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PROJECT MANAGER TREVOR JAHN THURSDAY LAB 2/18/2016

Semester Schedule and Expectations

SEMESTER TIMELINE Project Legacy – Semester Schedule Subject to Change

Week

: 6

Feb 14 – 20 (2.16.2016) Three Copies of 1 Page Resume in lecture

(2.16.2016) First Peer Evaluation

(2.18.2016) Action items are assigned in lecture

(2.20.2016) 10:00 pm first five pages are due for the final

report DRAFT to PM via email Week

: 7

Feb 21 – 27 (2.23.2016) Three copies of 1 Page Resume in lecture

(2.25.2016) Action items from week 6 are resolved

(2.25.2016) After lecture Design Freeze is in effect Week

: 8

Feb 28 –

March 5 (2.29.2016) Preliminary Design Review (PDR)

(3.1.2016) Action items assigned as a result of PDR

(3.1.2016) Three copies of Long Resume due on Tuesday

(3.5.2016) 10:00 pm second five pages of the Final Report

DRAFT to PM via email (ten pages total) with the first five

pages and any revisions included Week

: 9

March 6 –

12 (3.8.2016) 3 Copies of Long Resume due on Tuesday

(3.10.2016) Action items resolved and presented in lab

(3.10.2016) After lecture Design Freeze in effect

(3.11.2016) 10:00 pm third set of five pages of the Final Report

DRAFT to PM via email (fifteen pages total) with the first ten

pages and any revisions included

SEMESTER TIMELINE Week:

10

March 13

– 19 Spring Break

Week:

11

March 20

– 26 (3.22/24.2016) Report writing exercises in class

(3.22.2016) Second Peer Evaluation

(3.25.2016) Critical Design Review (CDR) Week:

12

March 27

– April 2 (3.31.2016) PM and APM present to AAE Industrial Advisory Council

(Tentatively 9:30 am)

(3.31.2016) Report groups are assigned to finish up report topics Week:

13

April 3 –

9 (4.4.2016) Final report due to PM via email for assembly into near final

draft

(4.5.2016) Go over near final draft of the final report for review

(4.7.2016) Final report due to Professor Longuski and Professor Minton

(4.7.2016) Mike Griffin visit (lunch and afternoon class visit) Week:

14

April 10 –

16 (4.14.2016) PM and APM give dry run of final presentation 8:30 am –

11:20 am Week:

15

April 17 –

23 (4.19.2016) Website and video are due

^more info on this in the coming weeks (as of 2.16.2016)

(4.21.2016) Final Formal Presentation is given by PM and APM –

Stewart Room 206 from 8:00 am – 12:30 pm

(4.21.2016) CLASS ENDS FOR THE SEMESTER

SYSTEMS ALEXANDRA DUKES

Deliverable Calendar Formulation

Deliverable Framework

Final Report Structure

PDR Structure and Expectations

FINAL REPORT FRAMEWORK OVERALL EXPECTATIONS AND STRUCTURE – ALEXANDRA DUKES

Report Structure:

Executive Summary

State of the Industry

• Brief summary of current technologies and industry plans

Requirements and Assumptions

• Describes basis of project and assumptions

Mission Overview

• Mission architecture (“big picture” design overview)

Project Elements

• What, why, how of our design decisions

Project Considerations

• Interaction of elements

• Risk Assessment

Conclusion and Recommendations

• Reiteration of overall design and “big picture” justifications

• Improvements on finalized design

• Looking forward

• Includes technology tests for future Mars missions

Appendices

• Trade study outcomes

• In depth calculations including code

Report Requirements:

• 1000 pages

• 25 pages per person

• Project elements will vary in the amount of page numbers

• This does not mean extra work for certain people

• Editing and review will be a team effort

• Report writing exercises in Week 12

• Final report review from Week 12 to Week 13

PRELIMINARY DESIGN REVIEW (PDR) EXPECTATIONS, STRUCTURE, AUDIENCE, ETC. – ALEXANDRA DUKES

Structure:

• Brief Overview of the Final Report

• i.e. Will cover from “State of the Industry”

to “Trade Study Outcomes” in the

Appendix

• Slides will be organized by vehicle

groups, NOT mission groups

• Back up slides should contain a brief

summary of design justifications

• Project Elements will require CAD for

major structures

• Back up slides should contain CADs

supporting the main presentation

Schedule:

• PM will provide room and times (expected

to be scheduled for two hours)

• PM and APM will go through presentation

once without interruption

• Each person will have a complete copy of

the presentation for notes

• At the end, we will go through each slide

(maximum time of 2 min) and discuss

Individual Deliverables:

• One slide per person

• Include a summary of what should be said in “Click to add notes”

Purpose:

• Draft of the Final Presentation

• PDR will not require specifics, these will need to be defined for CDR

• Slides need to answer why the specific designs were chosen

• PDR audience will consist of only the class; CDR will consist of outside reviewers

BACKUP - FINAL REPORT FRAMEWORK DETAILS OF EACH SECTION (FOR FUTURE REFERENCE) – ALEXANDRA DUKES Executive Summary

The executive summary will act as the “foreword” of our project and provide the purpose of this project. This is intended to be able to be read

in lieu of the report and should stand as a separate document. It will summarize the design decisions, justifications and recommendations as

outlined within the report.

State of the Industry

The “State of Industry” section will provide a short and concise overview of the current activities within the space industry focusing on those

which have provided useful assumptions leading to design decisions within our project.

Requirements and Assumptions

This sections will summarize the basis of what we built our project from including the project specifications document and Dr. Aldrin’s initial

presentation. Additionally this section will provide the assumptions derived from the “State of Industry” section. An example is the necessity

to use Cape Canaveral as our sole launch site due to the launch constraints of international sites.

Mission Overview

The mission overview will provide the reader with our mission architecture which consists of the mission timeline. The mission timeline

includes the launch schedule, phases of the overall mission and purpose of each phase. This will provide the reader a basic understanding

of our design concepts before diving into the project element details.

Project Elements

This section provides a detailed description and justification for the final designs. This excludes any large calculations and codes. These

sections should consist of written descriptions of each element (XMs, Habs, Radiation Equipment, etc.) which will provide the reader and

understanding of each vehicle, the purpose of the vehicle and justifications for major design decisions. The appendix should contain your

detailed analysis of each element and can be referenced in this section.

Project Considerations

The project considerations brings all of the different project elements into a comprehensible vision for the reader. The section will include

how each of the elements interact with each other and the risk assessment of the overall project.

Conclusion and Recommendations

The conclusion will reiterate the overall design and major design justifications. At this point we have thrown nearly 1000 pages at the reader

and they may need to refresh on the mission overview and overall purpose of our design project. Within this section, we will create a “looking

forward” plan which includes the technologies which will be tested on the Moon for future Mars missions. Additionally we will include a

section of what should be done at the conclusion of these designs and improvements that could be made in the future.

Appendices

The appendices will contain all trade studies discussed throughout the semester which affected our design decisions but were not necessary

to mention within the “Project Elements” sections (i.e. why we have chosen not to include an asteroid mission, etc.) and all detailed analysis

and calculations behind out design project.

SYSTEMS KATHRYN O’CONNOR

Refined Mission Timeline, IMLEO Calculations/Updating, Project Assumptions

Thus Far

2/18/2016

8

MAJOR PROJECT LEGACY ASSUMPTIONS TOP TEN ASSUMPTIONS TO DESIGN FOR, AS DECIDED BY THE TEAM

1. Launching from Cape Canaveral with SLS, Atlas V (Falcon Heavy)

2. EUS sole use on SLS (Electric Xenon System sole use on Falcon

Heavy)

3. 8 people always on base after 1st two years; rotate crew of 4 every

2 years*

4. Launches from Moon Surface only done by Ferrying Lander

5. 9 upright cylinder Habs- 3 floors, 1st floor buried

6. Cargo landers will not be reusable

7. ISRU used solely for fuel, miscellaneous water use- drinking

water launched from Earth

8. ISRU in a PSR (or location that requires least amount of power)

9. Aeroponics will be used for 2 meals- supplemented by MREs from

Earth

10. Launching to Cycler from Earth – New crew to Mars

*unless medical tests decide unfit to stay after 2 years Kathryn O’Connor

9

REASONING BEHIND ASSUMPTIONS EXAMPLES:

Kathryn O’Connor

EUS sole use on SLS:

• Already equipped with 105 Mg LH2/LOX Tank

• Capable of putting 24 Mg/Cargo on Lunar surface (not including lander)

• Would need to put a fairing inside of a fairing – detrimental loss of volume

• Engine inside of fairing – detrimental loss of volume

• Falcon Heavy could implement other systems until full development of Raptor

Aeroponics will be used for 2 meals:

ISRU solely for fuel, miscellaneous water needs:

Sending All Food All Aeroponics 1 Meal Aeroponics 2 Meals Aeroponics

5 Mg/year 96 Mg 6 Mg, 2.5 Mg/year 14 Mg, 1.25 Mg/year

- 5.25 Mg/year 1 Mg/year 2.85 Mg/year

Methane* LOX* H2O* ISRU Output H2O Payload

4 Mg/trip 16 Mg/trip 10 Mg/year 50 Mg/year 5.5 Mg/year

*Assuming 95% recycled water, 1 emergency trip storage

10

FULL ASSUMPTIONS LIST Assumptions for Presentation 3:

1. The habs will be upright cylinders, 3 floors, with the first floor buried for structural support and radiation protection. There will be 9 of them with a configuration that has yet to be decided

2. The ISRU will be in a PSR such that the volatiles and fuel can be stored at a low temperature without having to use much power to maintain.

3. ISRU will also mainly be used solely for making fuel for the ferrying lander, as well as miscellaneous water use (showering, etc). Drinking water will be launched from earth.

4. The only place we are launching from is Cape Canaveral, with an SLS, Atlas V, and Falcon Heavy.

5. The rover will be a modular pressurized rover instead of 3 separate designs.

6. If we are using new people to send to Mars, it is more efficient to launch from Earth, thus we will be launching to the cycler from Earth.

7. There will be 8 people on the base at all times. The crew will rotate our in groups of 4 every 2 years, unless after the first 2 years the original crew is seen as unfit to stay, the crews will then be rotated out every 2 years until it is deemed safe to extend the stay.

8. Aeroponics will be used as a proof of concept for Mars and will only be used as a one meal supplement each day, thus two meals will be launched from earth for each person.

9. The cargo landers will not be reusable.

10. The only launches off the surface of the moon will be from the ferrying landers to the XM modules for return to Earth of the astronauts.

11. The upper stage propulsion for the launch vehicles used will solely be the exploration upper stage for the SLS. If using the Falcon Heavy, the upper stage will solely be an electric system using Xenon.

Kathryn O’Connor 11

CURRENT MISSION TIMELINE

Kathryn O’Connor

Launch Launch Vehicle Year Mission Name Description

Destination

1 Falcon Heavy 2018 Washington 1 XM 1 (+controls), Nuclear Reactor Test 1 LEO 2 Falcon Heavy 2019 Washington 2 XM 2 (+controls), Orion Docking Mechanism, 3 com sats, Surface Com, Nuclear Reactor Test 2 CLO 3 Atlas V 551 2020 Washington 3 1 Mg lander, Science Rover, Nuclear Reactor Test 3 CLO 4 SLS-Crewed 2021 Adams 1 2 Crewed Test to XM, docking test, and back (1 week duration) LS 5 SLS 2022 Adams 2 Hab, Modular Rover, Test ISRU LS

6 SLS 2022 Adams 3 Rec Center, Pressurization/Oxygen Test Machine, Urine/Water samples, Urine/Water Recycling Test

Machine, Rec Center Leftovers LS 7 SLS 2022 Adams 4 ISRU, Garage, Medical Hab, 3-D Printer LS 8 SLS 2023 Adams 5 4 Nuclear Reactors, (1/4) Solar Panels LS 9 SLS 2023 Adams 6 Fuel Depot, (2/4) Solar Panels, Walls LS

10 SLS 2023 Adams 7 (3) Connectors, (6) Doors, Pumps etc, (1) Walls LS 11 SLS 2023 Adams 8 Water Storage Tank, ISRU Storage, Power Generation/Storage/Grid LS 12 SLS 2023 Adams 9 (1/2) Ferry Landers, (1/2) Fuel for Landers LS 13 SLS-Crewed 2023 Jefferson 1 2 Crew to XM, 2 Crew to Surface (1 week): Test Systems, Science Probes CLO/LS 14 SLS 2024 Jefferson 2 Food Storage Area, Supplies, Med Supplies LS 15 SLS 2024 Jefferson 3 Hab Supplies, Garage Supplies, Oxygen/Pressurization Mech. LS 16 SLS 2024 Jefferson 4 Water Recycling Mech+Pumps, Garage Pressurization, Trash Compactor+Storage LS 17 SLS 2025 Jefferson 5 Airlocks, Spacesuits, Laboratory, Laboratory Supplies, (1) Connector, (2) Airtight Doors LS

18 SLS 2025 Jefferson 6 (1/4) Solar Panels, Rec Supplies, Hydroponics/Irrigation Equipment, XM Module Supplies, (1/4)

Oxygen Tanks LS 19 SLS 2025 Jefferson 7 Food, Water, Walls, Leftover science things LS 20 SLS - Crewed 2025 Jefferson 8 Personal Items, (4) Humans, Orion Capsule and Service Module CLO/LS 21 SLS 2026 Jefferson 9 Hab, Hab Supplies, (1) Connector, (2) Doors, Food, Water LS 22 SLS 2026 Jefferson 10 (2/2) Ferry Landers, (2/2) Fuel for Landers CLO 23 SLS - Crewed 2027 Jefferson 13 Personal Items, (4) Humans, Orion Capsule and Service Module CLO/LS 24 SLS 2027 Jefferson 14 Water LS 25 SLS - Crewed 2029 Madison 1 (3/4) Personal Items, (4) Humans, Orion Capsule and Service Module CLO/LS 26 SLS 2029 Madison 2 Food, Water LS 27 SLS - Crewed 2031 Madison 5 Personal Items, (4) Humans, Orion Capsule and Service Module CLO/LS 28 SLS 2031 Madison 6 Water LS 29 SLS - Crewed 2033 Madison 9 Personal Items, (4) Humans, Orion Capsule and Service Module CLO/LS 30 SLS 2033 Madison 10 Food, Water LS 31 SLS - Crewed 2035 Madison 11 Personal Items, (4) Humans, Orion Capsule and Service Module CLO/LS 32 SLS-Crewed 2035 Monroe 1 1st Crew to Cycler Cycler 12

APPROXIMATE IMLEO THIS WEEK Payload

(Mg) SLS Lander

Mass Falcon 9 Lander

Mass # of SLS

Launches # of Falcon 9

Launches SLS Fuel

Mass Falcon 9 Fuel

Mass Cargo

Current 334.667 28.48 28.48 17 - 105 -

Cargo Ideal 334.667 28.48 28.48 6 17 105 6

Cargo Goal 334.667 28.48 28.48 9 - 105 -

Crew 0.32 - - 7 - 105 -

SLS Inert Mass

Falcon 9 Inert Mass

SLS IMLEO (Mg)

Falcon 9 IMLEO (Mg)

SLS Cost ($Bil)

Falcon 9 Cost ($Bil)

Total IMLEO (Mg)

Total Cost ($Bil)

Cargo Current 10.3937 - 2780.5199 - 8.5 - 3698.6959 12

Cargo Ideal 10.3937 6 1539.6692 1022.827 3 1.6575 2457.8452 6.5

Cargo Goal 10.3937 - 1917.0314 - 4.5 - 2835.2074 8

Crew 25.848 - 918.176 - 3.5 - - -

Kathryn O’Connor 13

SYSTEMS NICHOLAS HOBAR

ISRU Subsystem Mapping

Tentative Mass Requirements

Nicholas Hobar

OVERALL SUBSYSTEM MAP Objective: To create an overview of the necessary functions of the ISRU

Reasoning: For groups to gain an understanding of how the ISRU will function

Nicholas Hobar

BOTTOM LINE NEEDED PRODUCTION RATES + MASSES

Total Annual Volatile Mass Output Needed: 216.5 Mg + 312 Mg for shielding

Mass of Regolith to be Processed: 8018.6 Mg + 11555.6 Mg for shielding

Power Required Annually: 4.22E8 Joules

Recommendation: Do not rely on the ISRU for water output for radiation shielding. Will

need to discuss with ISRU group to determine maximum possible output of system.

