PROJECT MANAGER TREVOR JAHN

LAB 4.7.2016

APM/SYSTEMS MIKE YOUNG

Final presentation slides

WATER PRODUCTION ARCHITECTURE COST SAVINGS ANALYSIS

Recommendation: After mission reaches steady state, stop shipping H2O and make all

H2O in-situ (less IMLEO and cheaper)

Limitations:

• Capacity of rover to supply 2 ISRU units

• Radiation contamination of ice deposits

• ISRU TRL low until later in mission (start-up time)

• Assume no need for additional power generation

Water production costs Shipped

(Mg) Shipped one-

time (Mg) Cost/yr % Savings

Ship all H2O but prop 49.9 16 $ 313,825,476.19 0% Make all but drinking in situ 17.9 52 $ 332,873,095.24 -6%

Make all H2O in situ 0 52 $ 247,619,047.62 21%

Using cost of ~$4.76M per Mg to lunar surface (SLS cost per launch/launch capacity)

MISSION TIMELINE

Washington Series (2022-2023, 4 missions)

Testing and XM2 Phase

• Raising TRL of power/orbiters

• XM2 Delivered to CLO

• Moon Impact Mission for habitat foundation

Adams Series (2023-2029, 12 missions)

First Construction Phase

• Deliver crew to orbiter for shakedown

• Validate life-support systems and resources

• Landing first five habs, ISRU equipment, consumables, rovers

• Crew construction mission 1

Jefferson Series (2029-2031, 5 missions)

Second Construction Phase

• Landing remaining four habs

• Crew construction mission 2

Madison Series (2032-2035, 11 missions)

Crew and Resupply

• Deliver personal items and consumables

• Crew delivery and rotation via Orion

capsule

Monroe Series (2035+)

• Crew to cycler rendezvous

• Crew rotation

4

MISSION DESIGN PAUL WITSBERGER

Final Presentation – Hyperbolic Rendezvous (Main slides and appendix)

04/06/2016

2035 MONROE 1: FERRY TO CYCLER

Exploration Upper Stage (EUS) • Places FEMAC into elliptical orbit

Booster • Performs hyperbolic insertion maneuver

Service Module • Identical ΔV capabilities as Booster • Only used if Booster fails

Crew Capsule • Supports a crew of 4 (payload of 22 Mg) • Heat shield capable of Earth reentry • Performs docking maneuver

Ferry to Mars Cycler (FEMAC) stage configuration:

37

Crew Capsule

2035 MONROE 1: FERRY TO CYCLER Maneuver breakdown:

1. EUS burn: Hohmann transfer to 1000 km altitude – 0.32 km/s

2. EUS burn: Transfer to high energy ellipse – 2.09 km/s

3. EUS jettison

4. FEMAC Booster burn: Inject into hyperbolic departure orbit –

1.69 km/s

5. FEMAC Booster burn: Trajectory correction maneuver – 0.066

km/s

6. FEMAC Booster and Service Module jettison

7. FEMAC RCS burn: Rendezvous/dock with cycler – 0.030 km/s

Total required ΔV – 4.194 km/s

Launch Vehicle

SLS Block 2

IMLEO 273.2 Mg

IVLEO 1171 m3

Total ΔV Capability

7.979 km/s

Time of Flight

46.11 hr

FEMAC Specifications

Abort Options • Safe return with controlled reentry: 3 hours after

insertion • Prevent escape: 34 hours after insertion

Crew survival rate 96.96%

Mission success rate 93.96%

78

APPENDIX: HYPERBOLIC RENDEZVOUS

• Previous designs were based on FEMAC-35, which requires the

lowest ΔV

• 2039 flyby requires the most ΔV – the vehicle must accommodate

future missions, so we are sizing the vehicle based on the 2039 flyby

Reasoning: A FEMAC (FErry to MArs Cycler) will dock with a Mars

cycler every two years, and each flyby is different

Year Vinf (km/s) rmin (km) ΔVtot (km/s)

