Reach

LEO

Guidance and

Navigation Issues Launch Failure

LV Single Launch

Failure Rate

Launch Process

Issues

Fro

m L

V R

elia

bility

Weath

er Issu

es

Dela

yin

g L

au

nch

Ro

cket P

rep

ara

tion

Dela

yin

g L

au

nch

Co

llision

Avo

idan

ce

Failu

re

Tra

jecto

ry C

orre

ction

Failu

re

Co

ntro

l Syste

ms F

ailu

re

Missed

Launch

Window

P(fail) = 0.1

P(fa

il) = 0.3

P(fa

il) = 0.0

01

P(fail) = 0.301

P(fail) = 0.311 P(fail) = 0.001

P(fa

il) = 0.0

1

P(fa

il) = 0.0

1

P(fa

il) = 0.0

1

P(fail) = 0.03

P(fail) = 0.342

MISSION SURVIVABILITY FOR KEY EVENTS

PHASE 2

7

KRISTA GARRETT | MISSION DESIGN

TRAJECTORIES

o CREW MOON LANDER ASCENT PROFILE

o TRANS-EARTH INJECTION ΔV

o OTHER WORK: RADIATION EXPOSURE IN UNSHIELDED

ROVERS

2/13/2014

CREW MOON LANDER ASCENT PROFILE

8

Vertical rise to 10 km

Two-Point Boundary-Value Problem • Minimal time = minimal

propellant

• Reach desired altitude and velocity

Results: • Time for lunar ascent:

5 minutes, 10 seconds

• Propellant needed: 7.74 Mg

• Initial mass for ascent: 10.80 Mg

Krista Garrett | Mission Design

TRANS-EARTH INJECTION ΔV

9

Δv = 0.9705 km/s

Re-entry parameters:

• Altitude = 123.833 km

• Velocity = 10.9620 km/s

• Flight Path Angle = -5.7977°

Krista Garrett | Mission Design

Hyperbolic Trajectory at Moon (Krista Garrett)

Patched Conics Return (Krista Garrett)

11

MICHAEL CREECH | MISSION DESIGN

LUNAR OUTPOSTS

o OUTPOST LOCATIONS

o ENERGY COSTS FOR TRAVERSING THE LUNAR SURFACE

2/13/2014

OUTPOST LOCATIONS

12 Michael Creech | Mission Design

7 outposts per colony

Both mare and craters

Selected for optimal path between habitats and checkpoints

MISSION PROFILE AND ENERGY

13

Mission profile for a rover path

3D view of mission profile

Power profile and total energy requirement

Example – Shackleton to checkpoint 3

Total Energy = 3.091 MJ

Michael Creech | Mission Design

0 50 100 150 200 250 300 35023

23.5

24

24.5

25

25.5

26

26.5

27

27.5

Horizontal Distance [km]

Pow

er

[kW

]

Mission Profile - Power Map

26

JOSE MIGUEL BLANCO | MISSION DESIGN

LOW THRUST TRAJECTORIES

o LOW THRUST TRAJECTORY UPDATE, PATCHED TWO BODY

MOTION IMPLEMENTED

o MASS, POWER AND VOLUME FOR CARGO VEHICLE POWERED

BY ELECTRIC PROPULSION

2/13/2014

MASS POWER AND VOLUME OF CARGO VEHICLE

27

simulation

Jose Miguel Blanco / Mission Design / Cargo and Lander

power

volume Mass vs mission time

CARGO MASS TO LUNAR SURFACE

28 Jose Miguel Blanco / Mission Design / Cargo and Lander

mass volume

Mass on lunar surface

39

PROJECT ARTEMIS | AAE 450

WEEK 05 PRESENTATIONS 2/13/2014

SCHEDULE

40

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

Presentation Schedule

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.

41

SPENSER GUERIN | CONTROLS

CARGO VEHICLE

o ELECTRONIC CONTROL SYSTEM SIZE

o REACTION CONTROL THRUSTER SIZING

2/13/2014

ELECTRONIC CONTROL SYSTEM

42

Computing power still tentative.

Spenser Guerin | Controls

COMPONENT POWER [W] MASS [kg] VOLUME [] OPERATING

TEMPERATURES [°C]

Inertial Measurement

Unit (IMU) 22 4.5 7,206 -30→65

Star Tracker (x2)

24 7 15,776 -20→50

Altimeter 20 3.5 7,000 N/A

Computing 50 2 7,000 N/A

Total 116 17 36,982 -30→50

REACTION CONTROL THRUSTER SIZING

43

Allocate 50 m/s for trajectory and attitude control and 20 m/s for lunar orbit insertion clean up.