Necessary Annual Mass Output

• Fuel Components: 210.21 Mg

• H2O for Humans: 2095 kg annually + 2 years (4190 kg) of reserve

Mass of H2O for Radiation Shielding: Maximum of 312 Mg (Not annually)

Nicholas Hobar

BACKUP SLIDE BREAK DOWN OF OTHER PROCESSES: H2O EXTRACTION

BACKUP SLIDE ELECTROLYSIS

BACKUP SLIDE SABATIER PROCESS

BACKUP SLIDE BREAK DOWN OF MASS NUMBERS

- Assuming the use of Methalox engine

- Percent of H2O in Regolith = 2.7%

- Mass of H2O needed/percent = Mass of Regolith needed

- 1950 J to process 1 kg of water from regolith -> 4.22E8 J needed

REFERENCES

- Lots of number from previous presentations

- Project Aldrin Purdue Report

- Loftin, K., Captain, J., & Griffin, T. (2013, September). Presentation to HEMS. In

Integration and Ruggedization of a Commercially Available Gas Chromatograph and Mass

Spectrometer (GCMS) for the Resource Prospector Mission (RPM). Retrieved February 8,

2016, from http://www.hems-workshop.org/9thWS/Presentations/Loftin.pdf

- NASA In-Situ Resource Utilization (ISRU) Development & Incorporation Plans. (n.d.).

Retrieved from https://www.nasa.gov/pdf/203084main_ISRU TEC 11-07 V3.pdf

http://www.hems-workshop.org/9thWS/Presentations/Loftin.pdfd



MISSION DESIGN JAMES MILLANE

Lunar Radiation Environment Definition and Unshielded Dosage

Electrostatic Radiation Shield Analysis

2/17/2016

James Millane

LUNAR RADIATION ENVIRONMENT To design an effective radiation shield, we first need to understand what types of radiation are most harmful on the Moon.

James Millane

Conclusions:

• Astronauts must be protected from lethal SCRs

• We should focus on mitigating GCRs and SCRs, as it will effectively stop all other forms of radiation.

Particle Energy [MeV] Unmitigated

Dose [Sv] Source

GCR ~4 1.5 – 2 Omnidirectional

SCR 1 – 100 160 – 600 Sun (horizontal)

Solar Wind Negligible Sun (horizontal)

Lunar Regolith Neutrons

Negligible Lunar Regolith

*All environmental parameters are functions of particle energy (see backup slides)

ELECTROSTATIC RADIATION SHIELD Advantages:

• Effectively eliminates all SCR particles below 20 MeV

• TRL was 4-5 in 2004, but is possible to raise quickly

• Relatively low mass

• Little interference with comms

• Low energy use for active shield

James Millane

Conclusions:

Electrostatic Spheres will be a good option for future space colonies if we can prove reliability on the Moon, but they should not be primary shielding system for the Lunar base.

Concerns:

• Could interfere with electronics

• Future of TRL is uncertain

• Semi-active system: failure at

the wrong time could be

instantly lethal Figure 1: Sphere Diamond Array

REFERENCES AND ACRONYMS

[1] Miroshnichenko, Leonty I., “Dynamics of Galactic Cosmic Rays,” Radiation Hazard in

Space, Kluwer Academic Publishers, Boston, 2003, pp. 93-116.

[2] Miroshnichenko, Leonty I., “Cosmic Rays of Solar Origin,” Radiation Hazard in Space,

Kluwer Academic Publishers, Boston, 2003, pp. 117-142.

[3] Heiken, Grant H., Vaniman, David T., French, Bevan M., “The Lunar Evironment,”

Lunar Sourcebook, Cambridge University Press, New York, 1991, pp 47-56.

[4] Tribble, Alan C., “The Radiation Environment,” The Space Environment: Implications

for Spacecraft Design, Princeton University Press, New Jersey, 1995, pp 137-162.

Acronym Meaning

GCR Galactic Cosmic Ray

SCR Solar Cosmic Ray

TRL Technology Readiness Level

James Millane

RADIATION DOSAGE CALCULATION PROCESS AND PROGRAM OVERVIEW

Objective:

Calculate the radiation dose an astronaut experiences over their 4 year stay at the lunar base.

Process Flow:

1. Determine the most energetic particles and their fluxes (slides 7 and 8)

2. Determine how much energy is imparted upon human tissue from ionizing radiation given the flux of each particle; bind to LRO CRaTER data for each type of particle.

3. Assuming the particle travels through 15 cm of human tissue, determine the dose of radiation energy deposited in a person based upon particle fluences; report dose in Grays

4. Convert the dose to Sieverts by multiplying by the particle dosages by the quality factor

5. Since the doses are currently functions of energy, integrate over all energies to determine the overall dose

6. Total dose is combined sum of each particle dose

OLTARIS AND RADIATION SHIELD CALCULATION NASA OLTARIS PROGRAM

Problem:

Our developed radiation dose calculation tool is not perfect. We cannot take into account all energies of particles due to numerical issues, nor do we have access to enough data to average the fluxes. Because of this, our computed dose calculations may be several orders of magnitude off. To remedy this, we used OLTARIS for our final radiation shield calculations.

OLTARIS is a NASA tool for the computation of the radiation exposure through a shield. Users can specify a shield and subject it to different radiation environments, and calculate the dosage after the radiation encounters the shield.

This is not to say our program is useless. After we researched and built our flux and fluence models (slides 7 and 8) and were able to show the relative importance of each for a dose calculation (slide 10) we had great insight into the best solution for the total shield system design. We validated our design decision with OLTARIS, and used it to attain more accurate dose values.

SCR PARTICLE FLUXES AND FLUENCES

James Millane

Notes:

• Proton fluxes based on May 31, 1978 Solar Proton Event

• Heavy ion flux calculated from SCR relative abundances

• Fluence found by integrating over 4 years

DATA FROM LUNAR SOURCEBOOK AND “RADIATION HAZARD IN SPACE”

GCR PARTICLE FLUXES AND FLUENCES

James Millane

Notes:

• GCR flux is presented at Solar Minimum (maximum flux)

• Heavy ion flux calculated from GCR relative abundances

• Fluence found by integrating over 4 years

DATA FROM LUNAR SOURCEBOOK AND “RADIATION HAZARD IN SPACE”

LRO AND CRATER DATA DATA TAKEN FROM CRATER EXPERIMENT

James Millane

Notes:

• LET data is for Tissue Equivalent Plastic from LRO’s CRaTER

experiment

• Data is combined with the flux data for GCRs and SCRs to deduce

the LET of each type of particle

TOTAL MISSION DOSE FOR EACH PARTICLE DOES BROKEN UP BY PEARTICLE

James Millane

Particle Calculated Dose

SCR Protons 166.40

SCR Heavy Ions 11.42 Sv

GCR Protons 7.6 mSv

GCR Heavy Ions 0.0266 mSv

Notes:

• Programs only integrate over

available energies, thus

GCR dose is low

ELECTRIC SPHERE EFFECTIVENESS OUTPUT AND INTERPRETATION

To calculate effectiveness of electrostatic spheres, a program was developed to simulate individual particle trajectories inbound to the base at a random angle.

The program computes the percentage of particles that make it within 5, 10, and 20 meters of the center of the array.

James Millane

Above: 20 simulated particle trajectories at 25 MeV

ELECTRIC SPHERE FLUX EFFECTIVENESS FLUX CUT-OUT

Flux before and after electric spheres are implemented

HUMAN FACTORS KATE FOWEE

Radiation Shielding – Dose, Budget

Life support systems

Pressurized Rover

RADIATION Objective: Keeping the radiation dose down to 0.15 Sv in the habitat over 4 years

Reasoning: Legally astronauts need a budget for the radiation they will be subjected to

First floor buried

Passive Shielding

Regolith Bags

4 m thick on sides

2m thick on top

Carbon based fabric bags

Habitat Structure

All hab systems and

materials i.e. carbon fiber,

aluminum, insulation, etc.

(3cm thick on top of

carbon fiber)

Water

Tanks of water at the top

of habs of at least 5 cm

deep.

Kate Fowee

Fig.1: Habitat Radiation Mitigation Design. Cross section and top down views.

GCR SCR

lower

Inter-hab connector

Lunar ground

level Water Tank

Reg. bags

ground

storage

Characteristics

Exit 1 Garage

Exit 2

RADIATION

Mass: 10,160 Mg Regolith, 8 Mg bags, 22440 kg of water necessary

Volume: 5600 m3

Total Dose: 0.113 Sv over 4 years

Recommendation: Cluster the habs to allow for easy bag coverage. Bury the first floor.

Radiation Type

Unmitigated Dose

Mitigated Dose

Top Sides

Galactic Cosmic Rays

1.500-2.000 Sv 0.051 Sv

0.062 SV

Solar Cosmic Rays

160.0-600.0 Sv 0.660 mSv

0.542 mSv

Total 161.5 – 602.0 Sv 0.113 Sv

Kate Fowee

Parameters set before analysis:

<0.5 Sv per astronaut in one 4 year

mission

<0.2 Sv per astronaut in the hab

Reasons:

Career limits for 35 year old women

(1.75 Sv)

1 Sv increases the chance of cancer by

5% (nominal is 20%)

Allow for unshielded rover exploration

time

Allow for veteran astronauts

Rover budget and Recommendations:

No more than 15 hours at a time

Must have permission from Earth/check

that there is no solar event

Table 1: Habitat radiation dosage. Comparison of the unmitigated and mitigated dosage from the most harmful sources of radiation (over 4 years).

RADIATION ANGLE OF INCIDENCE

The radiation from the sun comes in at a fairly acute angle to the visible horizon. The moon’s axis has a 6.8 degree tilt. Combined with how far it is out of the suns equatorial plane at any one time the maximum angle of incidence from the sun to the moon is roughly 10 degrees.

Kate Fowee

Fig.2: Sun and Moon visual. The yellow lines show how the radiation from the sun will hit the Moon base

RADIATION

Angle of Incidence = 10 degrees

Sides: Habitat structure and 4 meters of regolith protect from strong solar

events. 4 meters was calculated to eliminate almost all of SCR.

On top: Habitat structure and 2 meters of regolith bags (and water) will stop

almost all radiation originating from the sun

ANGLE OF INCIDENCE OF SOLAR RADIATION

Inter-hab connector

Lunar ground level

Water Tank Cross section

of Hab

Regolith sand bags

lower

ground

storage

Kate Fowee

Fig.3: Cross section showing SCR. The SCRs come in at less than a 10 degree angle as shown above.

RADIATION

GCRs come from all angles and are significantly harder to stop.

However the likelihood of them hitting and damaging a cell is

significantly lower than the probability of damage from a solar

radiation event. Very few GCRs will get through.

GALACTIC COSMIC RAYS

Inter-hab airlock connection

Lunar ground level

Water Tank Cross section

of Hab

Regolith sand bags

lower

ground

storage

Kate Fowee

Fig.4: Cross section showing GCR. The GCRs come in from all angles above the horizon.

RADIATION RECOMMENDED PROCEDURES – HUMAN FACTORS

For Dose monitoring

Use a personal dosimeter to measure the dose absorbed by the astronauts.

Keep track of any symptoms for radiation sickness each month.

Evacuation of astronauts if more than half the 4 person crew has exceeded the dose limit of .5 Sv in less than 2 years.

During Large Solar Events

NASA receives early warning on the exact severity of a solar event (NOAA DSCOVR)

During severe solar events, the crew should all be located in a corner hab, or whichever hab is designated because these are the habs with the most protection from radiation.

The shielding was designed for the absolute worst case scenario and most solar events should be well below the maximum threshold.

Crew Considerations

While the number of particles that even make it to the crew will be minimal, for extra protection (specifically xrays and any GCR that makes it through), the crew can wear lead or other high density material lined clothing and blankets.

Crew can take medication to lessen any radiation effects. However, it is not known how medications will affect astronauts in lower gravity environments

RADIATION WATER

Recommended to use 5 cm of water to shield at the top of the habs

(0.05m per hab)(3.45m)2(π) = 1.869 m3 per hab

(1.869 m3 per hab)(1000kg/m3) = 1869 kg per hab

(1869 kg per hab)(12 habs) = 22,440 kg

Would like to have 1 year supply of water supply at any time

(14.35 kg/CM.day)(8CM)(365.25.day/ year) = 41,930 kg water / year

Assuming 95% recovery rate of water lose 2,097 kg a year.

The shielding water is well within the needed values for a year.

Need isru water to make up the loss of water.

The water used for shielding can be drank. It is not likely that heavy water would have a high enough concentration.

LIFE SUPPORT SYSTEM

Rapid Activation of Biological

Wastewater Treatment Systems

•Uses inoculum (bacteria) to remove

organic material (95%), ammonium

(95%) and nitrates from waste water

•Bacteria can be freeze dried to have

back up stores incase of emergency

•Techport has a TRL 4 by Dec. 2015

•Based on the speed of development it

is reasonable this tech could be TRL 8

by the beginning of the project

ADVANCED TECHNOLOGY

Lyophilization

•Microwave Enhanced Freeze Drying

of Solid Waste

•Last info from November 2006,

however this is a well developed tech

and could be reasonably forecasted to

be ready for launch date.

