Section I PresentationsFebruary 13, 2020

Nick Rich February 13, 2020

Project ManagerProject Schedule and Updates, Storyboard

Schedule, Project Updates, PDR Overview

Schedule

February 27 - PDR

- No longer landing on Mars surface vertically with chemical propulsion. We will use tether rendezvous and spindown for landing on Mars.

February 20 – PDR Slides Assigned

Project Updates

- Mass Driver will be used for Mars launch, tether sling will be used to catch taxis.

- Crew will be extracted from tether slings on Luna and Mars

- Due to the change in landing approach on Mars, significant propellant reduction in taxi. Expect taxi to get smaller and/or change shape. Number of taxis could change as well.

- Low ΔV from the ED tether will likely make missions from Earth to Mars system infeasible.

- Heat transfer on Mars re-entry will drive a change in taxi shape.

- Landing pad is gone but is being replaced by a “track” to seat the taxis upon spindown.

Risk Assessment

February 13 – On track!

APM – List of key events compiled and being refined as mission requirements change.

Updated Storyboard: OverviewMission Profile Martian Mission Requirements

Earth - Lunar Missions Requirements

Phobos

Phobos

To Mars

To/From EarthSystem

Landing + Takeoff

Cycler Departure/ArrivalRendezvous Only

Mission Start

Updated Storyboard: Earth-Luna SystemMission Profile Tether Sling – Luna and Mars

Taxi

Tether Sling Hub

Elevator

Path from Taxi tohabitation unit

Tether

ED Tether

Updated Storyboard: Mars-Phobos SystemMission Profile Tether Sling Spindown - Mars

Phobos

Tether Slings

Mass Driver Olympus MonsTrack to “catch” taxison spindown

Sodiq AdenijiFebruary 13, 2020

Discipline: CADVehicle & Systems Group: Tether Sling & ED Tether

Tether SystemsObjective: To provide a visualization of the Tether Sling & ED-Tether

Requirements:• 2D Drawings• 3D Models• Storyboard (Animations)

Constraints (ED-Tether & Tether Sling):• Model development is dependent on work from (MD, Struct, P&T)

Need to Finalize & Develop:• Length of Tether (Luna, Phobos, Mars)• Overall Height of the system and Motor Size (Luna, Phobos, Mars)• Docking Method• Animation Sequence (May be best to start after a design freeze)

Tether Sling (Current Model)

Tether Sling (Isometric)Not to Scale

Tether Sling (Top)Not to Scale

Generally Applies to both, Luna, Phobos, and Mars systems, but can be modified to

meet the requirement of each system.

ED-Tether still in development, design will be dictated by Mission Design requirements.

Hanson-Lee Harjono February 13, 2020

Discipline: CADVehicle: Communication Satellites

Satellite RequirementsCommunications [1]Geostationary:• 3 Radio Frequency antennas• 50 cm diameter laser

communications system (LCS)Lagrange Points:• 50 cm diameter LCS• 160 cm diameter LCS

Mars orbit:• 3 Radio Frequency antennas• 160 cm diameter LCS

LCS uses 500 W

Controls• LCS: rotate 360º & pitch 8º• 8 electric propulsion

thrustersPower & Thermal• Solar panels: 350 W/m²

Propulsion [3]• Electric propulsion: 2000 W

per engine

Structures [2]

[1] : Nick Oetting & Eric Smith, Communications Team[2]: Rachel Roth, Structures Team[3]: Chuhao Deng, Propulsion Team

Satellite Model

Megan BrownFebruary 13th, 2020

Area: Communications

Systems: Ground StationCycler RF Antenna

THE PROBLEM

• Determine the size of the cycler’s RF antenna based on the analysis of link budgets at different antenna diameters.

• Determine the power required to communicate from the ground station to the cycler at the Moon using RF.

• From the above value, determine whether RF or lasers should be used to communicate from the ground station to the cycler.

THE SOLUTION

• 3 meter RF antenna will be used on the cycler.• Validation: 3.7 meter antenna was used on Voyager

missions [1]

• Validation: Average of 5 Watts utilized for ISS to communicate with Earth ground station [2]

Antenna Diameter Power Required10 meters 0.45 Watts

5 meters 1.80 Watts

3 meters 4.03 Watts

2 meters 9.00 Watts

Table 1: Cycler power required at different antenna diameters

Location Antenna Diameter

Power Required (Moon)

Power Required (GEO) [3]

University of Alaska, Fairbanks

9 meters 353.18 Watts 3.5 Watts

Morehead State University

21 meters 70.46 Watts 0.7 Watts

Table 2: Ground antenna power required to reach cycler at Moon and in GEO

• RF will be used to communicate to relay satellites in GEO, not directly to the Moon.

Alex MooreJanuary 24, 2020

CommunicationsTaxi

Problem• How do we communicate during launch/atmospheric

flight?• External antenna would burn up in the atmosphere.• Where do we mount the antenna?

• Requirements• 4 dB gain margin• < 1kW transmission power

Solution• External Antenna mounted to taxi hull.

• Antenna would have to be mounted after launch.• Internal Antenna mounted inside taxi nose, surrounded

by radio transparent material [2].