Spenser Guerin | Controls

Isp = 220 sec

m_pl = 40 Mg

m_p = 1.3 Mg

44

RYAN ALLEN | CONTROLS

CREW TRANSPORT VEHICLE

o HUMAN MOON LANDER SPECIFICATIONS

o RADIO INTERFEROMETER ARRAY

2/13/2014

HUMAN MOON LANDER SPECS

45 Ryan Allen | Controls

Human Moon Lander Specification

Value

Mass to LEO (crew included)

37.79 Mg

Power Required 462.8 W

Volume to LEO 560 m3

Navigation System

Total Value

Mass *25.4 kg

Power *150.8 W

Volume *0.05 m3

CATIA Model from Scott Sylvester

Human Lunar Lander Navigation System: 1) Computer System 2) Inertial

Measurement Unit (IMU)

3) 2x Star Tracker 4) Altimeter *Values from Spenser Guerin

RADIO INTERFEROMETER ARRAY CUBE

46 Ryan Allen | Controls

• Field of 30 Radio Interferometer Array Cubes (RIA Cubes)

• Charge battery during lunar day using solar panels

• Operate during lunar night • Communicate with L2

satellite • Deployed by Heavy Rover

30 RIA Cubes Value

Mass 3.0 Mg

Power 15.75 W*

Volume (packed) 30.0 m3

*Value from Eric Menke

Figure based on Murchison Widefield Array

HANI KIM | HUMAN FACTOR

HEAVY & LIGHT ROVER 2/13/2014

o LIGHT ROVER AND HABITAT ENERGY ESTIMATION

HEAVY & LIGHT ROVER

48 Hani Kim 2/12/2014

Fig. 2:Light rover body

Fig. 1:Light rover Dimension

Fig. 3:Light rover inside

Cabin Oxygen Tank

Material Aluminum O2 0.02 m3

Empty cabin mass 1.594 Mg N2 0.092 m3

Inside Volume 10.623 m3 Total Volume 0.112 m3

Material Volume 0.588 m3

Burn up in the atmosphere (x)

Microwave freeze drying • 1~30 kW

• ~1 m3

• ~50 kg

Other waste : 13390 kg

HUMAN FACTOR

49 Hani Kim 2/12/2014

Weight of solid waste(4)

Wet Weight Dry weight

kg/person/day 0.10 0.02

Per colony for mission (kg) 1251.43 256.54

Habitat Energy consumption (4)

Energy provided : 88 kW

Energy (kW)

Lighting 0.918

Water recycle (6) 2.5

Oxygen system (7) 1.5

Game room 3

Computer 0.5

Total 6.918

Heavy Rover (For 2 crews) (1)

Hours of Available Weight(kg) Volume(m3)

Water 10 2 0.002

Food 10 1.5 0.006

total 10 3.5 0.008

Heavy Rover (For 4 crews) (1)

Hours of Available Weight(kg) Volume(m3)

Water 5, 10 3 0.003

Food 5, 10 2.2125 0.009

total 5, 10 5.2125 0.012

Light Rover (2 crews) (1)

Hours of Available Weight(kg) Volume(m3) Water 360 72 0.072 Food 360 53.1 0.2214 total 360 125.1 0.2934

52

BEN FISHMAN | HUMAN FACTORS

LIFE SUPPORT/OXYGEN

o PRESSURIZED ROVERS

2/13/2014

HF: PRESSURIZED ROVER (LIGHT)

53 Ben Fishman | Human Factors

Number of People Hrs Spent

Mass O2 [kg]

Mass N2 [kg]

Total Mass [kg]

Volume O2 [m^3]

Volume N2 [m^3]

Total Volume [m^3]

2 24 16.6 19 35.6 0.005 0.017 0.022

4 (emergency) 24 18.2 23 41.2 0.009 0.034 0.043

Number of People Hrs Spent

Mass O2 [kg]

Mass N2 [kg]

Total Mass [kg]

Volume O2 [m^3]

Volume N2 [m^3]

Total Volume [m^3]

2 360 83 95 178 0.025 0.085 0.11

Heavy Rover (Updated)

Light Rover (First iteration)