•Will freeze dry all wastes to separate

solids and particulates

See sources

Recommendation:

Use both a waste freeze dry system and an inoculum based treatment system in the water system. See next slide for a system diagram showing how the system changes

Kate Fowee

LIFE SUPPORT SYSTEMS MODIFIED

Waste Water Tank

Liquid/Particulate Separator

Filtration

Volatile Removal

Urine Tank

Ion bed

Vapor Distillation &

G/L separation

Microbe Check Valve

Potable Water

Hygiene, Food, Thermal Waste Water

Venting Waste water

Human Waste

Clean water

BASE LINE WATER MANAGEMENT

Waste Water Tank

Biological Waste water

treatment

Lyophilizer

Urine Tank

Vapor Distillation &

G/L separation

Microbe Check Valve

Potable Water

Hygiene, Food, Thermal Waste Water

Venting

ADVANCED WATER MANAGEMENT

Volatiles

Kate Fowee

LIFE SUPPORT SYSTEMS

Component Mass (kg) Volume (m3) Power (We)

Liophilizer and Biological waste treatment 591.2 2.050 2,869

Urine/Waste Water Collection* 4.550 0.020 4.000

Urine, Hygiene & Potable Water & Brine Tanks 181.6 0.470 17.80

Microbial Check Valve 5.720 0.020 0.000

Process Controller 36.11 0.080 156.2

Water Quality Monitoring 14.07 0.040 4.720

Product Water Delivery 51.73 0.120 3.440

Potable Water Storage 595.5 0.440 20.74

Totals 1,480 kg 3.24 m3 3,076 We

TABLE 5. ADVANCED WATER MANAGEMENT SYSTEM MASS, VOLUME, POWER DETAILED

Kate Fowee

Assume 4 needed water systems and two backup systems

Mass: 5920kg** Volume: 12.96m3** Power: 12,300 We**

* ISS proven

Table based on Table 6.14 and 6.9 from Hanford (2005)

**Only 4 running units considered Kate Fowee

REFERENCES • Project Aldrin-Purdue, 2015

• Larson, W. J., and Pranke, L. K., “Human Spaceflight: Mission Analysis and Design”,

Physiology of Space, McGraw-Hill Companies, Inc.New York, 1963, pp103-132

• Hanford, A., J., “Advanced Life Support Research and Technology Development Metric

- Fiscal Year 2005” NASA, CR - 2006 - 213694, February

2006. http://ston.jsc.nasa.gov/collections/trs/_techrep/CR-2006-213694.pdf [retrieved 3

February 2016].

• Cumbie, W., “Rapid Activation of Biological Wastewater Treatment Systems,” NASA

SBIR 2015 Solicitation, FormB – Proposal Summary, published online 2015.

http://sbir.gsfc.nasa.gov/SBIR/abstracts/15/sbir/phase1/SBIR-15-1-H3.02-9111.html [

retrieved 13 February 2016]

• Bailey, M., Lee, C., “Quantities, Units and Ionising Radiation Fundamentals,” Ionising

Radiation Metrology Forum, National Physical Laboratory, November

2010 http://www.npl.co.uk/upload/pdf/20101117_irmf_bailey.pdf [retrieved 17

February2016].

• “OLTARIS,” OLTARIS: On-Line Tool for the Assessment of Radiation in Space , Jul.

2010.

http://ston.jsc.nasa.gov/collections/trs/_techrep/CR-2006-213694.pdf










http://sbir.gsfc.nasa.gov/SBIR/abstracts/15/sbir/phase1/SBIR-15-1-H3.02-9111.html









STRUCTURES ROBERT WHITE

Shielding the HABS

Understanding the Scale

SHIELDING THE HABS ROBERT WHITE

Objective: Determine the most efficient means of construction a radiation

shield for the Habs

Reasoning: Reduce construction time and IMLEO

• Habs are 10 meters tall by 7.4

meter diameter

• Three floors with 4 airlock

hatches on second floor

• Burying depth is to bring airlock

hatches to ground level

• Bottom floor of living modules

will be for bedrooms for

maximum shielding

SHIELDING THE HABS ROBERT WHITE

RECOMMENDATION

Mass: 8 Mg of bags + mass of bagging equipment

Power: Power required by rovers

Volume: 6.4 m^3 for bags

Recommendation: Cover Habs in a group using bagged regolith

• Bury and Cover the Habs in one large group

• Bury the Habs 3.5 meters below surface

• Reduces bagged regolith by 1500 m^3 -> 2.1Mg of bags

• Cover in 2 meters thick ceiling and 4 meters thick wall of bagged

Regolith

• 3 cm thick carbon fiber shell of Habs adds radiation protection

• Store water in a tank at the top of each Hab for added protection

• 5620 m^3 of bagged regolith required for shielding

• Use 1 m^3 bags made of carbon fiber with mass of 1.4 kg per

UNDERSTANDING THE SCALE ROBERT WHITE

THE OUTER DIMENSIONS OF THE HABS 7.4 METERS DIAMETER

7.4 m diameter (distorted from camera orientation)

BACKUP SLIDES ROBERT WHITE

CROSS SECTION VIEW OF BASE

2 meters of bagged regolith

Water Tanks

Ground Level 4 meters of bagged regolith


HAB DESIGN • Habs are 10 meters tall by 7.4 meter diameter

• 0.25 thick walls (for structural support, insulation and basic electrical)

• Water storage tank integrated into the ceiling of the Hab

• Three floors with 4 airlock hatches on second floor

• Floors are removable through the airlock hatches to leave an entire hab open up to the insulation and tank above

• Burying depth is to bring airlock hatches to ground level

• Bottom floor of living modules will be for bedrooms for maximum shielding

• Connecting tunnels are 5 meters in diameter and 8 meters long. A cluster of 4 can fit in a volume less than that of the Hab so can be launched on an SLS and landed on the surface.

• 3 cm thick carbon fiber shell is 3 times the necessitated thickness but it reduces the required shielding at a ratio of 1 to 5 for carbon fiber to water

• Preliminary airlock hatch openings are 1.8 meters tall and .9 meters wide

• Airlock hatches route all electrical, communication and fluids between habs


HAB CROSS SECTION

Water Tank

Hatch Surface Level


Individually surrounding each hab with

bags

• 1,805 m^3 of regolith per hab

• 19,855 m^3 of regolith for all habs

• 27.7 Mg of bags for regolith

TERMINOLOGY

Individually Covering habs with loose

regolith

• 8,413 m^3 of regolith per hab

• Total of 92,543 m^3 of regolith for all

habs

Group Covering habs with loose

regolith

• 22,202 m^3 of regolith

Burying: defined as being below the surface of the moon

Covering: defined as being above the surface of the moon and regolith added on top.

Airlock Hatch: a single airtight door mechanism

Airlock: a structure containing 2 Airlock Hatches that can pressurize and depressurize

without effecting attached volumes

ALTERNATE METHODS OF BURYING HABS


PRELIMINARY CONSTRUCTION TIME ANALYSIS

• CAT 953D Track Loader of equivalent size to rover used as the

estimated bucket size

• Bucket capacity of 1.6 m^3

• Assume each bag fill time of approximately 10 minutes and placing

time of 10 minutes

• For 5620 bags: 1873 hours -> 78 days of 24/7 operation by 1 rover

• Does not include power consumption or charging time considerations

PROPULSION DAYLE ALEXANDER

• Preliminary design requirements for fuel depot and fluids diagram

• Mass, power and volume of required components

55

COMPONENTS AND VOLUME

56 Dayle Alexander

OBJECTIVES • Provide 1 launch of 10Mg re usable ferrying vehicle (now decided a methalox engine) • Produces fuel and oxidizer to accommodate frequency of 1 launch per every 2 years • Able to hold and produce 1 launch worth of extra fuel and oxidizer in case of emergency • Utilizes materials from the ISRU and demonstrates Sabatier and electrolysis processes

MASS, POWER, ISRU REQUIREMENTS

57 Dayle Alexander

POWER REQUIREMENTS

SABATIER PROCESS [m^3] Reaction tank (350/100C) 21.888

ELECTROLYSIS [Mg] Water to process per day 0.0141

STORAGE [m^3]

Methane storage (-173.15C) 20

LOX storage (-208.15C) 28

PROPELLANT REQUIREMENTS (PER LAUNCH)

MASSES [Mg] CH4 4

LOX 16 STORAGE TANK VOLUMES [m^3] CH4 20

LOX 28

ISRU REQUIREMENTS PER DAY [Mg]

Liquid Water 0.0200

Gas Methane 0.0071

Gas Carbon Dioxide 0.0087

Conclusion: • Produces 1 launch worth of fuel in 180 days • Utilizes cold PSR in heat exchangers for cooling of fluids (saves power and doesn’t need

condensers) • Components list in backup slides

REFERENCES • “The Sabatier System: Producing Water on the Space Station”, NASA

Space Station Research,

http://www.nasa.gov/mission_pages/station/research/news/sabatier.ht

ml [retrieved 15 February 2016]

• Barry, P., “Breathing Easy on the Space Station”, NASA Science

News, http://science.nasa.gov/science-news/science-at-

nasa/2000/ast13nov_1/ [retrieved 15 February 2016]

• Starr, M., “Breathe Deep: How the ISS Keeps Astronauts Alive”,

CNET News, http://www.cnet.com/news/breathe-deep-how-the-iss-

keeps-astronauts-alive/ [retrieved 16 February 2016]

• Sutton, G., Biblarz, O., “Propulsion Design Elements, Eight Edition”,

Wiley, [retrieved 17 February 2016]

58 Dayle Alexander

http://www.nasa.gov/mission_pages/station/research/news/sabatier.html

http://www.nasa.gov/mission_pages/station/research/news/sabatier.html

http://science.nasa.gov/science-news/science-at-nasa/2000/ast13nov_1/







http://www.cnet.com/news/breathe-deep-how-the-iss-keeps-astronauts-alive/















BACKUP SLIDE 1 – MATLAB CODE PG1

59 Dayle Alexander

% -MODEL FOR FUEL DEPOT INFO- % AUTHOR: DAYLE ALEXANDER % LAST UPDATED: 2/17/2016 % ASSUMPTIONS: - 1 LAUNCH INCLUDES UP AND DOWN % - LUNAR SURFACE TO LUNAR ORBIT (DV 2200KM/S) % - WILL NEED TO LAUNCH FROM THE SURFACE MIN ONCE EVERY 2 % YEARS % - NEED 1 LAUNCH WORTH OF EMERGENCY FUEL IN HOLDING TANK % - USING 10 MG LANDER % - CAN HARVEST 100 KG OF CH4 A DAY FROM ISRU % - INERT MASS FRACTION OF 0.15 clear all; close all; % -CONSTANTS- % % GENERAL boiloff=0.1; % Boiloff rate for cryogens in space in %/day freq_l=1; % Frequency of launch (with 1 extra) [launch/year] days_y=365; % Number of days in a year [days] % DENSITIES p_gh2=0.0899; % Density of GH2 at 350C [kg/m^3] p_gh2o=575; % Density of GH2O at 350C [kg/m^3] p_gco2=1.977; % Density of GCO2 at 350C [kg/m^2] p_gch4=0.6797; % Density of GCH4 [kg/m^2] p_gox=1.35; % Density of GOX [kg/m^3]

p_lox=1141; % Density of LOX [kg/m^3] p_lch4=421; % Density of LCH4 [kg/m^3] p_lh2o=1000; % Density of LH2O [kg/m^3] p_lco2=1101; % Density of LCO2 [kg/m^3] % MOLAR MASSES mm_h2=0.002; % Molar mass of H2 [kg/mol] mm_h2o=0.018; % Molar mass of H2O [kg/mol] mm_co2=0.044; % Molar mass of CO2 [kg/mol] mm_ch4=0.016; % Molar mass of CH4 [kg/mol] mm_o2=0.032; % Molar mass of O2[kg/mol] % METHANE ENGINE VALUES m_prop=10*2; % Mass of Propellants needed to launch 10 Mg lander [Mg] of=3.8; % O/F ratio for the Raptor engine m_lch4=m_prop/(of+1); % Mass of CH4 needed to launch 20Mg lander [Mg] m_lox=m_lch4*of; % Mass of O2 needed to launch [Mg] % -EQUATIONS- % % GENERAL mol_lch4=m_lch4*1000/mm_ch4; % Mols of CH4 required [mols] mol_lox=m_lox*1000/mm_o2; % Mols of LOX required [moles]


60 Dayle Alexander

% SABATIER PROCESS mol_h2_s=mol_lch4*4; % Mols of H2 required in Sabatier Process [mols] mol_h2o_s=mol_lch4*2; % Mols of H2O generated in Sabatier Process [mols] mol_co2=mol_lch4; % Mols of CO2 required [mols] m_h2_s=mol_h2_s*mm_h2; % Mass of H2 required in Sabatier Process [kg] m_h2o_s=mol_h2o_s*mm_h2o; % Mass of H2O generated in Sabatier Process[kg] m_co2=mol_co2*mm_co2; % Mass of CO2 required [kg] % ELECTROLYSIS mol_h2o_e=mol_h2_s; % Moles of H2O required from electrolysis H2 [moles] mol_o2_e=mol_h2_s/2; % Moles of O2 generated from electrolysis [moles] m_h2o_e=(mol_h2o_e*mm_h2o)/1000; % Mass of H2O in electrolysis H2 [Mg] m_o2_e=mol_o2_e*mm_o2; % Mass of O2 generated from electrolysis [kg] % -RESULTS- % % HOLDING TANK VOLUMES v_lox=(m_lox*1000)/p_lox; % Min volume of LOX needed [m^3] v_ch4=(m_lch4*1000)/p_lch4; % Min volume of LCH4 needed

[m^3] excess_lox=m_o2_e-m_lox; % Extra O2 generated [kg] excess_h2o=m_h2o_s-m_h2o_e*1000; % Extra H2O generated [kg] % GENERATION RATES days_ch4=180; % Number of days given to make required propellants [days] ch4h_perday=20; % Mass of CH4 provided from ISRU per day [kg/day] ch4g_perday=m_lch4*1000/days_ch4-ch4h_perday; % Mass of CH4 required per day [kg/day] co2r_perday=(ch4g_perday/mm_ch4)*1*mm_co2; % Mass of CO2 required per day [kg/day] h2r_perday=(ch4g_perday/mm_ch4)*4*mm_h2; % Mass of H2 required per day [kg/day] h2ogs_perday=(ch4g_perday/mm_ch4)*2*mm_h2o; % Mass of H2O generated in Sabatier per day [kg/day] h2ore_perday=(h2r_perday/mm_h2)*1*mm_h2o; % Mass of H2O required in electrolysis per day [kg/day] h2or_perday=h2ore_perday-h2ogs_perday; % Mass of H2O required from the ISRU per day [kg/day] o2g_perday=(h2r_perday/mm_h2)*0.5*mm_o2; % Mass of O2 generated per day [kg/day]