Image by Alex Moore

Parameter External Antenna Internal Antenna

Mass [1] 20 kg 1.25 kg

Power [3][4] 3.94 W 63.00 W

Diameter 2.0 m 0.5 m

BREAKResume at 2:10

Alexander James ChapaMonth Day, 2020

Vehicle and Systems Group: CyclerDiscipline: Cycler Orbital Controller

The Problem: Estimating the required propellant needed to account for external forces.Assumptions:• Cycler always has largest side exposed to the

source of force• Using 1000m² as surface area always exposed

to the sunGoals:• Find external force that would affect the space

craft.• Calculate the needed force to counter the

affects and fuel consumed in the process of correction

EstimationAverage Net Force From Each Satellite:

Satellite Force (N)Mercury 8.668*10-4

Venus 7.3*10-3

Earth 1.049*10-4

Moon 3.565*10-6

Mars 9.647*10-5

Jupiter 1.072*10-4

Saturn 1.723*10-5

Uranus 8.373*10-8

Neptune 2.316*10-8

Riley FranklinJanuary 23rd, 2020

Discipline: ControlsSystems: Cycler

Control Moment Gyroscope Sizing

Problem: Need to find the amount of angular momentum that will need to be stored in our CMGs. Requirements:

• Store angular momentum in any direction.• Store enough for length of the mission.

Need to Find:• Moment of inertia of the cycler• Layout of the CMG’s• Size and weight of the CMG’s

https://www.researchgate.net/figure/FIGURE-S1-Control-moment-gyroscope-CMG-A-CMG-consists-of-a-a-rotor-spinning-at-a_fig1_224594118

Drawn By: Sidharth Prasad

Results

Conclusions and Moving Forward:• Required flywheel for the CGMs is

most likely unreasonable• Look into means of desaturating the

CMGs without using Propellant• Investigate validity of assumptions

Moment Of Inertia of Cycler

(kg*m2)

Angular Momentum of

Cycler(kg*m2/s)

Radius of Flywheel (m)

Mass of Flywheel

(Mg)

Angle of Pyramid Configuration

(degrees)

[5.75 * 109,8.16 * 1011,8.10 * 1011]

[1.28 * 1011,-9.002 * 108,

0]

7.5 1.06 * 103 53.1

Above visual made by Eli Sitchin

Sarah CulpFebruary 13, 2020

Human FactorsCycler and Taxi: Waste Management

The Problem: How do we recycle our waste?70 people will produce 18.5 Mg of trash and 2.15 Mg of feces in 8 months

• Trash includes: clothing, hygiene items, life support supplies, etc.

• Current methods of trash disposal include storing it short term and then disposing of it in a separate spacecraft

• Burns in Earth’s atmosphere or returns to Earth• Not a plausible solution for our mission

• The goal is to repurpose this waste in a recyclable, useful way

0.128 kg/day/person of feces produced

1.1 kg/day/person of trash produced0.03 m3/day/person of

trash produced

X 70 people

The Solution: PHB (Feces Waste) and OSCAR (Trash)

Adapted from University of Calgary’s Astroplastic Project

Adapted from NASA’s Trash-to-Gas Initiative

Feces to Plastic (PHB) System

Taxi Infrastructure + Biomass (4 days)

Cycler Infrastructure + Biomass (8 months)

Mass (Mg) 0.40 4.32

Volume (m3) 5.76 47.2

Power (kW) 10.4 83.2

Trash to Gas System

Taxi Trash Generation (4 days)

Cycler Infrastructure + Trash (8 months)

Mass (Mg) 0.08 6.14

Volume (m3) 0.79 51.7

Power (kW) 0 8.75

Feces – converted to greenhouse fertilizer, buildable plastic (polyhydroxybutyrate)

Trash – converted to H2, H2O, char, vented gases

David FoxFeb. 13, 2020

Discipline: Human FactorsVehicle: Cycler

Topic: Storage, entertainment, interior design

1

The Problem(s)

Storage:- Simple/uniform ways to store luggage

Entertainment:- Crew will need ways to stimulate their senses

Interior design:- Psychological stressors [1]

- Crowding issues- Lack of privacy- Limited communication

2

Current Solution

Storage:- 2 double CTBs- 1 carry-on suitcase

Entertainment:- VR Oculus Quest- TVs

Interior design: [1]

- Have common recreational/dining area- Private crew quarters- Less about volume, more about layout

Mass (kg) Power (kW) Volume (m3)

CTBs (140) 3156[2] - 15[2]

Carry-on (70) 700 - 3.2

Storage Totals 3856 - 18.2

VR headsets (16) 9.1 0.24 -

TVs (4) 55.2 0.202[3] 0.752

Entertainment Totals 64.3 0.442 0.752

Next steps:- Floor surface area for different rooms- Layout of interior- Incorporate storage and entertainmentsystems

3

Kait HauberFebruary 13, 2020

Discipline: Human FactorsVehicle/System: Cycler

Topic: Lighting, Bathing, Psychology, Exercise

The Problem: Exercise, Hygiene, and Light

- Light/ UV Exposure and Vitamin D- Bone health- Psychological- Functional

- Hygiene and Toilet Availability- Psychological- Sanitation- Functional

- Urine Recycling and Management- Byproducts not used by other

systems

- Exercise- Psychological- Not essential for maintenance

- Clothes Washing- How often?