2033 4.41 26700 4.4632

2035 3.75 9700 4.1941

2037 4.25 9000 4.2879

2039 5.53 23900 4.9473

69

MISSION DESIGN ALEXANDER BURTON

Final Presentation Slides

4/7/2016

2035 MONROE 1: FERRY TO CYCLER Maneuver breakdown: Total required delta-v is 4.194km/s

1. EUS burn: Hohmann transfer to 1000 km altitude – 0.32 km/s

2. EUS burn: Transfer to high energy ellipse – 2.09km/s

3. EUS jettison

4. Booster Burn: Inject into hyperbolic departure orbit – 1.69 km/s

5. Booster Burn: Trajectory correction maneuver – 0.066 km/s

6. Booster and Service Module jettison

7. Ferry RCS burn: Rendezvous/dock with cycler – 0.0300 km/s

*Booster burn can be also done by service module in emergency case

Launch Vehicle SLS Block 2

IMLEO 273.21 Mg

IVLEO 1171.4 m^3

Total Delta-V capability

7.979 km/s

Time of Flight 46.11 hr

Basic specification of the FEMAC Configuration allows us to achieve; • Crew Survival rate: 96.96% • Mission Success rate: 93.96% Abort Options • Safe return with a more controlled

reentry: 45 min. (2039) to 3 hours (2035) after insertion

• Prevent escape: 3 hours, 11 min. (2039) to 34 hours (2035) after insertion

APPENDIX: FERRYING LANDER MISSION PARAMETERS

Mission: • Lunar Surface CLO XM-2 • Drop off old astronauts, pick up new ones • XM-2 Lunar Surface

Event Time-of-Flight

Takeoff 6.7 min

Ascent Hohmann Transfer

4.5 hours

Descent Hohmann Transfer

4.5 hour

Landing 13 min.

Total 9.33 hours

Maneuver ΔV (km/s) Number Required

Takeoff 1.961 1

Circularizing Burton to Enter CLO

0.231 1

15o Plane Change 0.271 2

Descent Hohmann and Landing

2.459 1

Total 5.193

APPENDIX: FERRYING LANDER TAKE ASTRONAUTS TO AND FROM XM-2 IN CLO

3. Enter CLO ΔV = 0.231 km/s

1. Takeoff ΔV = 1.96 km/s

2. Hohmann Transfer to CLO

4. Up to 15o Plane change to CLO ΔV = 0.271 km/s

5. Rendezvous with XM-2

6. Return and Land ΔV = 2.73 km/s

Event Time-of-Flight

Takeoff 6.7 min

Ascent Hohmann Transfer

4.5 hours

Descent Hohmann Transfer

4.5 hour

Landing 13 min.

Total 9 hours, 20 min

SYSTEMS KYLE BUSH

Updated Slides for Final Presentation

4/7/16

IMPACTORS VS DIGGING What is needed?

• Need to find the most efficient method to move regolith for hab construction

• Need a construction schedule to reflect the recommended method

Assumptions:

• Time numbers are assuming rover is constantly moving regolith or charging

• Full list of assumptions is located on following slide

Rectangle Dig Only

Rectangle Impact

Aldrin Dig Only

Aldrin Impact

Mass (Mg) 3676.8 3468.4 3542.6 3634.3

Power Required (kW) 1857.0 1542.6 1789.2 1101.3

Volume (m^3) 2451.2 2312.3 2361.8 2422.9

Excavation Time (days) 291.36 222.20 283.91 176.46

Fill in Time (days) 117.18 117.18 109.72 65.833

Table 1: Method Comparison Table

Kyle Bush

IMPACTOR ASSUMPTIONS

The rover:

• The ramps giving access to the pits all have an incline of 16°

• Volume of rover scoop: 0.66 m^3

• Power for a low work trip: 300 W

• Power for a high work trip: 500 W

• Battery charge time: 28.8 hrs

• Battery Life: 10000 W and 24 hrs

The habs:

• 7.4 m diameter

• Short connectors are 0.5 m long

• Habs are buried 2.7m

The Moon

• Regolith density: 1500 kg/m^3

• Rectangle crater is 32 x 32 x 3.11 m initially with 4 m long sloping walls

• Aldrin crater has diameter of 20 m that slopes into a diameter of 5 m over a depth of 4.17 m