61 Dayle Alexander

% ISRU TANK VOLUMES vh_ch4=ch4g_perday/p_gch4; vh_h2o=h2ore_perday/p_lh2o; vh_co2=co2r_perday/p_gco2; vh_rt=h2r_perday/p_gh2+co2r_perday/p_gco2; % % PRINT RESULTS fprintf(' MASS/VOLUME PER LAUNCH\n'); fprintf('--------------------------------------------------\n'); fprintf('-Mass of Methane Required [Mg] %.0f -\n',m_lch4); fprintf('-Volume of Methane Holding Tank [m^3] %.0f-\n',v_ch4*2); fprintf('--------------------------------------------------\n'); fprintf('-Mass of LOX Required [Mg] %.0f-\n',m_lox); fprintf('-Volume of LOX Holding Tank [m^3] %.0f-\n',v_lox*2); fprintf('--------------------------------------------------\n'); fprintf(' RAW MATERIALS REQUIRED/GENERATED\n'); fprintf('--------------------------------------------------\n'); fprintf('-Mass of H2O Required [Mg] %.0f-\n',m_h2o_e); fprintf('-Mass of CO2 Required [Mg] %.0f-\n',m_co2/1000); fprintf('--------------------------------------------------\n'); fprintf('-Mass of Excess LOX Generated [Mg] %.0f-

\n',excess_lox/1000); fprintf('--------------------------------------------------\n'); fprintf(' VOLUME OF RAW MATERIAL TANKS\n'); fprintf('--------------------------------------------------\n'); fprintf('-Volume of GCH4 Tank [m^3] %.4f-\n',vh_ch4); fprintf('-Volume of LH2O Tank [m^3] %.4f-\n',vh_h2o); fprintf('-Volume of GCO2 Tank [m^3] %.4f-\n',vh_co2); fprintf('--------------------------------------------------\n'); fprintf(' ISRU REQUIREMENTS PER DAY\n'); fprintf('--------------------------------------------------\n'); fprintf('-Mass of Gas CH4 Required [Mg/day] %.4f-\n',ch4h_perday/1000); fprintf('-Mass of Liquid H2O Required [Mg/day] %.4f-\n',h2or_perday/1000); fprintf('-Mass of Gas CO2 Required [Mg/day] %.4f-\n',co2r_perday/1000); fprintf('--------------------------------------------------\n');


62

fprintf(' POWER REQUIREMENTS\n'); fprintf('--------------------------------------------------\n'); fprintf('-Gas CH4 Heated/Kept at 21C [m^3] %.4f-\n',vh_ch4); fprintf('-Liquid H2O Heated/Kept at 21C [m^3] %.4f-\n',vh_h2o); fprintf('-Gas CO2 Heated/Kept at 21C [m^3] %.4f-\n',vh_co2); fprintf('-Gas H2 and CO2 Heated to 350C in RT [m^3] %.3f-\n',vh_rt); fprintf('-Liquid CH4 Heated/Kept at -173.15C [m^3] %.3f-\n',v_ch4*2); fprintf('-Liquid O2 Heated/Kept at -208.15C [m^3] %.3f-\n',v_lox*2); fprintf('-Current to process ? water in 1 day [kg] %.3f-\n',h2ore_perday); fprintf('--------------------------------------------------\n'); fprintf('-*RT: Reaction Tank: Sabatier process tank where -\n- CO2 and H2 react -\n'); fprintf('-*Cold temps are heated due to -233.15C temp of -\n- PSR (Permanently Shadowed Reigon) -\n'); fprintf('--------------------------------------------------\n');

BACKUP SLIDE 5 – CODE OUTPUT

63

MASS/VOLUME PER LAUNCH ------------------------------------------------------------------- -Mass of Methane Required [Mg] 4 - -Volume of Methane Holding Tank [m^3] 20- ------------------------------------------------------------------- -Mass of LOX Required [Mg] 16- -Volume of LOX Holding Tank [m^3] 28- ------------------------------------------------------------------- RAW MATERIALS REQUIRED/GENERATED ------------------------------------------------------------------- -Mass of H2O Required [Mg] 19- -Mass of CO2 Required [Mg] 11- -------------------------------------------------------------------- -Mass of Excess LOX Generated [Mg] 17- -------------------------------------------------------------------- VOLUME OF RAW MATERIAL TANKS -------------------------------------------------------------------- -Volume of GCH4 Tank [m^3] 4.6317- -Volume of LH2O Tank [m^3] 0.0142- -Volume of GCO2 Tank [m^3] 4.3791- -------------------------------------------------------------------- ISRU REQUIREMENTS PER DAY -------------------------------------------------------------------- -Mass of Gas CH4 Required [Mg/day] 0.0200- -Mass of Liquid H2O Required [Mg/day] 0.0071- -Mass of Gas CO2 Required [Mg/day] 0.0087- --------------------------------------------------------------------

POWER REQUIREMENTS ----------------------------------------------------------------------- -Gas CH4 Heated/Kept at 21C [m^3] 4.6317- -Liquid H2O Heated/Kept at 21C [m^3] 0.0142- -Gas CO2 Heated/Kept at 21C [m^3] 4.3791- -Gas H2 and CO2 Heated to 350C in RT [m^3] 21.888- -Liquid CH4 Heated/Kept at -173.15C [m^3] 19.794- -Liquid O2 Heated/Kept at -208.15C [m^3] 27.753- -Current to process ? water in 1 day [kg] 14.167- ---------------------------------------------------------------------- -*RT: Reaction Tank: Sabatier process tank where - - CO2 and H2 react - -*Cold temps are heated due to -233.15C temp of - - PSR (Permanently Shadowed Region) - ---------------------------------------------------------------------

BACKUP SLIDE 6 – COMPONENTS LIST

64

COMPONENTS LIST

COMPONENT REQUIRED TEMP [°C] REQUIRED VOLUME

[m^3] INLET FLUIDS EXIT FLUIDS

LH2O Tank 21 0.014 LH2O (From ISRU) GH2, GO2

GCO2 Tank 21 4.632 GCO2 (From ISRU) GCO2

GCH4 Tank 21 4.379 GCH4 (From ISRU) GCH4

Reactants Tank 350/100 83.17 GH2,GCO2 LH2O,GCH4

O2 Heat Exchanger -233.15 (no

insulation/heat) ? GO2 LOX

CH4 Heat Exchanger -233.15 (no

insulation/heat) ? GCH4 LCH4

LOX Storage Tank -208.15 28 LOX LOX (To lander)

LCH4 Storage Tank -173.15 20 LCH4 LCH4 (To lander)

BACKUP SLIDE 7 – CONCEPT SKETCH

65

STRUCTURES AUSTIN BLACK

Fuel Depot Systems Sizing, and final M/P/V values

LCH4 Tank

LOX Tank

Reactants Tank GCH4

Tank

GCO2

Tank

LH2O Tank

1

2

3

4 5

6

- Heat Exchanger MLI Outer Shield

PROBLEM COMBINE FUEL DEPOT SYSTEMS

INTO ONE STRUCTURE

Objective: Orient all systems of Fuel Depot in one combined cylinder, while considering heat

transfer from surroundings, radiation, and other tanks within cylinder.

Reasoning: Large cylinder casing allows all systems to be insulated together, allows for easier

SLS transport, and lunar setup.

Tube # Service

1 Liquid Water to LOX Tank

2 Gaseous Hydrogen to RT

3 Liquid Water to LH20 Tank

4 Gaseous Carbon Dioxide to Reactants Tank

5 Gaseous Methane to GCH4 Tank

6 Gaseous Methane to LCH4 Tank

Austin Black

SOLUTION MLI OUTER SHELL, ALUMINUM TANKS, WITH POLYURETHANE FOAM FILLING TANK GAPS,

6061 Aluminum used, due to highest strength to weight ratio, despite having worst conductance

TOTAL:

Mass: 18.74 [Mg] (dry) 85.74 [Mg] (wet)

Power: ~ 500 [W] (to maintain cryogenic storage temperatures)

Volume: 257.373 [m3] (outer vol. of MLI cylinder)

Recommendation: 6061 Aluminum Cylindrical containers for all liquids, MLI/Aluminum outer shell, Polyurethane Foam spacing

COMPONENTS LIST COMPONENT MASS [Mg] MATERIAL VOLUME [m^3]

LH2O Tank 0.0203 0.0075

GCO2 Tank 0.8183 0.3145 GCH4 Tank 0.8491 0.3031

Reactants Tank 2.3781 0.8808

LOX Storage Tank 2.7901 1.0334 LCH4 Storage Tank 2.2326 0.8269

MLI Container 2.69 28.41 Polyurethane Foam 6.96 145

Top Level

Bottom Level

Foam Gap: 6.23 cm

Austin Black

REFERENCE SLIDES SAMPLE PROBLEMS

Heat requirements for top floor

Austin Black

REFERENCE SLIDES

Austin Black

REFERENCE SLIDES

Plots from code created by Brian

O’Neill – Power/Thermo

Group

Austin Black

REFERENCE SLIDES

% Austin Black/Brian O'Neill

clear;clc;close all;

% [AMB LOX LCH4 GCH4 LH2O GCO2 RT]

t = [41.15;65;100;294.15;294.15;294.15;623.15]; % Storage Temp of Fuels [K]

% t = 41.15; % Ambient Temperature (avg. temp of PSR) [K]

tf = t.*(9/5)-459.67; % Temp Kelvin to Farenheit

delta_t = t(1)-t; % Temp Differnece across container wall

l = [0.01:.0025:1]; % Container thickness array

% [LOX LCH4 GCH4 LH2O GCO2 RT]

V = [28;20;4.6317;0.0142;4.3791;22]; % Fuel Volume [m^3]

n = 1;

ri(:,n) = nthroot((2.*V)/(4*pi),3); % Optimal inner radius to minimize surface area

h(:,n) = V./(pi.*(ri.^2)); % Optimal inner height to minimize surface area % Outer Radius [m]

% [LOX LCH4 GCH4 LH2O HCO2]

rho_fuel = [1141;438.89;0.6664;1000;1.98]; % Fuel density [kg/m^3]

% [Al SS Ti]

rho_con = [2700 8000 4430]; % Metal Density

% [LOX LCH4]

M_fuel = [31.2 16.043]; % Fuel Molecular Weight [kg/kmol]

m_fuel = [17 185 24]; % Fuel Mass required [Mg]

R = 8314; % Specific Gas Constant [J/kmol.K]

P_fuel = 101325;

% Calculating hoop and longitudinal stress [MPa]

% Columns - Fuel Type [LH2 LOX LLCH4]

% Rows - Wall thickness array (397 elements)

% Hoop = (Pressure*diameter)/(2*wall_thickness)

% Long = (Pressure*diameter)/(4*wall_thickness)

n = 1;

while n<=6;

hoop(:,n) = ((P_fuel*(2*ri(n)))./(2.*l))/1000000;

long(:,n) = ((P_fuel*(2*ri(n)))./(4.*l))/1000000;

n = n+1;

end

n = 1;

while n<=6;

figure(1)

subplot(2,3,n);

plot(l*100,hoop(:,n),l*100,long(:,n));

if n == 1

title('Hoop and Longitudinal Stress vs Wall Thickness - LOX');

elseif n == 2

title('Hoop and Longitudinal Stress vs Wall Thickness - LCH4');

elseif n == 3

title('Hoop and Longitudinal Stress vs Wall Thickness - GCH4');

elseif n == 4

title('Hoop and Longitudinal Stress vs Wall Thickness - LH2O');

elseif n == 5

title('Hoop and Longitudinal Stress vs Wall Thickness - GCO2');

else

title('Hoop and Longitudinal Stress vs Wall Thickness - RT');

end

xlabel('Wall Thickness [cm]');

ylabel('Hoop Stress [MPa]');

legend('Hoop','Longitudinal');

grid minor;

n = n+1;

end

% [Al SS Ti] Tensile Yield Strength [MPa]

yield = [276 215 880];

FoS = yield./2;

%

% [Al SS Ti] Minimum thickness [m]

% length_LH2 = [0.02 0.02 0.02];

% length_LOX = [0.02 0.02 0.02];

% length_LCH4 = [0.02 0.02 0.02];

% Combining minimum thicknesses into one vector

min = 0.02;

Austin Black

REFERENCE SLIDES

% Outer Diameters for repective fuel tanks

% [Al SS Ti]

n = 1;

while n<=6;

ro(n,:) = min+ri(n);

n = n+1;

end

A_end = pi*(ro.^2);

n = 1;

while n<=6;

vol_out(n,:) = pi*(ro(n,:).^2).*(h(n)+2*min);

vol_in(n,:) = pi*(ri(n).^2)*h(n);

vol(n,:) = vol_out(n,:)-vol_in(n,:);

n = n+1;

end

n = 1;

while n<=6;

mass(:,n) = (vol(n,:)*rho_con(1))/1000;

n = n+1;

end

n = 1;

while n<=4;

% Aluminum

al = [0.07918 1.09570 -0.07277 0.08084 0.02803 -0.09464 0.04179...

-0.00571 0]; % Constants for thermal conductivity calc

rho_al = 2700; % density

% 304 Stainless Steel

ss = [-1.04087 1.3982 0.2543 -0.6260 0.2334 0.4256 -0.4658 0.1650 -0.0199];

rho_ss = 8000; % density

% Ti-6Al-4V Titanium

ti = [-5107.8774 19240.422 -30789.064 27134.756 -14226.379 4438.2154...

-763.07767 55.796592 0];

rho_ti = 4430; % density

% Heat Transfer Coefficients

k_al(n,:) = 10.^(al(1)+al(2)*log10(t(n))+(al(3)*log10(t(n)).^2)...

+(al(4)*log10(t(n)).^3)+(al(5)*log10(t(n)).^4)+(al(6)*...

log10(t(n)).^5)+(al(7)*log10(t(n)).^6)+(al(8)*log10(t(n)).^7)+(al(9)*log10(t(n)).^8));

k_ss(n,:) = 10.^(ss(1)+ss(2)*log10(t(n))+(ss(3)*log10(t(n)).^2)+(ss(4)*log10(t(n)).^3)...

+(ss(5)*log10(t(n)).^4)+(ss(6)*log10(t(n)).^5)+(ss(7)*log10(t(n)).^6)+...

(ss(8)*log10(t(n)).^7)+(ss(9)*log10(t(n)).^8));

k_ti(n,:) = 10.^(ti(1)+ti(2)*log10(t(n))+(ti(3)*log10(t(n)).^2)+(ti(4)*log10(t(n)).^3)...

+(ti(5)*log10(t(n)).^4)+(ti(6)*log10(t(n)).^5)+(ti(7)*log10(t(n)).^6)+...