Solution

Space Requirements:- 10 showers at 1m2

- Twin bed

Power Requirements:- 500 W for lighting in essential

areas in cycler- 1.2 kW for tankless water heater

- LEDs used for all lighting in cycler

BREAKResume at 2:44

Jennifer Bergeson

February 13, 2020

Discipline: Mission DesignVehicles and Systems: Cycler Orbit

The Problem: Determine Burn Times and Eclipse Times for the CyclerRequirements:• Stay within predetermined thrust to weight ratio• Investigate all inner planetary bodies• Cycler follows predetermined trajectory

Assumptions:• Sun emits energy uniformly• Planets are spherical• Cycler orbit is locally circular• Mercury and Venus are far from cycler

Need to Determine:• Maximum burn time before refueling• Maximum eclipse time due to each planet• Relevance of Mercury and Venus

Necessary ParametersThrust to weight (N/Mg)

.01

Solar surface flux[1]

(MW/m2)63

θ

s/c

Numerical Results

Conclusions:• Inner planet eclipses are irrelevant• Cycler needs to be capable of 8 days of burn

Next steps:• More detailed eclipse times• Acceptable trajectory deviations

Solar Power Loss Due to Inner PlanetsMercury (%) 0.000624

Venus (%) 0.00756

Maximum Eclipse Time Due to Each Planet

Mercury (min) 9.45

Venus (min) 17.6

Earth (min) 7.14

Mars (min) 4.69

Jordan CuellarMonth Day, 2020

Misson Design for Phobos Tether SlingArrival and Departure of Objects from the

Tether Sling

Taxi Departure from Tether Sling on Phobos

• Need to define an orbit to reach the cycler.

• Calculate and initial ΔV for tether calculations.

• Calculate power needed for the ΔV

Calculations and Process

• Hyberbolic orbit to reach cycler velocity from Phobos.

• 𝑉𝑉∞ = 4.3 𝑘𝑘𝑘𝑘𝑠𝑠→ ∆𝑉𝑉 = 3.12 𝑘𝑘𝑘𝑘

𝑠𝑠

• 𝑉𝑉∞5.7 𝑘𝑘𝑘𝑘𝑠𝑠→ ∆𝑉𝑉 = 4.31 𝑘𝑘𝑘𝑘

𝑠𝑠• 𝑃𝑃𝑎𝑎𝑎𝑎𝑎𝑎 = 2.06 𝐺𝐺𝐺𝐺,𝑃𝑃𝑀𝑀𝑎𝑎𝑀𝑀 = 2.844 𝐺𝐺𝐺𝐺• Assumes 6 g’s on the taxi with a

mass of 11.2 Mg

Nicolas Martinez CrucesFebruary 13, 2020

Mission DesignMass Driver

Problem & Assumptions

Requirements• Mars to Phobos trajectory definition• 4 days on taxi• Minimize Δ𝑉𝑉

Assumptions• No orbital perturbations• Phobos orbit has 0 eccentricity• Instant Δ𝑉𝑉 boost

Need to Determine• Trajectory of Taxi from Mars to Phobos• Olympus Mons has high latitude• Must optimize Δ𝑉𝑉 as much as possible

Solutions & LimitationsΔ𝑉𝑉 𝑘𝑘𝑘𝑘

𝑠𝑠Time (hrs)

Launch 4.47 6.17

Burn 1 0.286 8.52

Burn 2 0.375 RNDVZ

Total 0.661 14.69

Limitations• Not fully optimized• No perturbations• Δ𝑉𝑉 thrust might differ with RCS Source [1] & [2]

Suhas AnandFebruary 13, 2020

Power and ThermalCycler

Thermal SystemsThe Problem - Sizing the thermal system for the CyclerAssumptions - 70% of power used will be converted into heat

Total Heat Generated: 1.22 GW

Human Factors: 70kW1

Propulsion: 1MW2Flight Controllers: 4kW3

Solar Power: 1.74GW4

[1] AAE 450: Human Factors Team[2] AAE 450: Propulsion: Griffin Pfaff[3] Burkey, Ronald (2009-08-21). "Virtual AGC — AGS — LVDC — Gemini: Launch Vehicle Digital Computer (LVDC): Saturn IB and Saturn V Rockets".[4] AAE 450: Power and Thermal: Jacob Nunez-Kearny

Mass and Power Sizing

Heat Exchangers

Coldplates

Radiators

Pumps

Heat PumpsFreon 215

[5] Hanford, Anthony J. and Michael K. Ewert. “Advanced Active Thermal Control Systems Architecture Study.” (1996).[6] Larson, Wiley J., and Linda K. Pranke. Human Spaceflight: Mission Analysis and Design. McGraw-Hill, 2007.

Passive Thermal Control

Total Sizing6

Total Mass 1636.2 Mg

Total Power 2435.7 kW

Total Volume 5.7674 x 104 m3

Vincent Bartels February 13, 2020

Power & Thermal Group LeadSurface (Phobos, Moon, Mars) Tethers

Problem: That’s a lot of power

Focus: Given power and mission requirements, determine the optimal motor & corresponding method of powering it

Objectives:- Analyze and compare various types of

potential motors- Strengths, weaknesses & efficiency

- Find solar panels which provide most efficient power production for given environment

- Determine surface area of panels to provide requisite power Fig 1: How motor power affects spin-up time

Created by Grace Ness (MD)

Motor Selection & Solar Panel SizingMotor Choices Examined

AC DC

InductionSynchronous

HTS

BrushedBrushless

HTS

HTS: High Temp Superconductor- 99.3% efficiency- Intended for high power

applications

Array Size (km²) based on GW reqSolar Array Efficiency 20 40 60 80 100

Top Production [2][3][4] 0.32 169.90 339.81 509.71 679.61 849.52SBT [1] 0.35 155.34 310.68 466.02 621.36 776.70Silicon (traditional) 0.169 321.71 643.42 965.13 1286.84 1608.55