Figure 4: One of Three Pits for Dr. Aldrin’s Hab Layout

Figure 3: Pit for Rectangular Hab Layout

Kyle Bush

CONSTRUCTION SCHEDULE The Procedure

1. The impactors hit

2. Rovers fill in the craters to create

pits

3. Hab 1 lands and is positioned in the

pit closest to landing site

4. Rover fills in regolith around hab

5. First connector is positioned and

attached

6. Regolith bag wall is built between pit

and landing site

7. Repeat steps 3 – 5 for other two

habs and connectors in cluster

8. Attach cluster connectors to Hab 3

9. Repeat steps 3 – 8 for the other two

clusters

Figure 1: First Hab Cluster in Pit

Figure 2: Full Layout with Placement Order

Kyle Bush

PROPULSION ANDREW CULL

A6. XM Attitude Control System

B4. Nuclear Thermal Propulsion

C1. Hydrolox vs Methalox Trade Study

1 ANDREW CULL

A6. XM ATTITUDE CONTROL SYSTEM ENGINE & PROPELLANT SELECTION Assumptions

• ΔV = 100 m/s/year

• XM Module weight = 20 tons

• 20 year life span of BA330/XM

Requirements

• Ability to pulse

• Quick start up

• High ISP

• 6 DOF control

Propellant/Engine Selection

• 16x MR-107 (220 N)

• Monopropellant Hydrazine

• Catalyst S405/LCH-202

• ISP = 229 sec

• Refuel every 10 years

18 Andrew Cull

System Parameters – 20 years

Propellant Mass 12Mg/10 years

Inert Mass 20.01 Mg

Power 34.5 W/thruster

Propellant Volume 12 m3/10 years

B4. NUCLEAR THERMAL PROPULSION

19

Pros Cons

High specific Impulse > 800s Never been flight tested

Decreased time of flight High development costs

Broader launch window Potential for spreading nuclear

material if disaster occurs

Public fear of nuclear energy

Long shutdown times

Despite advantages, nuclear thermal rockets will not be available during our mission schedule. Thus, we chose to exclude them from our designs.

Andrew Cull

C1. HYDROLOX VERSUS METHALOX

20

Parameter Hydrolox Methalox

Specific impulse (s) 450 375-400

Boiling Point (K) 20 111

Manufacture Method Electrolysis Electrolysis and Sabatier

Ease of Storage Will boil in PSR Can be stored in PSR

TRL # 9 6

Hydrolox is the most useful propellant for our application as it decreases the IMLEO due to its low density and high specific impulse.

20 ANDREW CULL

Andrew Cull

HUMAN FACTORS KELLY KRAMER

4/7/2016

Final designs

HABITAT MODULES

Hab number Hab type Total mass,

Mg

Power,

kW

1L Living quarters 1 16.14 6.44

2L Living quarters 2 16.14 6.44

3R Recreation 14.10 4.64

4R Exercise 15.80 5.16

5W Waste/water management 18.91 13.15

6LL Laboratory/work station 16.58 9.12

7M Medical bay 16.99 9.79

8A Aeroponics 18.09 8.87

9F Food preparation/storage 19.25 12.64

Airlock

Kelly Kramer

STEADY STATE OPERATIONS 2032: First Crew to Lunar Surface 40.6 Mg water, 5.35 Mg packaged food already on base 2033: Resupply mission 3 Mg water, 2 Mg food 2034: Resupply mission 3 Mg water, 2 Mg food 2035: Resupply mission 3 Mg water, 2 Mg food 4.8 Mg water First crew to cycler

Week 64, Day 3

Time Time

08:00 Wake up 16:00

Rover excursion

08:30 Breakfast

16:30

09:00 17:00

09:30

Medical test

17:30

10:00 18:00

10:30 18:30

11:00 Wallyball game

19:00

11:30 19:30 Dinner

12:00 Rest/free time

20:00

12:30 20:30 Priority task

13:00 Lunch

21:00

13:30 21:30 Exercise program

14:00

Rest

22:00

14:30 22:30 Personal time

15:00 23:00

15:30 23:30 Sleep

CREW SELECTION AND MENTAL HEALTH Crew screening, selection, composition

• Height: 60-72 in.