(ti(8)*log10(t(n)).^7)+(ti(9)*log10(t(n)).^8));

%k = [k_al k_ss k_ti];

n = n+1;

end

k = [k_al k_ss k_ti];

k_MLI = 0.00004;

n = 1;

while n<=6;

ISA(n) = 2*pi*ri(n).^2+2*pi*ri(n).*h(n);

OSA(n,:) = 2*pi*ro(n,:).^2+2.*pi.*ro(n,:).*(h(n)+2*min);

s(n,:) = ISA(n) + OSA(n,:) ./2; %meters/2

n = n+1;

end

emm = 0.9; % average emissivity of lunar regolith

e = 0.04; % emissivity of MLI aluminum

emm = 0.3; % emissivity of mylar

boltz = 5.67e-8; % Boltzman constant W/m^2*K^-4

Tsun = 5779; % Temperature of sun [K]

qr1_MLI = (e./((n+1).*(2-e))).*boltz.*(Tw1.^4-Tw2.^4); % heat transfer per unit area for radiation heat transfer through layers [Watts/m^2]

qr2_MLI = e*boltz*(Tsun - Tw2);

q_MLI = qr1_MLI+qr2_MLI

n = 1;

while n<=4;

R_shell(n,:) = (log(ro(n,:)./ri(n)))./(2*pi*k(n,:).*h(n));

n = n+1;

end

layers = 1 : 100;

L = 0.0127 .* layers;

n = 1;

while n<=9;

if n<=3

R_MLI(:,n) = (log((ro(n)+L)./ro(n)))/(2*pi*k(n)*h(n));

elseif 4<=n && n<=6

R_MLI(:,n) = (log((ro(n)+L)./ro(n)))/(2*pi*k(n)*h(n-3));

else

R_MLI(:,n) = (log((ro(n)+L)./ro(n)))/(2*pi*k(n)*h(n-6));

end

R_total(:,n) = R_shell(n)+R_MLI(:,n)+Rrad(n);

n = n+1;

end

n = 1;

while n<=9;

if n <=3

Q_total(:,n) = (delta_t(n)./R_total(:,n))/1000;

elseif 4<=n && n<=6

Q_total(:,n) = (delta_t(n-3)./R_total(:,n))/1000;

else

Q_total(:,n) = (delta_t(n-6)./R_total(:,n))/1000;

end

n = n+1;

end

Austin Black

REFERENCE SLIDES

figure(2);

n = 1;

while n<=9;

if n<=3

subplot(2,2,1);

plot(layers,Q_total(:,n));

hold on;grid on;

title('MLI Layers vs Total Heat Transfer - Aluminum');

xlabel('MLI Layers');

ylabel('Heat Transfer [kW]');

legend('LH_2','LOX','LCH_4');

elseif 4<=n && n<=6

subplot(2,2,2);


hold on;grid on;

title('MLI Layers vs Total Heat Transfer - Stainless Steel');




else

subplot(2,2,3);


hold on;grid on;

title('MLI Layers vs Total Heat Transfer - Titanium');




end

n = n+1;

end

Austin Black

REFERENCE SLIDES

%% Heat Transfer Analysis for Insulating Habs

%% Weronika Juszczak

%% Defining Variables

n = [0:1:100]; % number of layers

ks = 0.00004; % Thermal conductivity of MLI spencer [W/m*K]

Tw1 = 293; % Temperature of Inside Wall [K]

Tw2 = 41.15; % Temperature of Outside Wall [K]

L = .002; % length between layers of MLI [m]



emm = 0.3; % emissivity of mylar



%% Heat Transfer Across MLI with Varying n

% Conduction Heat Transfer of Spacer

qc_MLI = ks*(Tw1-Tw2)./(n*L);

% Radiation Heat Transfer


qr2_MLI = e*boltz*(Tsun - Tw2);

qtot_MLI = qr1_MLI+qr2_MLI+qc_MLI;

figure

plot(n, qtot_MLI,n,qr1_MLI+qr2_MLI,n,qc_MLI)

legend('total','radiation','conduction')

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

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

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

figure

plot(L.*n,qtot_MLI)

title('Heat Transfer vs. Thickness of MLI','FontSize',20)

xlabel('Thickness [m]')


% Effective Emmitance of MLI (test of efficiency)

e1 = e;% emissivity of surface 1

e2 = e; % emissivity of surface 2

eff = (((2.*n)./emm)-n-1+(1/e1)+(1/e2)).^-1;

plot(n,eff)

title('Effective Emittance of MLI','FontSize',20)

xlabel('Number of Layers')

ylabel('Eff')

%% Heat Transfer Through Regolith of Varying Thickness

k = .015; % Thermal conductivity of Lunar Regolith 1 M thick [W/m*K]

x = [.5:.1:4]; % Thickness

qc = k*(Tw1-Tw2)./(x); % Heat flux of regolith of varying thickness

qr = boltz*emm*(Tsun-Tw2);

q_tot = qc + qr;

%% FOR CYLINDER

%Brian O'Neill

Thickness = n*L

ri = 4 %meters INSIDE RADIUS OF CYLINDER

OSA = 2*pi*(ri+Thickness)+2*pi*(ri+Thickness).^2 %OUTSIDE SURFACE AREA

Austin Black

REFERENCE SLIDES

ISA = 2*pi*(ri)+2*pi*(ri).^2 %INSIDE SURFACE AREA

ASA = (OSA + ISA)/2 %AVERAGE BETWEEN INSIDE AND OUTSIDE SURFACE AREA'S

t = 1:.1:28.5

day_night_cycle = .5*sin(2*pi*t/28.5)+.5 %CHECK WITH SCIENCE IF YOU WANT TO USE THIS

figure

plot(t,day_night_cycle)

Q = ASA.*qtot_MLI;

figure

plot(Thickness,Q)



ylabel('Heat transfer [Watts]')

figure

plot(n,Q)

title('Heat Transfer vs. Layers of MLI','FontSize',20)

xlabel('Layers of MLI')


Austin Black

REFERENCES

file:///U:/Personal/Downloads/NGL-PUB-17021-co-storage-of-cryogenic-propell%20(7).pdf

http://www.lpi.usra.edu/lunar/documents/NASA%20CR-190014.pdf

http://www.foamforyou.com/Foam_Specs.htm

http://www.engineeringtoolbox.com/polyurethane-insulation-k-values-d_1174.html

http://web.mit.edu/16.unified/www/FALL/thermodynamics/notes/node118.html

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19890012048.pdf

http://ntrs.nasa.gov/search.jsp?R=19900040654

http://www.ulalaunch.com/uploads/docs/Published_Papers/Exploration/LunarLanderConfig

urationsIncorporatingAccessibility20067284.pdf

Austin Black

//roger.ecn.purdue.edu/black34/Personal/Downloads/NGL-PUB-17021-co-storage-of-cryogenic-propell (7).pdf















http://www.lpi.usra.edu/lunar/documents/NASA CR-190014.pdf



http://www.foamforyou.com/Foam_Specs.htm










http://web.mit.edu/16.unified/www/FALL/thermodynamics/notes/node118.html


http://ntrs.nasa.gov/search.jsp?R=19900040654

http://www.ulalaunch.com/uploads/docs/Published_Papers/Exploration/LunarLanderConfigurationsIncorporatingAccessibility20067284.pdf

http://www.ulalaunch.com/uploads/docs/Published_Papers/Exploration/LunarLanderConfigurationsIncorporatingAccessibility20067284.pdf

SCIENCE GROUP CALEB ENGLE

18 February 2016

Extraction techniques for water/volatiles

WATER IS TRAPPED IN REGOLITH

Objective: Develop efficient method for extracting water from lunar

regolith

Reasoning: Water is critical for our mission

• We have lunar regolith and ice crystals mixed together and we want to separate the ice crystals and turn them into liquid water.

• We need a method that is efficient and uses low power.

regolith + water Liquid water

USE MICROWAVES TO EXTRACT VOLATILE

Mass: 35.3kg

Power: 250kW

Volume: 89.77m3

Recommendation: Use microwaves to sublimate volatile then condense it to a liquid

• Microwaves can

efficiently sublimate

volatiles in regolith

• Connect a

condenser to

change the gas

into a liquid

Microwave Condenser

Volatile and Regolith Mix Liquid Volatile

BACKUP SLIDE

• 1,000/27.93 = 35.8

• So with this method, 35.8kg of regolith makes

1kg of water

Mass of regolith (kg)

Mass of water in regolith (kg)

Mass sublimated (kg)

Mass condensed (kg)

1,000 30 28.5 27.93

BACKUP SLIDE

• Volume of microwave needed to hold regolith

Mass of water pressurized

rover can hold (kg)

Mass of regolith for 1kg water

(kg)

Total regolith needed (kg)

Density of regolith (kg/m3)

Microwave volume needed

(m3)

3,610 35.8 129,238 1500 86.16

• Volume of condenser will need to be 3.61m3

• If made from carbon fiber total weight would be 35.3kg

BACKUP SLIDE

• Pressurized rover tanks to transport water from PSR to HAB.

• Holds 3,610kg liquid water

Designed by Ariel Dimston

BACKUP SLIDE

• Experiment extracted 2g

of water from 200g

regolith in 2 minutes from

regolith stimulant using

1KW microwave

• 95% of water in regolith

simulant was sublimated

• 98% of sublimated water

was captured by cold trap

Ethride, C. Edwin, Kaukler, William. Extraction of Water From Lunar Polar Regolith. January 2009.

BACKUP SLIDE

• Advantages of using microwave method

• Low power requirement

• Penetrates deep into regolith

• Fast process

• Lunar regolith transports microwaves

very well

POWER/THERMO BRIAN O’NEILL

Insulation for Fuel depot and power required

PROBLEM DESCRIPTION THERMAL ENVIROMENT

Objective: Determine Insulation and heating required for Fuel depot

Reasoning: Fuels need to be stored at particular temperatures in

LC

H4

LC

O2

LH2

O

LCH(

h)

LOX

Fuel Depot

254.15

254.15

254.15

458.15

25 60

ΔT in Degrees Celsius

LH2O GCO2 GCH4 Reactants LOX Storage LCH4 Storage

RESULTS OF THERMAL ANALYSIS

Mass: 3.136Mg(for insulation)

Power: 176.86 Watts for heat

exchanger(ideal)

Volume: 47.52 m^3

Recommendation: 8cm thickness of

Insulation for super structure and an

immersive heat exchanger system

CYLINDRICAL RADIAL/CONVECTIVE/CONDUCTIVE MODEL

Liquid Tank Liquid Temperature [C] Delta WRT Enviroment [C] VOLUME [m^3] Heat Loss [Watts]

LH2O 21 254.15 0.014 0.2362

GCO2 21 254.15 4.632 7.4031

GCH4 21 254.15 4.379 7.083

Reactants 225 458.15 22 158.4164

LOX Storage -208.15 25 28 1.2896

LCH4 Storage -173.15 60 20 2.428

BACK UP SLIDES

BACK UP SLIDES

%% Heat Transfer Analysis for Insulating Habs

%% Weronika Juszczak

close all

%% Defining Variables

n = [0:1:100]; % number of layers

ks = 0.00004; % Thermal conductivity of MLI spencer [W/m*K]

Tw1 = 100

; % Temperature of Inside Wall [K]

Tw2 = 40; % Temperature of Outside Wall [K] now the temp of Permenately shadowed region

L = .002; % length between layers of MLI [m]



em = 0.3; % emissivity of mylar



%% Heat Transfer Across MLI with Varying n

% Conduction Heat Transfer of Spacer

qc_MLI = ks*(Tw1-Tw2)./(n*L);

% Radiation Heat Transfer


qr2_MLI = e*boltz*(Tsun - Tw2)*.216;

qtot_MLI = qr1_MLI+qr2_MLI+qc_MLI;

figure

plot(n, qtot_MLI,n,qr1_MLI+qr2_MLI,n,qc_MLI)

legend('total','radiation','conduction')

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

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


figure

plot(L.*n,qtot_MLI)




MATLAB SCRIPT USED

% Effective Emmitance of MLI (test of efficiency)

e1 = e;% emissivity of surface 1

e2 = e; % emissivity of surface 2

eff = (((2.*n)./emm)-n-1+(1/e1)+(1/e2)).^-1;

plot(n,eff)

title('Effective Emittance of MLI','FontSize',20)

xlabel('Number of Layers')

ylabel('Eff')

%% Heat Transfer Through Regolith of Varying Thickness

k = .015; % Thermal conductivity of Lunar Regolith 1 M thick [W/m*K]

x = [.5:.1:4]; % Thickness

qc = k*(Tw1-Tw2)./(x); % Heat flux of regolith of varying thickness

qr = boltz*emm*(Tsun-Tw2);

q_tot = qc + qr;

%% FOR CYLINDER

%Brian O'Neill

Volume = 20

%m^3

Cyl_length = .9 %axial length of cylinder in meters

% Thickness = n*L

Thickness = .08;

ri = sqrt(Volume./(pi*(Cyl_length-2*Thickness))) %meters INSIDE RADIUS OF CYLINDER

OSA = 2*pi*(ri+Thickness)*Cyl_length+2*pi*(ri+Thickness).^2 ;%OUTSIDE SURFACE AREA

ISA = 2*pi*(ri)*Cyl_length-(2.*Thickness)+2*pi*(ri).^2; %INSIDE SURFACE AREA

ASA = (OSA + ISA)/2; %AVERAGE BETWEEN INSIDE AND OUTSIDE SURFACE AREAS

t = 1:.1:28.5;

day_night_cycle = (1-.213)/2*sin(2*pi*t/28.5)+1.213/2; %CHECK WITH SCIENCE IF YOU WANT TO USE THIS

figure

plot(t,day_night_cycle)

title('Moon cycle','FontSize',20)



Q = ASA.*qtot_MLI%(.08/.002)

figure

plot(Thickness,Q)

title('Heat Transfer vs. Thickness of MLI, For Super Assembly','FontSize',15)



figure

plot(n,Q)

title('Heat Transfer vs. Layers of MLI,For Super Assembly','FontSize',15)

xlabel('Layers of MLI')


%% Volume calculator

ri2 = 3.5; %m

Cyl_length = 2 %m

Thickness = .08 %m

Volume = pi*(ri2+Thickness)^2*Cyl_length-pi*ri2^2*(Cyl_length-2*Thickness)

density = (95+37)/2 %[kg/m^3]for MLI A144 cyrostat

Mass = density*Volume/1000

Q = ASA.*qtot_MLI(.08/.002)

POWER AND THERMAL TYLER MURRAY

Power Required for Base Per Mission

POWER CONSUMPTION PER MISSION Objective: Determine max power consumption per mission.

Reasoning: To understand how much power generation equipment is needed and to

properly allocate energy.

• Used updated mission timeline to determine power increase per mission on the base

• As base expands, electrical components added (refrigerators, additional water recycling, printers, rovers, etc.)