Power = 120 W/tank kept at 3000 psi

HF: CTV OXYGEN/FIGURES

54 Ben Fishman | Human Factors

Days

Mass (O2)

[kg]

Mass (N2)

[kg]

Total Mass

[kg]

Volume

(O2)

[m^3]

Volume

(N2) [m^3]

Total

Volume

[m^3]

4 72.8 124.2 197 0.072 0.268 0.34

5 91 155 246 0.09 0.335 0.425

6 109.2 185.8 295 0.108 0.402 0.51

7 127.4 216.6 344 0.126 0.469 0.595

• Power = 120 W/tank kept at 3000 psi • Still working on placement/storage of oxygen on CTV • Starting work on Habitat oxygen

Photo credit (teammate): HaNi Kim

58

JESSICA CALLINAN | HUMAN FACTORS

RESUPPLY MISSIONS

o RESUPPLY MISSION MASS AND VOLUME OPTIONS

o CLOTHING VOLUME UPDATE

o MEDICAL DEVICES

o TYPES OF PEOPLE TO SEND

2/13/14

RESUPPLY MISSIONS

59

Multiple options for resupply per colony

* The optimal resupply is 6 months for food, as food losses its nutrients drastically after one year of storage. However, water can stay for the entire mission. [7] **Water mass after initial 5.965 Mg and 5.965 m3 are there

Jessica Callinan | Human Factors

Resupply 3 months 6 months* 9 months 12 months

Food mass (Mg) 1.456 2.912 4.384 5.840

Water mass (Mg)**

0.2786 0.5572 0.8358 1.114

Total mass (Mg) 1.735 3.469 5.220 6.954

Food volume (m3)

6.067 12.13 18.27 24.33

Water volume (m3)

0.2786 0.5572 0.8358 1.114

Total volume (m3)

6.345 12.69 19.10 25.44

CLOTHING | MEDICAL | PEOPLE |

60

Clothing per colony for 2 week cycles vacuum packed[1] • Volume (m3): 0.1011

• Mass (kg): 79.05

Medical devices • IntraVenous Fluid Generation

(IVGEN) [6][8][4]

0.002163 m3 [2]

• Advanced Life Support Pack (ALSP) [2]

0.0152 m3

• Reference International Space Station Integrated Medical Group Medical Operations Book (ISS IMG)[3][5]

People • Assuming all have at minimum

Masters degree in their field

• 2 engineers, 2 doctors, 2 scientists, 2 psychologists

Jessica Callinan | Human Factors

Figure: Advanced Life Support Pack, Jessica Callinan

69

ERIC MENKE | COMMUNICATIONS

VEHICLE SATELLITE DISHES

o DISH SIZING

o NEW HEAVY ROVER DIMENSIONS

2/13/2014

VEHICLE DISHES

70 Eric Menke | Comm

Link Budget L2 Satellite Vehicle to L2 Interferometer to L2 Unit

Carrier Frequency 4 4 4 GHz

Transmitter Power 245 165 0.525 Watt

Transmitter Line Loss -1 -1 -1 dB

Transmitter Antenna Beamwidth 10 20 5 deg

Transmitter Antenna Pointing Offset 0.4866 5 0.25 deg

Distance Between Transmitter and Receiver 68181 68181 68181 km

Receiver Antenna Diameter 0.8 3 3 m

Receiver Antenna Beamwidth 20 10 20 deg

Receiver Antenna Pointing Error 5 0.4866 5 deg

Desired Data Rate 3.00E+06 7.00E+06 5.00E+05 bps

Bit Error Rate 1.00E-05 1.00E-05 1.00E-05 probability

Implementation Loss -2 -2 -2 dB

Transmitter Antenna Diameter 3 0.8 0.09 m

Gain Margin 3.0166 3.0801 3.0493 dB

Dish Volume 0.206 0.004 9.11E-05 m3

Dish Mass 300 20 3 kg

HEAVY ROVER

71 Eric Menke | Comm

Cabin Volume 3 m3

Shielding Volume 14.78 m3

Total Volume 18.133 m3

Battery Mass 10.87 Mg

Empty Shielding Mass 2.407 Mg

Total Mass 15.558 Mg

73

CAMERON HORTON | AERODYNAMICS

SAMPLE CARRIER RE-ENTRY

o SCIENCE SAMPLE CARRIER LAUNCH PLAN

o POD DESIGN AND SPECS

o RE-ENTRY CALCULATIONS

2/13/2014

SAMPLE CARRIER LAUNCH PLAN

74 Cameron Horton | Aerodynamics

Mass of Sample (kg)