• Blood Pressure: 140/90

• Visual acuity: 20/100

• Degrees: medical, engineering, science

Risk Mitigation and Support

• Autonomy for crewmembers

• Delegate tasks

• Goal-oriented, meaningful work

• Rotation of leadership

• Uplink of news, media, email

• Psychological conferences

• Private Earth contact

Sleep and Circadian Rhythms

• Strict work-rest schedule

• NASA developed LED light systems

• Rest during day, long sleep at night

SCIENCE ELLEN CZAPLINSKI

April 7, 2016

- Science rover overview

- Updated traverse maps

- Updated final presentation slides

25

SAMPLE RETURN BASE TO SCHRÖDINGER AND BACK ~ 1200 KM

Science Rover Details:

- Initial traverse to Schrodinger Basin (~ 600 km)

- Sampling impact melt sheet areas along the

way and at Schrodinger

- Rover will return these samples to base

- Samples analyzed and dated (if able) in lunar

lab

- Samples unable to be fully analyzed in lunar

lab sent to Earth with astronauts

- Further analysis done on Earth

- Other science rover traverses further into SPA

- Testing mineralogy of samples

- Ground-truth measurements

- All analysis performed in situ

26 Ellen Czaplinski

EXTENDED TRAVERSE TO SPA

- After sample return

from Schrödinger, rover

goes further into SPA

- Sampling exposed

bedrock in central peak

rings

- Mafic Mound may

contain remnants of

impact melt layer (Hurwitz and Kring, 2014)

- Rover will store

samples in cache

- Allows rover to travel on

alternate routes

27 Ellen Czaplinski

EXTENDED TRAVERSE TO SPA

28 Ellen Czaplinski

2023 WASHINGTON 4: TRAVERSE MAP

FOR SCIENCE ROVER

29

CABEUS CRATER TO SCHRÖDINGER BASIN = ~600 KM. TOTAL TRAVEL DISTANCE ~1200 KM

2023 WASHINGTON 4: SCIENCE TRACEABILITY MATRIX

Science Objective Justification Measurement Objective

Measurement Requirement

Instrument Selected

Constrain Bulk Composition of

the Moon

Constrain age of SPA and Late

Heavy Bombardment (LHB) theory

Sample return from SPA to

analyze mineralogy and

volatile distributions

Age SPA melt sheet within 20 My ppb level –

measure high FeO areas

Drill, Sample, NSS, SuperCam, hand

lens

Example of one row from the STM. Full STM encompasses 3 goals.

The South Pole-Aitken Basin (SPA) has high Iron Oxide levels (yellow) that we want to sample using our science rovers.

30

BACKUP

31 Ellen Czaplinski

UPDATED SCIENCE ROVER INSTRUMENT LIST

Instrument Mass (kg) Power (W) Volume (m3)

DAN 4.7 17.5 0.0045

NIRVSS Unknown Unknown Unknown

MastCam-Z 4.5 11.8 0.009

RAD 1.6 4.2 0.00024

ChemCam 5.778 Unknown 0.00133

APXS 0.37 Unknown 0.000368

CheMin 10 40 0.027

Mars Compass 0.5 2.5 Unknown

Ground penetrating

radar 4.53 Unknown 0.0063

Drill (5 cm) 4 Unknown Unknown

Dust Counter 1.6 5.1 Unknown

Total 37.578 81.1 0.048738

REFERENCES Duke, M. B., et al. Sample return mission to the South Pole-Aitken basin. Workshop on

New Views of the Moon II. #8017.

Hurwitz, D., Kring, D., (2014). Differentiatin of the South Pole-Aitken basin impact melt

sheet: Implications for lunar exploration. Journal of Geophysical Research:

Planets, 119, 1110-1133.

Hurwitz, D., Kring, D., (2015). Potential sample sites for South Pole-Aitken basin

impact melt within the Schrödinger basin. Earth and Planetary Science Letters,

427, 31-36.

32 Ellen Czaplinski

SLIDES FOR FINAL PRESENTATION

AUSTIN BLACK - STRUCTURES

Final CAD Renders of rover attachments and Fuel Depot, Finalized tables and values.

33

PROJECT LEGACY – LUNAR SURFACE

(ORIGINAL SLIDE #6) Permanently Shadowed Region

• In-Situ Resource Utilization – water extraction

• Fuel Depot

• Ferrying Lander

LOX

Tank

LH2

Tank

H2O Tank

Heat

Exchanger

MLI

Insulation

Aluminum

Support 1

5

4

3 2

34 – Austin Black

Permanently Shadowed Region

Fuel Depot

PROJECT LEGACY – LUNAR SURFACE

35 – Austin Black

Permanently Shadowed Region Fuel Depot

LOX

Tank

LH2

Tank

H2O Tank

Heat

Exchanger

MLI Insulation

Aluminum

Support

Permanently Shadowed Region

In-Situ Resource Utilization – water extraction

Fuel Depot

Ferrying Lander

Scoop: Volume: 0.66 [m3] Mass: 1.987 [Mg/arm]