• Considered 2nd option of food growth (Growing part of the food on base)

Spikes:

• Adams 3: Testing ISRU, Medical Module delivered and heated to avoid freezing electrical components

• Jefferson 9: ISRU production begins

• Jefferson 14: 2nd living module introduced

Plateaus:

• Jefferson 5-9: Landing power gen equipment and rec center equipment

• Jefferson 11-13: Landing for second hab components

• Madison 1-2: Missions to cycler, power on base will remain the same

Tyler Murray

POWER CONSUMPTION PER MISSION

Mass: 1.024 Mg

Power: 166.094 kW

Dimensions: 1.85 m^3 (nuclear reactors) 16 m^2 (solar panels)

Recommendation: 2 Nuclear Reactors and .013 kg III Solar Panels (~220 kW)

Tyler Murray

BACK UP SLIDES - CODE

Tyler Murray

%Tyler Murray

%AAE 450

hab_t = 7; %[kW]

water_t = 9.4; %[kW]

hab_o = 4.2; %Additional energy needed to heat to 75 F

(same as ISS) [kW]

hab_f = hab_t + hab_o; %Energy needed to heat directly to

75 F

rover = .2; %[kW]

atmo = 4.195; %[kW]

wu_recycle = 3.24; %[kW]

seis = .01; %[kW]

ISRU = 10; %[kW]

med_mod = 3.787; %[kW]

print = .1; %[kW]

food = 5.4; %[kW]

inside_hab = 2.152; %[kW]

misc = .96; %[kW]

fuel_depot = 5; %[kW]

rec_c = .26; %[kW]

%Washington Series

%Missions 1 & 2 do not reach lunar surface

Wash1 = 0;

Wash2 = 0;

Wash3 = rover;

%Adams Series

%Mission 1 does not reach lunar surface

Adams1 = Wash3;

Adams2 = wu_recycle + seis + rover;

Adams3 = ISRU + med_mod + print + 4 * seis + hab_t +

rover; %Test equipment, shut off when done

Adams4 = hab_t + seis + rover + atmo; %medical hab left

on surface, heat so electrical components don't freeze

Adams5 = Adams4;

Adams6 = hab_t + wu_recycle + Adams5; %medical hab and

food storage need heating avoid freezing

Adams7 = water_t + hab_t + Adams6 + atmo; %water storage

heated to avoid freezing

Adams8 = hab_t + 2*wu_recycle + Adams7; %Add waste

management module

%Jefferson Series

Jeff1 = 2*hab_o + (water_t - hab_t) + inside_hab +misc +

atmo + wu_recycle + print + Adams8;

Jeff2 = fuel_depot + Jeff1 + atmo;

Jeff3 = rec_c + hab_f + atmo + wu_recycle + Jeff2;

Jeff4 = hab_f + Jeff3;

Jeff5 = water_t + atmo + atmo + Jeff4; %Garage

Jeff6 = Jeff5;

Jeff7 = Jeff6;

Jeff8 = Jeff7;

Jeff9 = ISRU + Jeff8;

Jeff10 = Jeff9;

Jeff11 = hab_f + inside_hab + misc + atmo + wu_recycle +

print + Jeff10;

BACK UP SLIDES - CODE

Tyler Murray

Jeff12 = Jeff11;

Jeff13 = Jeff12;

%Madison Series

%No difference in power, sending crew to cycler for Mars

Mission

Mad1 = Jeff13;

Mad2 = Mad1;

figure(1)

plot(0:2, [Wash1 Wash2 Wash3],'g');hold on

set(gca,'XTick',[0 2 10 22])

set(gca,'XTickLabel',{'Washington'})

plot(2:10, [Adams1 Adams2 Adams3 Adams4 Adams5 Adams6

Adams7 Adams8 Jeff1],'b')

% set(gca,'XTick',3)

set(gca,'XTickLabel',{'Adams'})

plot(10:22, [Jeff1 Jeff2 Jeff3 Jeff4 Jeff5 Jeff6 Jeff7

Jeff8 Jeff9 Jeff10...

Jeff11 Jeff12 Jeff13],'r')

plot(22:23, [Mad1 Mad2],'k')

legend('Washington','Adams','Jefferson','Madison')

set(gca,'XTickLabel',{'Washington','Adams','Jefferson','M

adison'})

title('Lunar Base Power Consumption')

xlabel('Mission Series')

ylabel('Power Consumption [kW]')

grid on;

BACK UP SLIDES

Work done by Kelly Kramer

Tyler Murray

BACK UP SLIDES

Work done by Kelly Kramer

Work done by Mike Waldmann

Work done by Rachael Hess

Tyler Murray

REFERENCES

"Nuclear Reactors and Radioisotopes for Space." Nuclear Reactors for Space. World Nuclear, Feb. 2016. Web. 17 Feb. 2016. <http://world-nuclear.org/information-library/non-power-nuclear-applications/transport/nuclear-reactors-for-space.aspx>.

SCIENCE GROUP RACHEL MAXWELL

STM (Science Traceability Matrix)

Science Objectives

Analogs for Mars

99

SCIENCE TRACEABILITY MATRIX OBJECTIVES THAT DETERMINE THE INSTRUMENTS

Objective: To determine science instruments for the mission

Reasoning: Provide data to lower risk factors and increase science returns Science Objective Measurement Objective Measurement Requirements

Sample Return Age and trace elements Quality of sample Age within ±20My

Regolith composition Look for volatiles Analyze mineralogy

Detectability at ppb scale Quality of sample

Landscape Analysis Determine hazardous morphology

Resolution of images (best Mastcam-Z res: ~0.5mm/pix at 2m)

Radiation Determine radiation environment

Detectability (LET* of ~0.2 keV/µm – several hundred keV/µm)

Dust Determine dust density Sensitivity (LDEX: 0.4 – 4.0 x 10-3 m-3)

Moonquakes Determine seismologic hazards

Sensitivity

100 Rachel Maxwell, Science *LET = Linear Energy Transfer

Prio

rity High

est to Lo

west

MISSION OBJECTIVES HOW THE INSTRUMENTS MATCH OUR NEEDS AND THE DECADAL SURVEY

101 Rachel Maxwell, Science

Satisfy the needs of the mission

• Identify exact locations best suited for ISRU

• Safety planning

Understand origin and diversity of terrestrial planets

• Constrain the bulk composition of planets to understand formation and evolution

• Understand geologic processes and their effects

Understand how the evolution of terrestrial planets enables and limits the origin and evolution of life

• Composition and distribution of volatiles

Based on previous instruments (see backup slides)

Mass: 83.92 kg

Power: > 136.8 W [data missing for one instrument]

Volume: 0.0492 m3

Recommendation: Use

BACKUP SLIDES SUGGESTED SCIENCE INSTRUMENTS

Instrument Description/ Use Current Instruments Placement TRL

Drill / Sample Cache Bring samples back to Hab

TBD Science Rover

Gas Chromatograph/ Mass Spectrometer

Composition SAM (MSL) Hab 9

Imaging Analyze Lunar surface Mastcam-Z (Mars 2020) Science Rover

8/9

Spectrometer Mineralogy and Composition

ChemCam (MSL) SuperCam (Mars 2020)

Science Rover

9

Neutron Analysis Measure hydrogen content

DAN (MSL) Science Rover

9

Radiation Detector Direct radiation measurements

RAD (MSL) Science Rover, Probe

9

Seismometer Study Moonquakes TBD (SEIS (ExoMars)) Probe (5)


BACKUP SLIDES

Acronym Meaning

MSL Mars Science Laboratory

SAM Sample Analysis at Mars

ChemCam Chemistry and Composition

DAN Dynamic Albedo of Neutrons

RAD Radiation Assessment Detector

SEIS SEISmometer

ExoMars Exobiology on Mars

ACRONYM GUIDE


MASS, POWER, VOLUME

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

Mastcam-Z 4.5 11.8 0.00882

ChemCam 5.778 [unavailable] 0.000013271

RAD 1.6 4.2 0.00024

DAN 2.6 13 0.0019025

SEIS 6.5 3.4 0.0013

Total 20.98 32.4 0.0123

Total for four rovers, four probes

83.92 136.8 0.0492

104

BASED ON INSTRUMENT SPECIFICATIONS IN LITERATURE

Rachel Maxwell, Science

REFERENCES

105

Bell, J.F. III, Maki, J.N., Mehall, G.L., Ravine, M.A., Caplinger, M.A., and the Mastcam-Z Science Team. “Mastcam-Z: A Geologic, Stereoscopic, and Multispectral Investigation on the NASA Mars-2020 Rover,”

Champion, S., and Shearer, C., “Lunar Geophysical Network ( LGN ) Planetary Science Decadal Survey,” Evolution. Elphic, R.C., Horanyi, M., Colaprete, A., Benna, M., Mahaffy, P.R., Delory, G.T., Noble, S.K., Halekas, J.S., Hurley, D.M., Stubbs, T.J., Sarantos, M., Kempf, S.,

Poppe, A., Szalay, J., Sternovsky, S., Cooke, A.M., Wooden, D.H., Glenar, D., “LADEE Science Results and Implications for Exploration,”46th LPSC, 2015 Grotzinger, J. P., Crisp, J., Vasavada, A. R., Anderson, R. C., Baker, C. J., Barry, R., Blake, D. F., Conrad, P., Edgett, K. S., Ferdowski, B., Gellert, R., Gilbert, J. B.,

Golombek, M., Gómez-Elvira, J., Hassler, D. M., Jandura, L., Litvak, M., Mahaffy, P., Maki, J., Meyer, M., Malin, M. C., Mitrofanov, I., Simmonds, J. J., Vaniman, D., Welch, R. V., and Wiens, R. C., Mars Science Laboratory mission and science investigation, 2012.

Hassler, D. M., Zeitlin, C., Wimmer-Schweingruber, R. F., Böttcher, S., Martin, C., Andrews, J., Böhm, E., Brinza, D. E., Bullock, M. A., Burmeister, S., Ehresmann, B., Epperly, M., Grinspoon, D., Köhler, J., Kortmann, O., Neal, K., Peterson, J., Posner, A., Rafkin, S., Seimetz, L., Smith, K. D., Tyler, Y., Weigle, G., Reitz, G., and Cucinotta, F. A., “The Radiation Assessment Detector (RAD) investigation,” Space Science Reviews, vol. 170, 2012, pp. 503–558.

Mahaffy, P. R., Webster, C. R., Cabane, M., Conrad, P. G., Coll, P., Atreya, S. K., Arvey, R., Barciniak, M., Benna, M., Bleacher, L., Brinckerhoff, W. B., Eigenbrode, J. L., Carignan, D., Cascia, M., Chalmers, R. A., Dworkin, J. P., Errigo, T., Everson, P., Franz, H., Farley, R., Feng, S., Frazier, G., Freissinet, C., Glavin, D. P., Harpold, D. N., Hawk, D., Holmes, V., Johnson, C. S., Jones, A., Jordan, P., Kellogg, J., Lewis, J., Lyness, E., Malespin, C. A., Martin, D. K., Maurer, J., McAdam, A. C., McLennan, D., Nolan, T. J., Noriega, M., Pavlov, A. A., Prats, B., Raaen, E., Sheinman, O., Sheppard, D., Smith, J., Stern, J. C., Tan, F., Trainer, M., Ming, D. W., Morris, R. V., Jones, J., Gundersen, C., Steele, A., Wray, J., Botta, O., Leshin, L. A., Owen, T., Battel, S., Jakosky, B. M., Manning, H., Squyres, S., Navarro-Gonzelez, R., McKay, C. P., Raulin, F., Sternberg, R., Buch, A., Sorensen, P., Kline-Schoder, R., Coscia, D., Szopa, C., Teinturier, S., Baffes, C., Feldman, J., Flesch, G., Forouhar, S., Garcia, R., Keymeulen, D., Woodward, S., Block, B. P., Arnett, K., Miller, R., Edmonson, C., Gorevan, S., and Mumm, E., “The Sample Analysis at Mars Investigation and Instrument Suite,” Space Science Reviews, vol. 170, 2012, pp. 401–478.

Maurice, S., Wiens, R. C., Saccoccio, M., Barraclough, B., Gasnault, O., Forni, O., Mangold, N., Baratoux, D., Bender, S., Berger, G., Bernardin, J., Berth, M., Bridges, N., Blaney, D., Bouye, M., Ca??s, P., Clark, B., Clegg, S., Cousin, A., Cremers, D., Cros, A., Deflores, L., Derycke, C., Dingler, B., Dromart, G., Dubois, B., Dupieux, M., Durand, E., D’Uston, L., Fabre, C., Faure, B., Gaboriaud, A., Gharsa, T., Herkenhoff, K., Kan, E., Kirkland, L., Kouach, D., Lacour, J. L., Langevin, Y., Lasue, J., Le Moulic, S., Lescure, M., Lewin, E., Limonadi, D., Manh??s, G., Mauchien, P., McKay, C., Meslin, P. Y., Michel, Y., Miller, E., Newsom, H. E., Orttner, G., Paillet, A., Pares, L., Parot, Y., Perez, R., Pinet, P., Poitrasson, F., Quertier, B., Sall, B., Sotin, C., Sautter, V., S??ran, H., Simmonds, J. J., Sirven, J. B., Stiglich, R., Striebig, N., Thocaven, J. J., Toplis, M. J., and Vaniman, D., “The ChemCam instrument suite on the Mars Science Laboratory (MSL) rover: Science objectives and mast unit description,” Space Science Reviews, vol. 170, 2012, pp. 95–166.

Maurice, S., Wiens, R. C., Anderson, R., Beyssac, O., Bonal, L., Clegg, S., DeFlores, L., Dromard, G., Fischer, W., Forni, O., Gasnault, O., Grotzinger, J., Johnson, J., Martinez-Frias, J., Mangold, N., McLennan, S., Montmessin, F., Rull, F., Sharma, S., Fouchet, T., Poulet, F., and Team, T. S., “Science Objectives of the SuperCam Instrument for th Mars2020 rover,” Lunar Planetary Sciences Conference, vol. 10, 2015, pp. 6–7.

Mitrofanov, I. G., Litvak, M. L., Varenikov, A. B., Barmakov, Y. N., Behar, A., Bobrovnitsky, Y. I., Bogolubov, E. P., Boynton, W. V., Harshman, K., Kan, E., Kozyrev, A. S., Kuzmin, R. O., Malakhov, A. V., Mokrousov, M. I., Ponomareva, S. N., Ryzhkov, V. I., Sanin, A. B., Smirnov, G. A., Shvetsov, V. N., Timoshenko, G. N., Tomilina, T. M., Tret’Yakov, V. I., and Vostrukhin, A. A., “Dynamic Albedo of Neutrons (DAN) experiment onboard NASA’s Mars Science Laboratory,” Space Science Reviews, vol. 170, 2012, pp. 559–582.

REFERENCES

106

LINKS TO SCIENCE INSTRUMENTS MENTIONED

SuperCam (Mars 2020) http://mars.nasa.gov/mars2020/mission/science/for-scientists/instruments/supercam/

MastCam-Z (Mars 2020) http://mars.nasa.gov/mars2020/mission/science/for-scientists/instruments/mastcam-z/

DAN (MSL) http://mars.nasa.gov/msl/mission/instruments/radiationdetectors/dan/

RAD (MSL) http://mars.nasa.gov/msl/mission/instruments/radiationdetectors/rad/

SAM (MSL) http://msl-scicorner.jpl.nasa.gov/Instruments/SAM/

CONTROL BECCA PIETRZYCKI

Communication Scheme

Ground sites and Comm. Sat. Antenna Selection

COMMUNICATION SCHEME Objective: Understand communication scheme and choose antennas for Earth, Moon, and

Communication Satellites to set the stage for the rest of the vehicles.