Total Launches Launches per

Year

Launches per Colony

(per Year)

Total Volume (m3)

10 300 75 25 0.01

25 120 30 10 0.03

250 12 3 1 0.24

Pod Empty Mass (Mg)

Pod Final Mass (Mg)

Pod Volume

(m3)

Frontal Surface

Area (m2)

0.27 0.52 0.16 1.57

All renderings done in CATIA

RE-ENTRY

75 Cameron Horton | Aerodynamics

Original code by Nicholas LaPiana -20 -15 -10 -50

0.2

0.4

0.6

0.8

1

X: -8.1

Y: 0.5064

Heat Load, Heat Flux, and G vs. Entry Angle

entry angle gamma (degrees)

X: -10.2

Y: 0.4047

X: -6.5

Y: 1

Heat Load

Heat Flux

G

Atmosphere Deflection

Optimal Entry Angle

Ballistic Coefficient

Max Heat Rate (Watts/cm2)

Heat Load (Joule/cm2)

Max G Load (G)

-8.1o 256.27 672.95 21753.23 33.67

78

ERIK SLETTEHAUGH | STRUCTURES

PAYLOAD CONFIGURATION

o PAYLOAD CONFIGURATION

o COSTS FOR LAUNCH VEHICLES

2/13/2014

PAYLOAD CONFIGURATION

79 Erik Slettehaugh | Launch Vehicle & Com Sats

Assume: Payload to Lunar Surface w/ SLS, Mass: 12Mg; Volume: 479.5m^3

# of Launches Delta Mass Delta Volume

72 166.82 24,977.47

Total

Launch #Launch

IdentificationDescription

Launch

VehicleMass Volume Delta Mass Delta Volume

1#1 Communication

SatelliteSatellite at L1

Delta IV

(Small)1.77 3.44 7.42 339.07

2#2 Communication

Satellite

#1 Satellite in Polar

Obit, #2 Satellite in

Polar Obit

Delta IV

(Small)3.54 6.88 5.65 335.63

3#3 Communication

SatelliteSatellite at L2

Delta IV

(Small)1.77 3.44 7.42 339.07

4 #1 Pod Habitat

Sleeping

Area/Bathrooms

Eating/Meeting

Food Storage/Prep

Lab/Medical

SLS 8.671 301.44 3.33 178.05

5 #2 Pod Habitat Rec Room SLS 8.67 301.44 3.33 178.05

6 #3 Habitat Cargo

Initial Water, System &

Storage

ECLSS & Fire

Suppression

SLS 11.319 13.5 0.68 465.99

7 #4 Habitat Cargo

Lunar Fission Reactor,

Solar Panels,

RTGs, 1 Way Point

Battery, Laser

Communication

Terminal

SLS 11.928 50.286 0.07 429.20

8 #5 Habitat Cargo

Habitat Shield Support

Structure, 3D Printing

Machine, 1 Way Point

Battery

SLS 10.8 14.8 1.20 464.69

9 #6 Habitat Cargo 3 Way Point Batteries SLS 12 3 0.00 476.49

10 #1 Heavy RoverHeavy Rover - Tires,

Cab, etc.SLS 10.45 52 1.55 427.49

11 #2 Heavy Rover Battery - Heavy Rover SLS 10 1.3 2.00 478.19

12 #3 Heavy RoverHeavy Rover - Tires,

Cab, etc.SLS 10.45 52 1.55 427.49

13 #4 Heavy Rover Battery - Heavy Rover SLS 10 1.3 2.00 478.19

14 #1 Light Rover Light Rover SLS 14.4 40 2.40 439.49

15 #2 Light Rover Light Rover SLS 14.4 40 2.40 439.49

16 #3 Light Rover Light Rover SLS 14.4 40 2.40 439.49

17 #4 Light Rover Light Rover SLS 14.4 40 2.40 439.49

All Bases

Launch #Launch

IdentificationDescription

Launch

VehicleMass Volume Delta Mass Delta Volume

18#1 Construction

Vehicles

Soil Transport Vehicle,

1 Way Point BatterySLS 9.48118 88.799 2.52 390.69

19#2 Construction

VehiclesBulldozer SLS 14.848 36.5397 2.85 442.95

20#3 Construction

Vehicles3 Helper Robots SLS 12 81 0.00 398.49

21#1 Science Sample

Carrier

Sample Capsule, Rail

Gun, etc., 1 Way Point

Battery

SLS 8.81 5.61 3.19 473.88

22 #1 Crew Transport

Crew Transport

Vehicle, Crew Moon

Lander

SLS 80.78 640.6 48.95 859.51

23 #1 ResupplyFood, Water, Personal

Items, Spare PartsSLS 12 479 0.00 0.49

21*

#1 Interferometer

(Packed with #1

Science Sample

Carrier)