2025 ADAMS 2: UNIFIED ROVER SYSTEM (ORIGINAL SLIDE #31)

Universal Pallets

JVA Bed

JVA-01: Volume: 14 [m3] Mass: 1.2 [Mg] Power: 390 [W]

36 – Austin Black

2025 ADAMS 2: UNIFIED ROVER SYSTEM

Universal Pallets

JVA Bed

37 – Austin Black

Scoop

Volume: 0.5 [m3/scoop]

Mass: 1.987 [Mg/arm]

JVA-01

Volume: 14 [m3]

Mass: 1.2 [Mg]

Power: 390 [W]

Extended configuration Collapsed configuration

REGOLITH BAGGING (ORIGINAL SLIDE

#34) Fill Process: 1) Roll bag

2) Catch opening

3) Fill

4) Cut perforations

5) Cinch

6) Drop

This system is one

of our rover pallets

which can moved

on/off our rover.

Bag Spool Fill Tube

Spring loaded hook

[1]

[3]

[6] [5]

[4]

38 – Austin Black

REGOLITH BAGGING

39 – Austin Black

Bag Spool

Fill Tube

Spring loaded hook

Regolith scoop

Bagging Process

1. Roll bag

2. Scoop regolith

3. Catch opening

4. Fill trough

5. Cut perforations

6. Cinch

7. Drop

This system is one of our rover pallets which can be moved on/off our rover.

2026 ADAMS 5: FUEL DEPOT

(ORIGINAL SLIDE #37)

• Location: PSR At mining site • Function: Processes water

into fuel/oxidizer • Mass (dry): 16.64Mg • Power: 300 Watts • Volume: 313.9m^3

LOX

Tank

LH2

Tank

H2O Tank

Heat

Exchanger

MLI

Insulation

Aluminum

Support 1

5

4

3 2

Tube # Service

1 Liquid Water to H2O

Tank

2 Gaseous Hydrogen to

LH2 Tank

3 Liquid Hydrogen output

line

4 Gaseous Oxygen to LOX

Tank

5 Liquid Oxygen output

line

40 – Austin Black

2026 ADAMS 5: FUEL DEPOT

Tube # Service

1 Liquid Water to H2O Tank

2 Gaseous Hydrogen to LH2 Tank

3 Liquid Hydrogen output line

4 Gaseous Oxygen to LOX Tank

5 Liquid Oxygen output line

41 – Austin Black

Fuel Depot

Location: PSR At mining site

Function: Processes water into fuel/oxidizer

Mass (dry): 16.64 [Mg]

Power: 300 [Watts]

Volume: 313.9 [m3]