Reasoning: Complete requirement for constant 2-way video throughout the mission.

Moon Base

Earth bases

XM-1

Orion

Comm Sats (x3)

XM-2/3

Lander

Habitats

Rovers

Earth

EARTH, MOON, & COMM. SAT. ANTENNAS

Recommendations:

• Use 9.4 Meter Earth Station Antenna from ASC Signal for Earth and Moon bases. Use

2 smaller antennas for communication satellites.

• Smaller antennas will be added to communication satellite design.

• Moon base antenna will need to be launched on cargo mission.

Location Diameter (m)

Mass (Mg)

Power (kW)

Number of antennas

Earth (U.S. and Japan?)

9.4 4.7 100 2

Moon Orbit (to bases)

0.1064 0.47 0.65 1 (per satellite)

Moon Base 9.4 4.7 100 1

EARTH GROUND BASE TO COMM. SATS.

Inputs

Frequency [Hz] 8.00E+09

Transmitter Power [W] 100000

Transmitter Line Loss [dB] -1

Transmit Antenna Beam Width [deg] 2.47E+01

Transmit Antenna Pointing Offset [deg] 2.34E+01

Propagation Path Length [m] 3.844E+08

Propagation/Polarization Loss [dB] -0.18

Receive Antenna Diameter [m] 9.4

Receive Antenna Pointing Error [deg] 0.005

System Noise Temperature [K] 135

Data Rate [bps] 1.00E+08

Bit Error Rate - 1.00E-05

Eb/No for Bit Error Rate [dB] 9.6

Implementation Loss [dB] -2

Outputs

Peak Transmit Antenna Gain [dBi] 16.45

Transmit Antenna Diameter [m] 0.1064

Transmit Antenna Pointing Loss [dB] -10.830

Space Loss [dB] -222.2

Peak Receive Antenna Gain [dBi] 55.34

Receive Antenna Beam Width [deg] 0.279

Receive Antenna Pointing Loss [dB] -0.0038

Eb/No [dB] 14.86

Carrier to Noise Density Ratio [dB-Hz] 94.86

Margin [dB] 3.264

COMM. SATS. TO EARTH

Inputs

Frequency [Hz] 8.00E+09

Transmitter Power [W] 1000

Transmitter Line Loss [dB] -1

Transmit Antenna Beam Width [deg] 2.79E-01

Transmit Antenna Pointing Offset [deg] 2.65E-01

Propagation Path Length [m] 3.800E+08

Propagation/Polarization Loss [dB] -0.18

Receive Antenna Diameter [m] 0.1064

Receive Antenna Pointing Error [deg] 0.005

System Noise Temperature [K] 135

Data Rate [bps] 1.00E+06

Bit Error Rate - 1.00E-05

Eb/No for Bit Error Rate [dB] 9.6

Implementation Loss [dB] -2

Outputs

Peak Transmit Antenna Gain [dBi] 55.38

Transmit Antenna Diameter [m] 9.3985

Transmit Antenna Pointing Loss [dB] -10.830

Space Loss [dB] -222.1

Peak Receive Antenna Gain [dBi] 16.41

Receive Antenna Beam Width [deg] 24.671

Receive Antenna Pointing Loss [dB] 0.0000

Eb/No [dB] 14.97

Carrier to Noise Density Ratio [dB-Hz] 74.97

Margin [dB] 3.372

RECOMMENDATION JUSTIFICATION

ASC Signal 9.4 M Earth Station Antenna

• Already designed and proven, just ship

it!

• Can be placed in an optimal location for

our mission

Communication Satellite antenna must

be fabricated or found elsewhere

Picture from http://ascsignal.com/

REFERENCES http://ascsignal.com/files/satellite/earth_station/9.4mKa-band/PBESA94MEA.C.pdf http://www.nasaspaceflight.com/2013/04/iss-communications-overhaul-boost-scientific-output/ http://amrc.ssec.wisc.edu/meetings/MGS/history.html

http://ascsignal.com/files/satellite/earth_station/9.4mKa-band/PBESA94MEA.C.pdf



http://www.nasaspaceflight.com/2013/04/iss-communications-overhaul-boost-scientific-output/











http://amrc.ssec.wisc.edu/meetings/MGS/history.html

CONTROLS MAO KONISHI

Rover Communication

February 18, 2016

ROVER COMMUNICATION OVERVIEW

Objective: Design antennas for rovers that can communicate to comsat

Reasoning: Mission requires a two-way HD video communication 24/7

Mao Konishi

Science Rover

• Receive control commands

from Earth/hab via comsat

• Commands: 1kbs

Pressurized Rover

• Stream HD videos of crew

to Earth via comsat

• HD video: 100 Mbps

Elevation map of base area

ROVER COMMUNICATION ROVER ANTENNAS

For each rover (UHF, High-Gain Antenna)

Mass (kg): < 1 for both

Power (W): 2, 1

Diameter (m): 0.7698, 1.3125

Recommendation: Pressurized rovers use UHF antenna for main communication, have High-Gain Antenna (HGA) for backup. Uncrewed rover only uses HGA.

Mao Konishi

BACKUP SLIDES LINK BUDGET ANALYSIS: UHF ANTENNA

Mao Konishi

BACKUP SLIDES LINK BUDGET ANALYSIS: HIGH GAIN ANTENNA

Mao Konishi

BACKUP SLIDES

4 communication satellites

• circular, polar orbit (i = 90 degrees)

• separated by Ω = 45 degrees

• 4400 km altitude

• 12 hour period

• Visible 4 hours per period

• Assume receiver can be in

line-of-sight within 128 degrees

COMMUNICATION SATELLITE ORBIT DETERMINATION

Mao Konishi

BACKUP SLIDES ELEVATION ANALYSIS

Mao Konishi

BACKUP SLIDES COMMUNICATION SATELLITE ORBIT DETERMINATION

Mao Konishi

STRUCTURES ARIEL DIMSTON

Unified Rover System

DESIGN AND CAD

Objective Transportation Tanks,

Fuel Cell tanks, Bumper,

Attachment Hard Points, Paint

Scheme, in-hub motors

Reasoning To design a unified

rover system that can accomplish

many different missions

Fig. 3: 10 horsepower in-hub motor within wheel hub

Fig. 2: Fluid Storage Module

Fig. 1: Bumper and Attachment Hard points

SPECIFICATIONS

Fig. 5: Lunar Rover

Configuration Max Side Slope [deg]

Max Slope [deg]

Empty 26 30

Fuel Tanks 23 40

Science Module 20 40

Operation Parameter

Value

Max Speed 30 [kph]

Cross country Speed

20 [kph]

Fluid Storage Module

0.709 [Mg]

FSM Capacity 3.80 [m3]

FSM Displacement

7.85 [m3]

BACKUP SLIDE 1

Modulus of Elasticity 229 [Gpa]

Tensile Yield Strength 3654 [Mpa]

Density 1.77 [g/cm^3]

Fig. 5: Attachment Points – Max Towing Capacity 6,000 [kg]

Carbon Fiber Composite Material Properties

HUMAN FACTORS MICHAEL WALDMANN

Hab assumptions

Internal Power, Mass, Volume

Michael Waldmann

INTERNAL HAB MASS, POWER VOLUME OBJECTIVE: WHAT IS GOING INSIDE OUR HABS

Reasoning: Directly related to IMLEO

Assumptions:

• Did not consider lift-off/inflight orientations

• I am not an architect

• I am not a psych major

Michael Waldmann

POWER, MASS, VOLUME

Figure 1. Sizing and layout example

Michael Waldmann

Michael Waldmann

Everything is work in progress There are excel sheets for every hab type

MISSION DESIGN DOMINIC VITELLO

Adams-1 detailed mission trajectory

THE ADAMS-1 MISSION

• Given the information in the

Mission Arcetecture, is the Adams-1

Mission dynamically possible?

• If so, what is the proposed timeline

for the mission itself?

• What launch dates are ideal for

the Adams series?

Requirement Value Units

Cape Canaveral

Launch

28.46, -80.56

°N, °S

Payload 33.5 Mg

Pre TLI LEO 450 km

Maneuver budget (Δv)

6.928 km/sec

Polar LMO (inclination)

85-90 °

LMO Duration <4 weeks

Deorbit LEO 450 km

TRAJECTORY AND BURNS

Maneuver TLI Lunar Capture

Circularization TEI Earth Capture

Circularization Total

Cost (Δv, km/s) 3.1 0.1 1 0.2 2 .5 6.9

LOX/LH Mass: 130 Mg IMLEO: 163.5 Mg

MISSION DESCRIPTION

• Use Hohmann transfer ballistic trajectories

• Launch January 3rd, 2021 • Time of flight

• 5.1 days • Viable for current proposed launch regime • Do not use plane change for date compensation

Event Date

TLI 1/3/2021

Lunar Capture 1/8/2021

Circularization 1/9/2021

XM Rondezvous 1/9/2021

Docked Adams-1 in LMO 1/10-2/5/2021

XM-Adams-1 separation 2/6/2021

TEI 2/6/2021

Earth Capture 2/11/2021

Circularization 2/11/2021

0

50000

100000

150000

200000

250000

300000

350000

400000

450000

0 5 10 15 20 25 30 35 40 45

Alt

itu

de

(km

)

Elapsed Days (days)

Adams-1 Altitude as a function of Mission Elpased Days

Earth Altitude Moon Altitude

% Script Mission - Lunar Transfer Example % % This script demonstrates how to set up a lunar transfer mission % %---------------------------------------- %---------- Spacecraft %---------------------------------------- %**************************************************************************%************ Create Objects for Use in Mission Sequence ******************%************************************************************************** %-------------------------------------------------------------------------- %----------------- SpaceCraft, Formations, Constellations ----------------- %-------------------------------------------------------------------------- Create Spacecraft Nux; GMAT Nux.DateFormat = UTCGregorian; GMAT Nux.Epoch = '20 Dec 2018 11:29:10.811'; GMAT Nux.CoordinateSystem = EarthMJ2000Eq;

MISSION DESIGN JOSHUA OSTMAN

Lunar Descent:

Elevation Clearance, Incoming Descent Trajectory Requirements

2/4/2016

THE PROBLEM

Objective:

• Ensure landing trajectory is able to clear terrain

Reasoning:

• Generate requirements for incoming trajectory angle and trajectory design

JOSH OSTMAN

RESULTS

From To

5° CW of North 110° CW of North

85° CCW of North 100° CCW of North

125° CW of North 150° CW of North

Restricted Incoming Trajectory Angles

BACKUP SLIDES JOSH OSTMAN

Grayscale map used to generate height map (Provided by Jake Elliot)

-8811 m 9142 m

Elevation Ranges

Corners (lon, lat)

Upper Left 275.90,-72.73

Lower Right 54.91, -79.03

BACKUP SLIDES JOSH OSTMAN

Action Length of Burn ΔV

Descent Orbit Insertion

-- 0.2761 km/s

Braking and Rotation

775 sec 2.183 km/s

Vertical Descent 145 sec .344 km/s

Total -- 2.8031 km/s

BACKUP SLIDES JOSH OSTMAN % Joshua Ostman Mission Design % CODE FOR PLOTTING TRAJECTORY ONTO 3D MOON MAP %clear all; % ===================================== % CONSTANTS % ===================================== rMoon = 1737 % km, radius of the Moon circ = 2*pi*rMoon % km, circumphrence of the Moon % Top left and bottem right corners of map lon1 = 275.90 lat1 = -72.73 lon2 = 54.91 lat2 = -79.03 [arclengthDiag,az] = distance(lat1,lon1,lat2,lon2) % arc length distanceDiag = circ*(arclengthDiag/360) % diagonal length across picture map % =========================================== % CREATING MAP % =========================================== c = imread('elevation.png'); c = rgb2gray(c); [y,x] = size(c); theta = atan(x/y) xDistance = distanceDiag*cos(theta) yDistance = distanceDiag*sin(theta) X = 1:x; Y = 1:y; [yy,xx]=meshgrid(X,Y); i = im2double(c);

% RESCALING MAP i = rescale(i,-8.811,9.142) xx = rescale(xx, 0, xDistance) yy = rescale(yy,0,yDistance) figure; mesh(xx,yy,i); % ======================================================== % PLOTTING TRAJECTORY ON TO MOON % ======================================================== hold on % TRAJECTORY DATA FROM OPTIMAL DESCENT CODE z = zeros(size(x_sol_mass))*(1/1000) x = x_sol_mass*(1/1000) y = y_sol_mass*(1/1000) % ROTATING TRAJECTORY AROUND LANDING SITE for theta = 0:5:360 [x_rotate_loop, z_rotate_loop] = rotate(x,z,max(x), min(z), theta*(pi/180)) plot3(x_rotate_loop-(max(x)-xDistance/2),z_rotate_loop+(yDistance/2),y-1.6298,'ro') end [x_rotated1, z_rotated1] = rotate(x,z,max(x), min(z), 0) [x_rotated2, z_rotated2] = rotate(x,z,max(x), min(z), pi) [x_rotated3, z_rotated3] = rotate(x,z,max(x), min(z), -pi/4) [x_rotated4, z_rotated4] = rotate(x,z,max(x), min(z), pi/4)

BACKUP SLIDES

% PLOTTING TRAJECTORIES ON TO MESH %plot3(x_rotated1-(max(x)-xDistance/2),z_rotated1+(yDistance/2),y-1.6298,'ro') %plot3(x_rotated2-(max(x)-xDistance/2),z_rotated2+(yDistance/2),y-1.6298,'ro') %plot3(x_rotated3-(max(x)-xDistance/2),z_rotated3+(yDistance/2),y-1.6298,'ro') %plot3(x_rotated4-(max(x)-xDistance/2),z_rotated4+(yDistance/2),y-1.6298,'ro') % PLOT DETAILS axis equal xlim([0,xDistance]) ylim([0,yDistance]) colorbar % figure; % imshow(i); % ========================================================== % PLOTTING HEIGHT CLEARANCE % ========================================================== figure newx = x_rotated4-(max(x)-xDistance/2) newz = z_rotated4+(yDistance/2) newy = y-1.6298 newz = round((newz/yDistance)*1609) newx = round((newx/xDistance)*850)

for n = 1:1:length(newx)

if (newz(n)>0)&&(newx(n)>0)

altitude(n) = newy(n)-i(newx(n),newz(n))

else

altitude(n) = 0

end

end

minAltitude(1:length(newx)) = 2

plot(newx, altitude)

hold on

plot(newx, minAltitude, '--')

xlabel('Downrange Distance (km)')

ylabel('Terrain Clearance (km)')

title('Terrain Clearance for Trajectory Incoming at 45\circ CCW of North')

legend('Trajectory Altitude', '2 km Terrain Clearance')

xlim([0,max(newx)])