Solar Cell Selection & Sizing

What’s next?- Determine voltage, current draw & requisite wiring- Look into alternate power sources- Develop system for heat dissipation

BREAKResume at 3:20

Yashowardhan GuptaFebruary 13, 2020

Discipline: Power and ThermalVehicle & Systems: Communication Satellites (Thermal)

Problem: TCS for GEO satellites Earth/MarsRequirements: Determining the thermal subsystem for:o Main Body Temperature Controlo Power Systemo Attitude and Control Systemo Solar Arrayso Antennas + Communicationo PropulsionAssumptions: 1. Half of the body is illuminated in the hot cases2. No effect from Earth/Mars’ Albedo and IR [1]3. Circular OrbitsConstraints: o Satellite Dimensions (3m x 3m x 5.5m)

System Power Attitude Control Solar Arrays Antennas Propulsion Main Body

Temperature Parameters

(ºC)-5 to +30 -30 to +30 -85 to +135 -90 to 90 +5 to +40 -10 to +45

5.5 m

3 m

3 m

[1],[2]

[3]

[1]

Initial thermal systems to be used for each subsystemSubsystem Methods for TCS

Main Body MLI [Multi Layer Insulation] Blanket (Passive)High Power dissipating systems attached to the body (Passive)

Power System MLI + Surface Paint (Passive)Radiators (maintain maximum)Thermostatic Heaters (maintain minimum)

Attitude and Control System

MLI Small Radiator

Solar Arrays High emittance paint Low conductive spacers to isolate support structure

Antennas + Communication

MLI + Surface Paint (Reflectors and Polarizer

Propulsion Isolate the system using low-conductivity stand-offs and attachment fittingsHeat Shields and MLI

Parameter Value (±20%)Mass 85 kg (5% of total)

Power 165 W (5.5% of total)

Volume Varies due to surfaces

Outer Body Condition

Min. Operating Temp

Min. Operating Temp

Max. Operating Temp

Max. Operating Temp

Perihelion Aphelion Perihelion Aphelion

⍺/ε 0.19 0.21 0.41 0.44

[2][4]

Chuhao DengFeb 13, 2020

PropulsionCommunication Satellites

End-of-Life Plan

Requirements:• Enough propellant for 100 years operation• End-of-life plan is achievable and efficient

Assumptions:• T = 0.01 𝑁𝑁, ⍺ = 0.01 𝑘𝑘𝑎𝑎

𝑊𝑊, η = 0.5, g = 9.8 𝑘𝑘

𝑠𝑠2• Total working time is 1 year within 100 years operation• At most 3 of the 8 thrusters work at the same time

Solutions:• Transfer satellites to another orbit and fall into atmosphere (Figure(1))

Analysis Process:

Make assumed values

Plug in equations and get 𝑘𝑘𝑢𝑢

𝑘𝑘0

Get the optimal Isp and find the

corresponding power

Calculate 𝑚𝑚𝑝𝑝𝑝𝑝, 𝑚𝑚𝑝𝑝𝑝𝑝𝑝𝑝, 𝑚𝑚𝑝𝑝 and 𝑚𝑚0

T: Thrust ⍺: Power Supply Specific Mass 𝑚𝑚𝑝𝑝: Mass of Useful Payloads η: Thruster Efficiency g: Gravitational Acceleration on Earth 𝑚𝑚𝑝𝑝𝑝𝑝: Mass of Propellant𝑚𝑚𝑝𝑝𝑝𝑝𝑝𝑝: Mass of Power Supply 𝑚𝑚0: Initial Mass

Figure (1)

Results

ΔV (m/s) Mass of Propellant (kg) Mass of Power Units (kg) Mass of Useful Payloads (kg) Total Mass of the Satellite (kg)2000 5.72 5.52 149 1611000 5.67 5.57 307 318

Kristen FleherFebruary 13, 2020

PropulsionElectrodynamic Tether

Electrodynamic Tether Re-orbit

After boosting the taxi the tether partially deorbits

The larger the mass ratio, the smaller the deorbit, but the more force required to reset the system

Can use electrodynamic tethers within the large momentum bank to boost the orbit back

Diagram from Pierre Venzin

Diagram from Joe Tiberi

Boost Orbit Using Electrodynamic TethersUsing mass ratio of 50 for the system [1]

System can scale with the increasing massEstimating taxi mass to be 100 Mg, can be scaled up too

Changing from a 300 km orbit to a 1000 km orbitLarger than the probable orbit decay

Copper cablesCan apply 770A of current

[1] Pierre Venzin, [2] Joe Tiberi

Tethers: 3,000 x 100m longPower Required: 8.02 MW

Tether Power Requirement: ~170MW [2] Voltage Required: 10.4 kVTime: 2.8 days

Equivalent propellant savings: 435 Mg

Adam BrewerJanuary 13, 2020

Discipline: StructuresVehicle: Cycler

Superstructure Design & Sizing

Problem: Design cycler superstructure to meet all criteria and incorporate all necessary systems.