LOX

Tank

LH2

Tank

H2O Tank

Heat

Exchanger

MLI Insulation

Aluminum

Support 1

5

4

3 2

H2O Input Line

LOX/LH2

Output Lines

PROPULSION MATT SCHURMAN

Possible Presentation Slide

4/7/2016

42

COMSAT SPECIFICATIONS

Matt Schurman

43

XIPS 25 cm

Mass of Propellant (Xenon) 282 kg

Power Usage (kW) 4.3 kW

Isp 3550 s

Thrust 0.165 N

#Engines 1

Main Engine

MR-111

Mass of Propellant (Hydrazine) 146 kg

Isp 229 s

#Engine 1

Circularization Engine

Total Mass 1200 kg

Dry Mass 722 kg

General Comsat Characteristics

HUMAN FACTORS RACHAEL HESS

Final Hab Images

HABS 1-5

45

Rachael Hess

Hab 1L (Living) Residential Areas

Hab 2L (Living) Residential Areas

Hab 3R (Rec Center) Wally ball/Multipurpose Court

Hab 4R (Rec Center) Exercise Equipment

Hab 5W (Waste/Water) Water Reclamation

HABS 6-9

46

Rachael Hess

Hab 6LL (Laboratory) Laboratory Equipment

Hab 7M (Medical) Medical Bay

Hab 8A (Aeroponics) Food Production

Hab 9F (Food Storage) Food Storage and Preparation

HAB 1L & 2L: LIVING QUARTERS Rachael Hess

HAB 3R: WALLYBALL/MULTIPURPOSE

COURT

Rachael Hess

HAB 4R: RECREATION CENTER Rachael Hess

HAB 5W: WASTE & WATER RECLAMATION Rachael Hess

HAB 6LL: LABORATORY STATION Rachael Hess

HAB 7M: MEDICAL BAY Rachael Hess

HAB 8A: AEROPONICS Rachael Hess

HAB 9F: FOOD STORAGE Rachael Hess

PROPULSION BROCK MILLER

Slide Recommendations for the Final Report

2023 WASHINGTON 4: CARGO LANDER - For the 20 Mg lander, the total DeltaV to land from our 4500 km orbiting radius will be 2.5

km/s.

- Both landers are powered by 1 Aerojet Rocketdyne RL10B-2 Engine.

- 20 Mg Lander is 10.5 meters tall 7 meters wide at the top and 14 meters wide at the struts

Variant LHy Vol. (m3)

LHy Mass (Mg)

LOx. Vol. (m3)

LOx. Mass (Mg)

5 Mg 7.43 0.53 2.69 3.09

20 Mg 31.0 2.20 11.23 12.91

Hab and Lander in SLS Fairing

Cargo Lander with Extended Struts

Brock Miller

APPENDIX SLIDE METHALOX TRADE STUDY FOR CARGO LANDER

[Table values are based on a 1 to 5 scale, 5 being the best]

Startability – Both Require ignition systems

Refueling – For now, the cargo landers are not planned to be refueled

Efficiency – Hydrolox have Isp values of about 450 s, Methalox is about 390 s

Storage – More volume will be needed for Hydrolox (less dense fuel)

TRL – Hydrolox is highly proven, having flown many missions as opposed to Methalox which has a TLR of 3-4

Fuel Type Startability Refuelling Efficiency Storage TRL Totals

Hydrolox 2.5 2.5 5 2 5 3.75

Methalox 2.5 5 3 4 1.5 2.975

Weighting 0.15 0.05 0.35 0.25 0.2

Brock Miller

POWER AND THERMAL RACHEL LUCAS

April 7, 2016

Reactor Summary

Possible Nuclear Risks

SAFE-400 SUMMARY

Rachel Lucas

Total Mass: 1.08 Mg Total Power: 200 kWe, 800 kWt,

Total Volume: 0.24 m3

Working Fluid: Sodium Distance from Base: 1.1 km

Fuel: Uranium Nitrate Lifetime: 5-7 years

Mission

Name

Year Description

Washington 3 2023 Raise TRL from 7 to 8

Washington 4 2023 Raise TRL from 8 to 9

Adams 2 2025 Delivery of two nuclear

reactors for base usage

Adams 11 2028 Delivery of one replacement

reactor for the base

Interior of SAFE-400 Reactor

POSSIBLE NUCLEAR RISKS

Nuclear Meltdown

• Occurs when a reactor is improperly cooled

• Constant supervision of the reactors is recommended as they haven’t

been tested in a lunar environment previously

• Regular checks of reactor components should be made

Radiation Shielding

• Reactor produces enough radiation that it could be dangerous and

possibly fatal to nearby inhabitants

• It is therefore recommended that the reactor be a distance of at least 1.1

km from the base

Orbital Failure

• Danger lies in the possibility of nuclear fuel being dispersed throughout

a planet’s atmosphere

Rachel Lucas

REFERENCES

[1] Poston, D., Kapernick, R., Guffee, R., “Design and Analysis of the SAFE-400 Space

Fission Reactor,” AIP Conference Proceedings, 608, 578, 2002.

[2] El-Genk, M., Tournier, J., “Performance Analysis of Potassium Heat Pipes Radiator for

HP-STMCs Spac Reactor Power System,” ResearchGate, 2004.

[3] Tournier, J., El-Genk, M., “Radiator Heat Pipes with Carbon-Carbon Fins and Armor

for Space Nuclear Reactor Power Systems,” Space Technology and Applications

International Forum, 2005.