PROPULSION HAKUSHO CHIN

Ferry to Mars Cycler

• Trajectories & Staging

• Stage specifications

• Volume Analysis

FERRY TO MARS CYCLER STAGING Objective: Mass and volume analysis of the ferry to cycler vehicle

Reasoning: To design the vehicle according to the specification and to validate the design

• Mission Requirements

• Performing Maneuvers

Payload DeltaV-St1 DeltaV-St2 DeltaV-St3 DeltaV-St4 DeltaV-Total

20Mg 2.5km/s 2.1km/s 0.94km/s 0.2km/s 5.74km/s

Illustration by Paul Witsberger Originally from “Guidance Strategy of Hyperbolic Transfer Vehicle” By J.Longuski St1 St2 St3 St4

1,2 3 4 5,6

*delta V found with cooperation with Paul Witsberger

FTMC PROPULSION SPECIFICATION

Stage 1-3: Hydrolox, Isp450s, O/Fratio 5.58, RS-25 SSME Stage 4: Hydrazine, Isp 230s,MR-107

Final result with SLS Block 2 Launch

Recommendation:

This design is able to launch 4 crews + 20Mg payload and to dock with cycler

Stage Propellant specification

Stages M0 M inert M prop Payload Oxidizer tank Vol Fuel tank Vol

1 129.29Mg 10.34Mg 55.94Mg 63.0Mg 7.45m^3 668.24m^3

2 62.83Mg 5.02Mg 23.80Mg 34.0Mg 3.17m^3 284.31m^3

3 33.65Mg 2.69Mg 6.46Mg 24.5Mg 0.86m^3 77.17m^3

4 24.5Mg 1.96Mg 1.09Mg 21.5Mg Monoprop tank volume: 1.07m^3

Payload Total Tankage Vol Payload Vol Vacant Vol

21.5Mg 1042.27m^3 20m^3 537.73m^3

FAIRING PACKAGE

Red: Fairing Blue: Package shell Yellow: Engine Green: LH2 tank Purple:LOX tank White: 20Mg Cargo

CAD Illustration by Hakusho Chin

REFERENCES

[1] Hydrazine Monopropellant engine, Aerojet Rocketdyne, https://www.rocket.com/propulsion-systems/monopropellant-rockets[retrieved 17 Feb 2016]

[2] Orion Spacecraft, NASA.gov, https://www.nasa.gov/exploration/systems/orion/index.html [retrieved 17 Feb 2016]

[3]Hydrazine, Encyclopedia Astronautica, http://www.astronautix.com/props/hydazine.htm [retrieved 17 Feb 2016]

[4] Landau, D. F., and Longuski, J. M. “Guidance Strategy for Hyperbolic Rendezvous.” Journal of Guidance, Control, and Dynamics. (2007): 1209-1213.

[5]RS-25 Engine, Aerojet Rocketdyne, http://www.rocket.com/rs-25-engine [retrieved 17 Feb 2016]

[6] David, L.,Akin, “Mass Estimating Relations,” University of Maryland,

http://spacecraft.ssl.umd.edu/academics/483F09/483F09L13.mass_est/483F09L13.MER.pdf [retrieved 20 Jan 2016]

https://www.rocket.com/propulsion-systems/monopropellant-rockets





https://www.nasa.gov/exploration/systems/orion/index.html

http://www.astronautix.com/props/hydazine.htm

http://www.rocket.com/rs-25-engine





BACKUP SLIDES 01

Dimension List

• SLS Block 2 fairing envelope size: 29.9 x 9.1m, shell size:31.1 x 10m

• Package Shell 20cm thick to envelope fuel tank, engine, and cargo

• RS-25, 4.3m height, 2.4m diameter

• LOX tank sizing are significant, since the volume is very small, should fit between stages

• LH2 tank sizing

There is a gap on upper surface of the LH2 tank to fit the engine of the upper stage

Gap size is same as RS-25 engine size

• Cargo is same dimension as Orion Capsule module

• Thrust to weight ratio: Stage 1, 1.79, Stage 2, 3.69, Stage 3. 6.90

• Assumptions: No pipe, insulation considered

• Density of LH2 70.99, LO2, 1141, Hydrazine 1021, all in kg/m^3

Stages LH2 tank radius LH2 tank height

1 4.35m 11.48m

2 4.35m 5.04m

3 3.5m 2.4m

BACKUP SLIDES 02

Volume reduction plan

Hydrolox to RP-1/LOX use of Merlin D.

Switching 1st stage or 2nd stage into RP-1/LOX MerlinD to conserve volume.

Switching to Merlin D does not satisfy the requirement.

Max Payload 1st stage Merlin D

Max Payload 2nd stage Merlin D

Stage 1 54.6Mg 63.0Mg




PROPULSION ZACHARY RADY

Crewed Ferry Methalox Engine and Tank Sizing

151

ENGINE CONFIGURATION

Objective: Determine Number of Methalox Engines, Propellant/Tank Mass/Volume

Reasoning: Propulsion Mass/Volume Numbers for Crewed Ferry

152 Zachary Rady Methalox Engine based on designs by ANDREW CULL

Number of Engines:7

G Force = 7 g’s

Thrust Per Engine = 118421 N

Total Thrust = 828947 N Assuming Thrust Required to escape Mars Atmosphere

PROPELLANT TANKS

Total Mass: 11.428 Mg

Total Volume: 9.1515 m3

153 Zachary Rady

Methane Tank

LOX Tank

Mass 3.5428 Mg 0.8857 Mg

Volume 8.3753 m3 0.7762 m3

Total Propellant Mass

4.4285 Mg

BACKUP SLIDE 1

% AAE 450

% Zachary Rady, Andrew Cull

% Crewed_Farry_Engine_Calculations

clear;clc

% Constants

dV = 1870; %m/s

ISP = 389; %s %ISP from Centaur RL10 engine

% minert=1732.136; %kg

g0 = 9.80665; %m/s^2

mpay = 5000; %Kg

IMF=0.05;% inert mass fraction

OF=4; %Fuel Ratio

DMeth=423; %kg/m^3 Density of L Mehtane

DLOX=1141; %kg/M^3 Density of LOX

Dthroat=12*0.0254; % m Throat diameter (first value in in)

AR=165; %Expansion Ratio

gma=3.8; %Mars Gravity m/s

gmo=1.6; %Moon Gravity m/s

Pe=0.05212*10^5; %Pa (first value in bar)

Ve=1793.4; %m/s

TW=180/g0; %thrust to weight ratio based on Merlin 1D

MaxG=16; %Max g force for humans

MATLAB CODE 1

154 Zachary Rady

%Calculations

%% Initial Mass Calc

c = ISP*g0;

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

PMF = MR-IMF; % Propellent mass fraction

Minitial = mpay/PMF;

Mprop=Minitial*(1-MR);

Minert=Minitial-Mprop-mpay;

%%Num Thruster Calc

At=pi()*(Dthroat/2)^2; Ae=At*AR;

Fte=Pe*Ae/(1-Ve/(ISP*g0));

W=Minitial*gma;

TWe=Fte/W; %Thrust to weight of 1 engine

Gv1e=((Fte-W)/Minitial)/g0; %G's on vehicle for 1 engine

Fnma=TW*Minitial*gma;

Gv=((Fnma-W)/Minitial)/g0; %G's on vehicle

NumEnginesneeded=Fnma/Fte;

NumEnginesi=NumEnginesneeded+(1-mod(NumEnginesneeded,1));

Fe=Fte*NumEnginesi;

Gvf=((Fe-W)/Minitial)/g0; %G's on vehicle final

BACKUP SLIDE 2

%%Mass Calc w/num Thruster

NumEngines2=NumEnginesi;

Minert=Minert*NumEngines2*.75; %Factor based on assumption of how much of base inert mass is related to 1 engine

Mprop=(Minert+mpay)/MR-Minert-mpay;

Minitial=Mprop+Minert+mpay;

W2=Minitial*gma;

TWe2=Fte/W2; %Thrust to weight of 1 engine

Gv1e2=((Fte-W2)/Minitial)/g0; %G's on vehicle for 1 engine

Fnma2=TW*Minitial*gma;

Gv2=((Fnma2-W2)/Minitial)/g0; %G's on vehicle

NumEnginesneeded2=Fnma2/Fte;

NumEngines2=NumEnginesneeded2+(1-mod(NumEnginesneeded2,1));

Fe2=Fte*NumEngines2;

Gvf2=((Fe2-W2)/Minitial)/g0; %G's on vehicle final

NumEnginesi=NumEngines2;

if NumEngines2 > NumEnginesi

Minert=Minert*NumEngines2*.75; %Factor based on assumption of how much of base inert mass is related to 1 engine

Mprop=(Minert+mpay)/MR-Minert-mpay;

Minitial=Mprop+Minert+mpay;

W2=Minitial*gma;

TWe2=Fte/W2; %Thrust to weight of 1 engine

Gv1e2=((Fte-W2)/Minitial)/g0; %G's on vehicle for 1 engine

Fnma2=TW*Minitial*gma;

Gv2=((Fnma2-W2)/Minitial)/g0; %G's on vehicle

NumEnginesneeded2=Fnma2/Fte;

NumEngines2=NumEnginesneeded2+(1-mod(NumEnginesneeded2,1));

Fe2=Fte*NumEngines2;

Gvf2=((Fe2-W2)/Minitial)/g0; %G's on vehicle final

NumEnginesi=NumEngines2;

end

MATLAB CODE 2

155 Zachary Rady

%%Tank Sizing

MLOX=Mprop/(OF+1);

Mmeth=Mprop-MLOX;

VLOXT=MLOX/DLOX; %Volume of Lox tank

VmethT=Mmeth/DMeth; %Volume of methane tank

REFERENCES

156

"Merlin (Rocket Engine Family)." Wikipedia. Wikimedia Foundation, n.d. Web. 10 Feb. 2016. Sutton, G. P., & Oscar, B. (2010). Rocket Propulsion Elements. John Wiley & Sons Inc. Sutton, G. P., & Oscar, B. (2010). Rocket Propulsion Elements. John Wiley & Sones Inc. Cengel, Y. A., Cimbala, J. M., & Turner, R. H. (2012). Fundimentals of Thermal-Fluid Sciences. McGraw Hill.

Zachary Rady

PROPULSION CLAIRE ALEXANDER

• Adams -3 Science Probe Missions

• Recommended Propulsion System

157 Claire Alexander

SCIENCE PROBES ADAMS-3 (2022)

Propulsion System Requirements:

• Land science probes safely on Lunar surface

• Recommend efficient use of mass and volume

• Delta V = 2.183 km/s

• Must fit in SLS 1B fairing: 8.4 m diameter, 19.1 m height

• 4 science probes containing seismometer and other equipment

• 4 separate landing sites

• All 4 probes Launched from LEO to CLO using SLS 1B upper stage

• Estimated total of 100 kg of science instruments


Benefits of Hypergolic Propellant:

• Easy/small storage of propellants

• Easy pulsed

• Can be throttled

• Isp and Thrust closely resembles needs

SCIENCE PROBES RECOMMENDED PROPULSION SYSTEM

Propulsion Recommendation:

Aerojet Rocketdyne HiPAT 445N

Mass: 110 kg

Power: 46 Watts

Volume: 0.40 m^3


Mass: Mass of Propellant: 104.6 kg Mass of Engine: 5.4 kg Payload Mass: 100.0 kg Total Mass: 210 kg

Volume: Volume of Fuel: 0.06 m^3 Volume of Oxidizer: 0.04 m^3 Engine Volume: 0.30 m^3 Total Volume: 0.40 m^3

BACKUP SLIDES

160

REFERENCES

Braeunig, A. Robert. “Rocket & Space Technology”. Rocket Propellants.

http://www.braeunig.us/space/propel.htm. [Retrieved 13 February 2016].

“Aerojet Rocketdyne Capabilities”. Bipropellant Fact Sheet.

http://www.rocket.com/files/aerojet/documents/Capabilities/PDFs/Bipropellant%20Data%2

0Sheets.pdf. [Retrieved 13 February 2016].

Liou, Larry. “NASA Archive”. Advanced Chemical Propulsion for Science Missions.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080014199.pdf. [Retrieved 13

February 2016].

http://www.braeunig.us/space/propel.htm

http://www.rocket.com/files/aerojet/documents/Capabilities/PDFs/Bipropellant Data Sheets.pdf

http://www.rocket.com/files/aerojet/documents/Capabilities/PDFs/Bipropellant Data Sheets.pdf


BACKUP SLIDES

161

CODE/CALCULATIONS

% Landing on Lunar Surface for Science Probes

% Author: Claire Alexander

% Based on code written by: Mason Buckman

clear all

close all

clc

% Known Values

deltav = 2183; % (km/s) from LLO to surface

g = 9.80655; % (m/s^2) for Earth

x = [1:1:9];

Isp = [294 280 315.5 323 329 303 327 333 293]; % Isp of engines (sec)

mi = [0.454 2 4.31 5.44 5.44 4.53 7.3 5.4 6.8]; % mass of engines (kg)

p = [5 36 46 46 46 46 45 45 70]; % power required to operate the valve (W)

t = [22 111 490 445 445 890 890 623 4000];

mg = 9.80655/6; % gravitational acceleration of the Moon (m/s^2)

mpay = 100; % (kg)

% Calculations to narrow down engine options

MR = exp(deltav./(g.*Isp));

mprop = MR.*(mi+mpay)-mpay-mi;

mfull = mprop+mpay+mi;

mempty = mi+mpay;

T = mfull*mg;

set(gcf,'color','w');

subplot(3,1,1)

plot(x,mfull)

xlabel('Engine number')

ylabel('Mass (kg)')

title('Total Mass of Probe')

subplot(3,1,2)

plot(x,p)


ylabel('Power (W)')

title('Power Required')

subplot(3,1,3)

plot(x,T)

hold on

plot(x,t,'r')


ylabel('Thrust (N)')

title('Comparrison of Thrust

Provded and Required by Engine')

legend('Required','Provided')

Ethrust = t(4);

Emass = mprop(4);

OF = 1; % Choosen Engine O/F ratio

fd = 880; % Density of Fuel (kg/m^3)

od = 1440; % Density of Oxidizer (kg/m^3)

mfuel = Emass/2; % Mass of fuel needed (kg)

mox = Emass/2; % Mass of oxidizer needed (kg)

vfuel = mfuel/fd; % Volume required for fuel (m^3)

vox = mox/od; % Volume required for oxidizer

(m^3)

Ev = (pi*((26.1+2.47)*0.0254)*(14.25*.0254)^2);

project manager trevor jahn - purdue university€¦ · feb 14 – 20 (2.16.2016) three copies of 1...

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