30 Radio Interferometer

Array CubesSLS 3 3.75 N/A N/A

11*

&

13*

# 1 Skylight Repel

System (1 Plate

Packed with #2 &

1 Plate w/ #2

Heavy Rover)

Skylight Repel System,

2 Elevator PlatesSLS 3.9375 2.25 N/A N/A

Shackleton Base

Skylight Base

COST PER LAUNCH

80 Erik Slettehaugh | Launch Vehicle & Com Sats

Assisted by: Finu Lukose, Sadie Holbert

Launch

Vehicle

Type Payload

Fairing

Payload

Volume

(m^3)

Mass (LEO)

(Mg)

Cost

($1M)/Mg

Cost

($1M)/(m^3)

Cost / Launch

($1M)

8.4m Diameter 1058.48 129.73 $ 3.85 $ 0.47 $500.00

10m Diameter 1500.11 129.73 $ 3.85 $ 0.33 $500.00

Falcon

HeavyComposite Fairing 278.21 53.00 $ 2.74 $ 0.52 $145.00

Dragon Spacecraft

& Trunk77.42 13.15 $ 4.30 $ 0.73 $56.50

Composite Fairing 278.21 13.15 $ 9.51 $ 0.45 $125.00

Large Payload

Fairing (LPF) 401280.47 9.80 $ 9.18 $ 0.32 $90.00

Extended Payload

Fairing (EPF) 411316.43 12.03 $ 7.90 $ 0.30 $95.00

Extra Extended

Payload Fairing

(XEPF) 431

356.41 15.26 $ 6.55 $ 0.28 $100.00

Short 401 485.26 9.80 $ 8.67 $ 0.18 $85.00

Medium 521 665.39 13.50 $ 7.04 $ 0.14 $95.00

Long 551 933.80 18.85 $ 6.90 $ 0.14 $130.00

Small Payload

Fairing342.51 9.19 $ 9.79 $ 0.26 $90.00

Medium Payload

Fairing583.34 11.06 $ 9.04 $ 0.17 $100.00

Large Payload

Fairing882.01 28.30 $ 4.95 $ 0.16 $140.00

SLS

Atlas V

Delta IV

Falcon 9

9m

9.8m

10m

9.738m

6.438m

19.1m

Cargo Transport

Vehicle

Payload Bay

3.6m Unused Space

Assisted by: Arika Armstrong

SLS Payload Fairing

84

ANDREW EMANS | STRUCTURES

HABITAT SHIELDING

o HABITAT SHIELDING SUPPORT STRUCTURE

o 3D PRINTING WITH REGOLITH

02/13/2014

HABITAT SHIELDING

85

Initial analysis shows a reduced launch mass with the use of a 3D lunar concrete printer instead of carbon fiber as the shield support material.

Pros/cons for concrete • Requires a 3D printer larger than any ever built which could fail

and leave the habitat unshielded • Reduces initial mass from Earth by ~30% compared to carbon

fiber (saves ~84% mass for additional check-points) • Printer can be reused for check-point construction

Pros/cons for carbon fiber • Easier to install • Reliable since more than one robot can complete the task of

installing

Andrew Emans | Structures

MATERIAL NUMBERS

86 Andrew Emans | Structures

Lunarcrete Total

Carbon Fiber

Mass [Mg] 8.2 11.7

Volume [m^3]

19.1 6.7

Power [kW] 60 0

Requires: 3D Printer Gantry Helper Robot

Construction Robot

Note: Mass and volume values were calculated by assuming thin walls and the primary mode of failure is due to buckling. Sketches drawn by Andrew Emans, except for the three arcs in the top sketch which were skillfully drawn by Arika Armstrong