Requirements:• Thrusters: 10 X3 Hall thrusters must be spaced evenly about CoM and fire

along all 3 axes• Docking: Allow 3 taxi vehicles to dock simultaneously• Communication: Antenna must point towards Earth at all times• Habitation: Habitation modules must rotate around superstructure and have

transport and support from the superstructure• Power: Solar panels must be exposed to the sun• Must be protected from space debris

Superstructure Design & Sizing

• Cylinder with 10 m diameter, 50 m height• Sections 4 and 6 rotate at a constant

angular velocity to induce artificial gravity• Sections 1 and 9 rotate so solar panels

are normal to incoming sunlight• All other sections do not rotate; control

used to keep antenna pointed at Earth• Whipple shield implemented to prevent

debris from breaching the hull• Aluminum 2219-T6 used• 3.1 mm outer wall, 42.2 mm inner wall

• Total volume: 3927 m3

• Total hull mass: 208.7 Mg

Eric Eagon February 12, 2020

Structures TeamCycler Vehicle Design

Cycler Habitation Module StructureThe Problem: Habitation module must be pressurized and protect from external threats

Requirements:• Protect against micrometeors• Protect against solar radiation

Assumptions/Constraints:• Vehicle radius is 400 m• Floor space required is 2550 m2

Needs to Determine:• Empty mass of habitation module Image credit: Eli Sitchin (CAD)

Cycler Habitation Module StructureFurther Assumptions & Decisions:

• Pressurized interior modeled as a thin cylinder

• Safety factor of 10 was applied• Interior structures were not factored

into these calculations

Cycler Habitation Mass & Volume

Cross Sectional Radius [m] 4Wall Thickness [m] [1] 0.1Mass [Mg] 2381

Pressurized Habitable Volume [m3] 8043

Conclusions:• Initial empty mass estimates were too high and based on scaling up the ISS

[3]

• Including interior structure will significantly increase this mass estimate• Habitable volume still exceeds requirements which leaves room for

community spaces

[1] Walter Manuel. Radiation shielding requirement with T6061 aluminum.

Backup SlidesPower:Only 4 thrusters will fire at the same time, so at most 8000 W of electric power will be used at any time for electric propulsionThe LCS uses 500 W and the other components use about 120 WTherefore, geostationary satellites use 8000+500+120=8620 WLagrange point satellites use 8000+(2*500)+120 = 9120 WAnd the Mars satellites use 8000+500+120 = 8620 WSolar panels generate 350 W/m², soGSC & Mars: 8620/350 = 24.63 m² solar panelsLagrange Point: 9120/350 = 26.06 m² solar panels

Moment of Inertia

Optics demonstration

Made in the ray optics simulator at https://ricktu288.github.io/ray-optics/simulator/This shows that the signals received by the telescope can be narrowed to a smaller beam for the laser transceiver

Maximum distance between Mars and Earth/Lagrange Points

This animation was created in MATLAB in an attempt to find the maximum distance that Mars is from the satellites in GEO or at L4/L5 at any point in time. This is a gross approximation, as evidenced by the circular orbits, and the Mission Design team’s calculations are favored over this. With this distance, the amount of power needed for the communications signals could be found using power = 1/distance^2

Megan Brown Backup SlidesFebruary 13th, 2020

Area: Communications

Systems: Ground StationCycler RF Antenna

Figure by Megan Brown

Figure by Megan Brown

Alex MooreJanuary 24, 2020

CommunicationsTaxi

Backup Slides

External Antenna Link Budget

Internal Antenna Link Budget

References1. AAE 450 Spring 2015 Project Aldrin-Purdue Report pg. 50-512. “Radome,” Wikipedia Available:

https://en.wikipedia.org/wiki/Radome#Radar_dishes.3. “Antenna Formulas,” University of Iowa, Iowa City, Iowa.

[http://user.engineering.uiowa.edu/~ece195/2006/docs/AntennaFormulas.pdf. Accessed 1/22/2020.]

4. “Link Budget,” Wikipedia. [https://en.wikipedia.org/wiki/Link_budget. Accessed 1/22/2020.]

Citations

• Currie, B. J., “Control of a Spacecraft Using Mixed Momentum Exchange Devices.”

• Nagabhushan, V., “Development of Control Moment Gyroscopes for Attitude Control of Small Sattelites.”

• “List of moments of inertia,” Wikipedia Available: https://en.wikipedia.org/wiki/List_of_moments_of_inertia.

• “Control Moment Gyro (CMG) Sizing and Cluster Configuration ...” Available: https://www.researchgate.net/publication/311665872_Control_Moment_Gyro_CMG_Sizing_and_Cluster_Configuration_Selection_for_Agile_Spacecraft.

Important Equations

• Torque exerted by CMG= Gimbal Rotation Rate x (Iflywheel * ωflywheel)

• Maximum Angular Momentum Storage = 4 * sin(Pyramid Angle) * Angular Momentum of Flywheel

Code• %%%%Moment of Inertia of the Cycler•• b = 52;%Length of the habitation modules in m• habwidth = 6+2.4;%width of one habitation module• lengthcenterhab = 36.3;%length of the central hab connector• habheight = 2.5;%height from ceiing to floor• widthcenterhab = 10+2.4;%Width of central hab in meters• d = 2 * habwidth +lengthcenterhab;• h = lengthcenterhab;• t = widthcenterhab;• s = habwidth;• l = 400;%length of support in meters• g = 9.81%acceleration in hab modules• Volumehab =(habwidth)* 2.5 *(b) -(habwidth-2.4)* 2.5 *(b-2.4);• %Moment of inertia of the Habitation Unit after parallel axis theorem