SYSTEMS OF STRUCTURES BENJAMIN MISHLER

State of the Industry Future Technologies

Benjamin Mishler

PERTINENT TECHNOLOGY DATES

Benjamin Mishler

Technology/Event When it’s supposed to be ready

Bigelow BEAM Test (inflatable structure)

Scheduled for a 4/8/2016 launch on CRS-8

Results from Scott Kelly’s (almost) year in space

6 months+

SpaceX Falcon Heavy Later this year

SpaceX Mars Colonial Transporter (using Raptor Engines)

Mid 2020’s

POWER AND THERMAL WERONIKA JUSZCZAK

Thermal Control System Overview

THERMAL CONTROL SYSTEM

Weronika Juszczak

R11

Ammonia Q

Heat Pump, Rankine Cycle

W

MOON

RADIATOR ARRAY

HABS

COMPRESSOR

HEAT EXCHANGER

Active Thermal Control System

Passive Thermal Control System

Insulation Layers Thickness (cm)

MLI 10 2

TOTAL MASS: 1.471 Mg

THERMAL SYSTEM OVERVIEW

Weronika Juszczak

Project Legacy • Insulation

• MLI (aluminized Mylar) • Power Requirement: 154 kW • Coolant: Ammonia, R11

• 1 acquisition coolant loop • R11 outer loop (greater

pumping power) • Radiated Heat: 250 kW/m2

• Radiator • 200 m2

• Fixed, vertically oriented • Thermal louvers

ISS • Insulation

• MLI (aluminized Mylar) • Power Requirement: 75-90 kW • Coolant: Ammonia

• 2 loops (two-temperature loops) • Radiated Heat: 275 kW/m2

• Radiator • 154 m2

• Rotate to reject maximum heat

BACKUP SLIDES

Weronika Juszczak

System Mass (Mg)

Compressor 0.01645

Evaporator 0.247

Pipes and Fluid 0.174

Radiator Array 0.5001

Heat Exchanger (Cold Plates, Fluid, Piping)

0.432

Properties Mass (Mg)

Fin Efficiency 0.8

Rejection Loop T [K]

362

Coolant Loop T [K] 275

BACKUP SLIDES

Weronika Juszczak

Figures: NASA, ATCS

REFERENCES [1] NASA, 1992, Thermal Control Systems for Low-Temperature Heat Rejection, http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19920011027.pdf [2] https://www.nasa.gov/pdf/473486main_iss_atcs_overview.pdf

Weronika Juszczak

CONTROLS BECCA PIETRZYCKI

Final Presentation Slides:

Satellite Parameters

Vehicle Controls

IMLEO

BACKUP: SATELLITE PARAMETERS Table X provides the mass, power, and size off all antennas in the communication scheme for the

whole mission.

Vehicle/Location of Antenna Mass (Mg) Power (W) Diameter (m)

Earth Bases (3) 4.70 100 9.40

Moon Base 1.50 25 3.00

Comm. Sats to Moon (3) 0.047 18 0.044

Comm. Sats. to Earth (3) 0.047 55 1.31

Ferrying Lander 0.003 25 0.131

Cargo Lander 0.003 8.0 0/188

Ferry-Cycler 0.003 65 1.220

Pressurized Rover (X-Band) 0.001 2.0 0.769

Pressurized Rover (HGA) 0.001 1.0 1.31

Science Probes 0.047 2.0 0.769

XM-2 0.047 50 0.088

Table X: Parameters for all satellites in the communication

scheme for Project Legacy IMLEO: 1.887 Mg

BACKUP: VEHICLE CONTROLS Table X provides the mass, power, and volume of the full control system on each

vehicle. These vehicles must be controlled due to the forces listed below.

Vehicle Control Method Mass (Mg) Power (W)

XM CMGs & Thrusters 0.826 200 1.00

Ferry to Cycler CMGs 0.544 552 3.24

Ferrying Lander Reaction Wheels &

CMGs

1.108 356 1.63

Cargo Lander Reaction wheels &

CMGs

0.241 1015 1.63

Comm. Satellites Reaction Wheels &

Thrusters

0.296 430 0.047

Table X: Parameters for all control systems for each vehicle in Project Legacy

Environmental forces we are concerned with:

•Gravitational forces

•Reflected solar radiation

•Solar radiation

•Gravity gradient

•Magnetic field force

IMLEO: 3.015 Mg

CONTROLS ZARIN BARI

Final Presentation Slide

Communication Satellite

Requirement: 24-hour HD video communication with crew throughout mission

Communication Satellite Specifications:

• Number of satellites: 3

• Orbit altitude: 1,200 km

• Period: 4 hours

• Inclination: 88 degrees (polar orbit)