90

ARIKA ARMSTRONG | STRUCTURES

CARGO SIZING

o CARGO POD

o ALLOWABLE PAYLOAD MASS

02/13/14

CARGO SIZING

91 Arika Armstrong | Structures

Cargo Pod

External Diameter [m] 9.800

External Height [m] 6.500

Internal Diameter [m] 9.738

Internal Height [m] 6.438

Mass [Mg] 4.741

Lander Vehicle

External Dimensions* [m] 9.8 D x 9 L

Inert Mass [Mg] 10.53

Max Landed Mass [Mg] 12.17

Max Cargo with Pod [Mg] 7.430

Propellant Mass [Mg] 103.0

ADDITIONAL CARGO SYSTEMS

92

Landing Strut System

• Based on Apollo – 3%

• ~ 0.8-1 Mg for landing gear

Marman Clamp

• ~0.1 Mg for a clamp

Arika Armstrong | Structures

93

FINU LUKOSE | PROPULSION

THRUSTER SIZING & CHASSIS DESIGN

o THRUSTER SIZING AND PROPELLANT MASS

o CHOICES FOR THRUSTERS

o PROGRESS ON CHASSIS DESIGN

2/13/2014

COM SAT PROPULSION SIZING Basic Forces Model on

cylindrical satellite

Refined model for greater maneuverability/pointing capability

Code estimating thrust necessary for maneuvers (Controls Team)

Sizing of thrusters based on thrust

• Off the shelf solutions

• Modified AAE 539 code

94 Finu Lukose | Propulsion

CHOICES FOR THRUSTERS

95

Controls sizing for micro-thrusters to resist various environmental torques (~5E-5N)

Approximate delta V for station keeping ~750 m/s for 4.5 year (based on LRO)

Finu Lukose | Propulsion

Bi-Propellant Thrusters

Model Thrust

(N) Propellant Isp (s)

System Mass (kg)

Propellant Mass (kg)

Total mass (kg)

Total Volume (m^3)

R-6D 22 MMH/NTO 294 0.454 297 307.896 0.349881818

R-1E 110 MMH/NTO 280 2 314 362 0.411363636

R-4D 490 MMH/NTO 316 3.4 274 355.6 0.404090909

Mono-Propellant Thrusters

Model Thrust

(N) Propellant Isp (s)

System Mass (kg)

Propellant Mass

Total mass (kg)

Total Volume (m^3)

MR-103D 1 Hydrazine 224 0.33 410 417.92 0.40972549

MR-111C 4 Hydrazine 229 0.33 399 406.92 0.398941176

MR-106E 22 Hydrazine 229 0.635 400 415.24 0.407098039

99

BRYAN FOSTER | PROPULSION

SCIENCE SAMPLE RETURN

o ELECTROMAGNETIC LAUNCHERS

o BIPROPELLANT ROCKET

2/13/2014

RAILGUN CONCEPT

100 Bryan Foster | Science Sample Return

Sample Mass (Mg)

Sample and Capsule Mass (Mg)

Force (kN)

Current (kA)

Voltage (V)

Energy (MJ)

Launches

0.250 0.5194 1592.43 1556.69 416.23 647.94 1

0.025 0.0722 221.36 653.39 138.72 114.15 10

0.010 0.03076 94.31 454.45 121.51 55.22 25

Ballistic Return Velocity 3065.9 m/s

Railgun Mass (Mg) Railgun Volume (m^3)

0.662 0.900

SAMPLE RETURN ROCKET

101 Bryan Foster | Science Sample Return

Sample Mass (Mg)

Sample and Capsule Mass (Mg)

Dry Mass of Rocket (Mg)

Propellant Mass (Mg)

Gross Mass of Rocket (Mg)

Volume of Rocket (m^3)

0.250 0.5194 0.65 1.06 1.71 13.18

Airbus Aestus Rocket Engine MMH/N2O4 propellant Uses same low energy ballistic

trajectory Releases samples and burns up in

reentry.

ERIC A FLORES / PROPULSION

CREW TRANSPORT VEHICLE

o CLM PROPELLANT ANALYSIS UPDATE

o CTV PROPELLANT ANALYSIS UPDATE

2/13/14

CLM PROPELLANT ANALYSIS UPDATE

112

Vehicle Propellant Mass

[Mg] Propellant Volume

[m^3]

LM Ascent 7.74 21.65

LM Descent 25 + 0.27 70.68

***Propellant Mass Credit to Krista Garret

Hovering Capabilities (Descent Stage) • 60 Seconds = 0.27 Mg of Propellant

o Assuming complete use of 25 Mg before hovering.