Code Continued• Volumehabcenter = (widthcenterhab)*(lengthcenterhab)*(2.5) -

(widthcenterhab-2.4)*(lengthcenterhab)*(2.5);•• %moments of inertia for the habitation unit as a cross section of an I

beam• IyyI = (b*d^3- h^3*(b-t))/12;• IzzI = (2*s*b^3+h*t^3)/12;• IxxI = IzzI+IyyI;•• %moment of inertia of the habitation Unit before parallel axis theorom• Ibeforepa = [IxxI 0 0; 0 IyyI 0;0 0 IzzI];•• densitycyc = 2710;%kg/m^3 desnity of aluminum• MassI = (Volumehab*2+Volumehabcenter)*(densitycyc);• Ihab = Ibeforepa +[0 0 0;0 400^2 0;0 0 400^2]*MassI;

Code Continued• Ihabbar = [Ihab(1,1) Ihab(2,2) Ihab(3,3)];• wCyc = [0,0,sqrt(g/l)];• rSS = 10; %radius of Superstructure • hSS = 50; %height of superstructure• VSS = rSS^2 *pi * hSS;%Volume of SS in m^3• mSS = densitycyc * VSS;%mass of Superstructure with density

assumedot be of aluminum•• %The super structures moment of Inertia modeled as a cylinder • ISuper = [1/12*mSS*(3*rSS^2+h^2),1/12*mSS*(3*rSS^2+h^2),1/2 * mSS *

rSS];• odElev = 2.5;%m Outer diameter of Elevator• idElev = 2;%m Innderdiameter elevator• hElev = 388.88;%m Length of Elevator• VElev = h * pi * (odElev^2-idElev^2);

Code Continued• mElev = densitycyc * VElev;%mass of elevator• IxxE = 1/2 * mElev * (odElev^2+idElev^2);• IyyE = 1/12 * mElev * (3*(odElev^2+idElev^2)+hElev^2);• IzzE = IyyE;• %Moment of Inertia for the elevator as a hollow cylinder• IElev = [IxxE,IyyE,IzzE];•• %Moment of Inertia of the cycler modeled with two elevators, • %two Habitation modules and the superstructure • ICycler = 2 * (IElev+Ihabbar)+ISuper;•• %angular momentum of the cycler at a rate of rotation that gives us• %1 g of acceleration in the habitation modules• LCycler = cross(ICycler,wCyc)• Lreq = norm(LCycler)/2

Code Continued

• %test of the mass of the habition units that gave us a result similar to• %the other method using the approximate density of the ISS• %MassI2 = (3220/2+2363/2)*1.0106e+03•• %%%%%CMG Sizing and equations•• wf = 6600/60*2*pi;%radians per second rotation rate of the flywheel• densityf = 8000;%kg/m^3 %density of the flywheel as stainless steel• grate = [0,deg2rad(3.1),0];%rad/s max rotation rate of the gimbals• wf = [0 , 0,wf];• thetap = deg2rad(53.1);%angle of the orientation of the pyrimad

orientation• %of the gyroscopes

Code Continued

• %calculates the maximum Angular Momentum in any direction for a given• %radius of gyrscope in the pyrimadal configuration• for rf =1:.1:100• hf = rf/10;• vf = hf* rf^2*pi;%m^3• mf = densityf*vf;%kg• Ixf = .5 * mf * rf^2;%kg*m^2• Lf = wf(3)*Ixf;• Lmax = 4*sin(thetap)*Lf;• if Lmax > Lreq• break• end• end

Code Continued

• Iyf = 1/12*mf*(3*rf^2+hf^2);%kg*m^2• Izf = Iyf;• If = [Ixf, 0, 0; 0,Iyf,0;0,0,Izf]; %moment of inertia of the flywheel• t = cross(grate,If*wf')•• %%below is the code used to find the optimal pyramid angle for • %%a very even momentum envelope it was used to find an agle of • %%53.1 so that is what will be used for the code above, that is sizing• %%the flywheel• % for i = 0:.1:90 • % thetap = deg2rad(i);• % hf = wf(3)*Ixf

Code Continued

• % hmax = 4*sin(thetap)*hf• % hcheck = hf * 2*(1+cos(thetap))• % t = norm(norm(hcheck) - norm(hmax));• % if t< check• % pyramidangle = i• % check = t• % end• % end

Code: Main Function• function [output,output2] = Forces(input, SurfaceArea)• time = input.Vehicle1.julian_date;• x1EMO = input.Vehicle1.position(:,1);• y1EMO = input.Vehicle1.position(:,2);• z1EMO = input.Vehicle1.position(:,3);• x2EMO = input.Vehicle2.position(:,1);• y2EMO = input.Vehicle2.position(:,2);• z2EMO = input.Vehicle2.position(:,3);• x3EMO = input.Vehicle3.position(:,1);• y3EMO = input.Vehicle3.position(:,2);• z3EMO = input.Vehicle3.position(:,3);• x4EMO = input.Vehicle4.position(:,1);• y4EMO = input.Vehicle4.position(:,2);• z4EMO = input.Vehicle4.position(:,3);