• Relays data from Earth to Moon or from Moon to Earth

Control Characteristics: - 3 axis stabilization - Using 3 reaction wheels - Desaturate about once every 6 months

2022 WASHINGTON 2: COM SATS

Figure X.x. Front side view Communication satellite

Figure Y.y. Orbit of communication satellite

EXTRA INFORMATION

Figure Z.z. Backside view of communication satellite

CONTROLS CHAD OETTING

Updated Communications Map

COMMUNICATIONS MAP EARTH - MOON

Chad Oetting

COMMUNICATIONS MAP LUNAR

Chad Oetting

SYSTEMS MASON BUCKMAN

Similarity Between the Moon and Mars – Final Trade Study

ESI AND MSI

0.0 – 0.2 Completely Dissimilar

0.2 – 0.4 Dissimilar

0.4 – 0.6 Somewhat Dissimilar

0.6 – 0.8 Somewhat Similar

0.8 – 1.00 Very Similar

RESULTS AND RECOMMENDATION

Results: Interestingly enough, Antarctica is considered the most “Mars-like” based off of the MSI calculations. The Moon would also still be considered somewhat “Mars-like.” In fact, some past research for future Mars missions has been done in Antarctica.

Future Considerations: Early testing for key systems (ISRU, habs, rovers, etc.) could be done in Antarctica to gain at least basic functionality or data on how they need to operate. Doing so could potentially save millions or billions of dollars in testing and would eliminate many unknowns when being used on the Moon or Mars.

Location in our Solar System ESI

Earth (Average Conditions) 1.0000

Venus (High Atmosphere) 0.9712

Antarctica 0.8473

Mars 0.6975

Moon 0.5606

Venus (Surface) 0.4398

Location in our Solar System MSI

Mars 1.0000

Antarctica 0.7080

Moon 0.6892

Venus (High Atmosphere) 0.6451

Earth (Average Conditions) 0.6309

Venus (Surface) 0.3557

BACKUP SLIDES The weight parameters that are used to calculate ESI are obtained by using the definitions

of terrestrial planets. The lower and upper bounds of these definitions are put into the ESI

equation and equated to 0.8 which is the boundary for “like-ness.” The equation is then

solved for w to obtain the weight factor for each boundary. The two boundary values are

averaged to obtain the official weight exponent.

For example, terrestrial planet definition for radius is between 0.1 and 10 Earth radii.

These two values help to define the upper and lower boundary values which are averaged

to obtain the weight exponent. An exponent with a higher value has a greater effect on the

similarity index. Each parameter has its own unique weight exponent.

Since the weight exponents are created based on parameters written in Earth units, new

weight exponents were made using Mars units to create the Mars similarity index. ESI is

commonly used as a way to gauge potential habitability because the temperature

exponent for ESI is calculated from a range of 0°C to 50°, the most suitable temperature

range for life as we know it.

SYSTEMS ANAIS ARNAIZ

Regolith Bagging Considerations

April 7, 2016

REGOLITH BAG OPTIONS

Anais Arnaiz

Criteria Considered When Choosing Bags: 1. Stability 2. Volume 3. Mass 4. Movability of bags to final location 5. Loss of regolith in bags over time 6. Failure Consequences of bags

4 Bag Options 1. Bag with clinched top (Tear Drop) 2. Large rectangular shaped bags 3. Large sandbag 4. One large tarp that would contain all the regolith

REGOLITH BAGGING CONSIDERATIONS

Anais Arnaiz

Criteria TearDrop LargeRectangular

LargishSandbags

OneBigTarp

Stability Un-proportional Cubes Flat Stable

Volume 0.1m3to1m3 1m3to10m3

0.1m3to0.6m3 All

Mass <1kg <10kg <1kg All

Criteria TearDrop LargeRectangular

LargishSandbags

OneBigTarp

Movability 2 4 1 5

Lossofregolith 3 5 3 0FailureConsequence 3 1 2 5

1 (Good) to 5 (Bad)

BACK UP – OTHER CONSIDERATIONS

Old design: • Relied on 2 rovers • Bags laid flat • Wouldn’t be able to open bags

on lunar surface

New Design: • Uses only one rover since scoop is

attached to this pallet. • Bags are in rolls • Considers criteria for bag type used

CAD by Adrian and Jay

CAD by Anais Arnaiz