***Hovering Code Credit to Sean Snoke

CTV PROPELLANT ANALYSIS UPDATE

113

Trajectory Propellant Mass [Mg]

Propellant Volume [m^3]

LEO to LLO 58.40 169.89

LLO to LEO 12.17 35.40

Total 70.57 205.29 0

100

200

300

400

500

Selected Engines

Vo

lum

e (

m^3

)

Total CTV Propellant for the Mission

J-2 (Old Estimates)

J-2X (New Estimates)

50.7% Less Propellant

205.29

416.56

•Optimization Code Bringing Costs vs. Sending Costs?

117

ANDREW POWIS | POWER/THERMAL

ROVER POWER SYSTEMS

o VEHICLE TEAM LEADER: PRESSURIZED ROVERS

o FUEL CELLS TO BATTERIES

o ROVER SPECIFICATIONS

2/13/2014

TRANSITION FROM FUEL CELLS TO BATTERIES

Battery as shielding on Heavy Rover.

Decrease in final mass. Minimal change to launch mass.

Avoids the hydrogen storage issue.

Power system commonality between rovers and construction bots.

Easier to achieve maximum power output.

118 Andrew Powis | Power/Thermal

[3]

Eric Menke

Battery/ Shielding

Cabin

LIGHT ROVER / ROVER SUMMARY

119 Andrew Powis | Power/Thermal

Rover Heavy Light

Maximum Power (kW) 227.0 149.5

Total Energy (kWh) 1608 707.2

Battery Mass (Mg)[3] 10.87 4.779

Battery Volume (m3) [3] 4.717 2.074

Number of Battery Cells[3] 5973 2627

Rover Launch Mass (Mg) 18.79 10.00

Rover Launch Volume (m3) 31.37 26.51

Rover Final Mass (Mg) 45.55 10.00

Rover Final Volume (m3) 31.37 26.51

Jou

rney O

ut

Ret

urn

Jo

urn

ey

Habitat

Checkpoints

0 km

750 km

100 km

Ha Ni Kim

Light Rover Cabin Window

Life Support Systems

1.8 m

1.8 m

3.5 m

Standing Room

120

DIVINAA BURDER | POWER&THERMAL

CTV POWER SYSTEMS

o SOLAR ARRAYS

o LITHIUM-ION BATTERIES

2/13/2014

SOLAR ARRAYS

121

Stored in Service Module

Ultra-flex Butterfly wing shape (based on Orion)

Power: 15 kW

Volume: 30 m^3

Advantages:

• Low Mass

• Compact storage

• Durable

Divinaa Burder | Power & Thermal

5.5 m

LITHIUM-ION BATTERIES & SYSTEM DIAGRAM

122

Lithium Ion Batteries • Re-chargeable • Secondary power

– Dark Periods – Launch operations – Re-entry

• Power: 35 kWh • Mass: 400 kg (with insulation) • Volume: 0.18m^3

Power Systems Structure • Solar Arrays • Lithium-Ion Batteries • 2 Power Buses • 2 Power Control Units (PCUs)

Divinaa Burder | Power & Thermal

Solar Arrays

Battery

Battery

Battery

Battery

Power Bus

Power Bus

Secondary System

Primary System

PCU PCU

Sub Systems

Sub Systems

126

JOSEPH AVELLANO | PT/CARGO

CARGO VEHICLE THERMAL

o OPERATING THERMAL TEMPERATURES

o STORAGE OF ELECTRONIC SYSTEMS

o MULTI LAYER INSULATION

2/13/2014

OPERATIONAL TEMPERATURES

127

Low Earth Orbit Temps: • -75°C – 65°C

Operational Temp for Battery: • Lithium Ion (NCA) - 0°C

Total volume of systems : • .081 𝑚3

52 cm x 52 cm x 38 cm carbon fiber container • 1 cm thick

SIZING

128

Mass (kg) Power (W) Volume ()

Carbon Fiber Container

20.4 - .10275

Multi Layer Insulation

.48 - .14

Heating Pad .005 100 2.58E-6

Thermal Control Unit

.03 20 8.58E-6

• Carbon Fiber Container: 52 cm x 52 cm x 38 cm – 1 cm thick

• Multi Layer Insulation: 57.08 cm x 57.08 cm x 43.08 cm – 5.08 cm thick