E = 23 ;%degrees, axial tilt

Cyclerpos1 = EMO2EME([x1EMO, y1EMO, z1EMO], E);Cyclerpos2 = EMO2EME([x2EMO, y2EMO, z2EMO], E);Cyclerpos3 = EMO2EME([x3EMO, y3EMO, z3EMO], E);Cyclerpos4 = EMO2EME([x4EMO, y4EMO, z4EMO], E);

x1 = Cyclerpos1(1,:);y1 = Cyclerpos1(2,:)z1 = Cyclerpos1(3,:);x2 = Cyclerpos2(1,:);y2 = Cyclerpos2(2,:);z2 = Cyclerpos2(3,:);x3 = Cyclerpos3(1,:);

Code: Main Function• y3 = Cyclerpos3(2,:);• z3 = Cyclerpos3(3,:);• x4 = Cyclerpos4(1,:);• y4 = Cyclerpos4(2,:);• z4 = Cyclerpos4(3,:);•• L = [length(Cyclerpos1(:,1)),length(Cyclerpos1(1,:))]• dist_Sun = [0,0,0,0];• Solar_rad = [];• udistSun=[];•• for count = 1:1:length(x1);• dist1 = sqrt((x1(count))^2+y1(count)^2+z1(count)^2);• dist2 = sqrt((x2(count))^2+y2(count)^2+z2(count)^2);

dist3 = sqrt((x3(count))^2+y3(count)^2+z3(count)^2);

dist4 = sqrt((x4(count))^2+y4(count)^2+z4(count)^2);

dist_Sun(count,:) = [dist1; dist2; dist3; dist4];

%udistSun(1:4,count) = [[Cyclerpos1(1:3,count)./dist1],[Cyclerpos2(1:3,count)./dist2], [Cyclerpos3(1:3,count)./dist3],[Cyclerpos4(1:3,count)./dist4]];end

Planetpos = xlsread('Planetpos.xls', 'Sheet1', 'A3:AB13110');

Code: Main Function• posMercury = Planetpos(:,2:4)./149597871;• posVenus = Planetpos(:,5:7)./149597871;• posEarth = Planetpos(:,8:10)./149597871;• posMars = Planetpos(:,11:13)./149597871;• posJupiter = Planetpos(:,14:16)./149597871;• posSaturn = Planetpos(:,17:19)./149597871;• posUranus = Planetpos(:,20:22)./149597871;• posNeptune = Planetpos(:,23:25)./149597871;• posMoon = Planetpos(:,26:28)./149597871;•• % L1 = length(x1EMO)• % L2 = length(posMercury(:,1))• % L3 = length(x1)• [uC_Me, distC_Me] = Distance(input, posMercury);

[uC_V, distC_V] = Distance(input, posVenus);[uC_E, distC_E] = Distance(input, posEarth);[uC_Ma, distC_Ma] = Distance(input, posMars);[uC_J, distC_J] = Distance(input, posJupiter);[uC_S, distC_S] = Distance(input, posSaturn);[uC_U, distC_U] = Distance(input, posUranus);[uC_N, distC_N] = Distance(input, posNeptune);[uC_Mo, distC_Mo] = Distance(input, posMoon);

Code: Main Function

• L1 = [length(uC_Me(1,:)),length(uC_Me(:,1))];• % Radiation from sun• Rss = dist_Sun;• c = 2.998*10^8;• k = 1;• fO=1361; %W/m^2• Force_rad = k.*SurfaceArea.*fO./(c.*Rss.^2);•• % Force_sunrad = udistSun.*Force_rad;•• % Reflected Solar rad • Me_radius = 2439.7;• V_radius = 6051.8;

E_radius = 6378.1;Ma_radius = 3396.2;J_radius = 71492;S_radius = 60268;U_radius = 25559;N_radius = 24764;Mo_radius = 1738.1;

Me_a = 0.119;V_a = 0.75;E_a = 0.29;Mo_a = 0.123;Ma_a = 0.16;J_a = 0.343;S_a = 0.342;U_a = 0.29;N_a = 0.31;

Code: Main Function• % sun_rad_Me =

reflective_solar_rad(SurfaceArea,Me_a,Me_radius,fO,c,posMercury(:,1:13108),distC_Me(:,1:13108));• % sun_rad_V =

reflective_solar_rad(SurfaceArea,V_a,V_radius,fO,c,posVenus(:,1:13108),distC_V(:,1:13108));• % sun_rad_E =

reflective_solar_rad(SurfaceArea,E_a,E_radius,fO,c,posEarth(:,1:13108),distC_E(:,1:13108));• % sun_rad_Ma =

reflective_solar_rad(SurfaceArea,Ma_a,Ma_radius,fO,c,posMars(:,1:13108),distC_Ma(:,1:13108));• % sun_rad_J =

reflective_solar_rad(SurfaceArea,J_a,J_radius,fO,c,posJupiter(:,1:13108),distC_J(:,1:13108));• % sun_rad_S =

reflective_solar_rad(SurfaceArea,S_a,S_radius,fO,c,posSaturn(:,1:13108),distC_S(:,1:13108));• % sun_rad_U =

reflective_solar_rad(SurfaceArea,U_a,U_radius,fO,c,posUranus(:,1:13108),distC_U(:,1:13108));• % sun_rad_N =

reflective_solar_rad(SurfaceArea,N_a,N_radius,fO,c,posNeptune(:,1:13108),distC_N(:,1:13108));

Code: Main Function

• % sun_rad_Mo = reflective_solar_rad(SurfaceArea,Mo_a,Mo_radius,fO,c,posMoon(:,1:13108),distC_Mo(:,1:13108));

• % • % Sun_ref_rad = uC_Me.*sun_rad_Me +

uC_V.*sun_rad_V + uC_E.*sun_rad_E + uC_Ma.*sun_rad_Ma + uC_J.*sun_rad_J + uC_S.*sun_rad_S + uC_U.*sun_rad_U + uC_N.*sun_rad_N + uC_Mo.*sun_rad_Mo;

•• % Solar wind•• Sun_wind = 2.3*10^(-9)./dist_Sun;• %Force_sun_wind = udistSun.*Sun_wind;•

%planet radiation

Me_temp = 438+273;V_temp = 468+273;E_temp = 25+273;Ma_temp = -28 + 273;J_temp = -119+273;S_temp = -139 + 273;U_temp = -196 + 273; N_temp = -202 + 273;

P_rad_Me = Planetradiation(SurfaceArea, Me_temp, Me_radius, c, distC_Me);P_rad_Ma = Planetradiation(SurfaceArea, Ma_temp, Ma_radius, c, distC_Ma);P_rad_Mo = Planetradiation(SurfaceArea, Ma_temp, Mo_radius, c, distC_Mo);

Code: Main Function

• P_rad_V = Planetradiation(SurfaceArea, V_temp, V_radius, c, distC_V);

• P_rad_E = Planetradiation(SurfaceArea, E_temp, E_radius, c, distC_E);

• P_rad_J = Planetradiation(SurfaceArea, J_temp, J_radius, c, distC_J);

• P_rad_S = Planetradiation(SurfaceArea, S_temp, S_radius, c, distC_S);

• P_rad_U = Planetradiation(SurfaceArea, U_temp, U_radius, c, distC_U);

• P_rad_N = Planetradiation(SurfaceArea, N_temp, N_radius, c, distC_N);

• %P_rad = uC_Me.*P_rad_Me + uC_V.*P_rad_V + uC_E.*P_rad_E + uC_Ma.*P_rad_Ma + uC_J.*P_rad_J+ uC_S.*P_rad_S + uC_U.*P_rad_U + uC_N.*P_rad_N+ uC_Mo.*P_rad_Mo;

•

% Total net force %Net_Force = P_rad + Force_sun_wind + Sun_ref_rad + Force_sunrad;%Plottingcounter = linspace(0,length(Force_rad(:,1)),length(Force_rad(:,1)));%vehiles forceuuu = sum(Force_rad(4,:));figure(1)plot(counter,Force_rad(:,1),'r');hold on;plot(counter,Force_rad(:,2),'k');plot(counter, Force_rad(:,3),'b');plot(counter,Force_rad(:,4),'g');

title('Solar Radiation Force Excerted on the Cycle at Each Positon')

Code: Main Function

• xlabel('Position')• ylabel('Force Felt (N)')• legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')•• counter =

linspace(0,length(P_rad_Me(:,1)),length(P_rad_Me(:,1)));• % figure(2)• % plot(counter,P_rad_Me(:,1),'r');• % hold on;• % plot(counter,P_rad_Me(:,2),'k');• % plot(counter, P_rad_Me(:,3),'b');• % plot(counter,P_rad_Me(:,4),'g');• % • % title('Solar Radiation Force Refleced off of Mercury')

% xlabel('Position')% ylabel('Force Felt (N)')% legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')% % counter = linspace(0,length(P_rad_E(:,1)),length(P_rad_E(:,1)));% figure(3)% plot(counter,P_rad_E(:,1),'r');% hold on;% plot(counter,P_rad_E(:,2),'k');% plot(counter, P_rad_E(:,3),'b');% plot(counter,P_rad_E(:,4),'g');% % title('Solar Radiation Force Refleced off of Earth')

Code: Main Function

• % xlabel('Position')• % ylabel('Force Felt (N)')• % legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')• % • % counter =

linspace(0,length(P_rad_V(:,1)),length(P_rad_V(:,1)));• % figure(4)• % plot(counter,P_rad_V(:,1),'r');• % hold on;• % plot(counter,P_rad_V(:,2),'k');• % plot(counter, P_rad_V(:,3),'b');• % plot(counter,P_rad_V(:,4),'g');• % • % title('Solar Radiation Force Refleced off of Venus')

% xlabel('Position')% ylabel('Force Felt (N)')% legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')% % counter = linspace(0,length(P_rad_Ma(:,1)),length(P_rad_Ma(:,1)));% figure(5)% plot(counter,P_rad_Ma(:,1),'r');% hold on;% plot(counter,P_rad_Ma(:,2),'k');% plot(counter, P_rad_Ma(:,3),'b');% plot(counter,P_rad_Ma(:,4),'g');% % title('Solar Radiation Force Refleced off of Mars')

Code: Main Function

• % xlabel('Position')• % ylabel('Force Felt (N)')• % legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')• % • % counter =

linspace(0,length(P_rad_J(:,1)),length(P_rad_J(:,1)));• % figure(6)• % plot(counter,P_rad_J(:,1),'r');• % hold on;• % plot(counter,P_radJ(:,2),'k');• % plot(counter, P_rad_J(:,3),'b');• % plot(counter,P_rad_J(:,4),'g');• % • % title('Solar Radiation Force Refleced off of Jupiter')

% xlabel('Position')% ylabel('Force Felt (N)')% legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')% % counter = linspace(0,length(P_rad_S(:,1)),length(P_rad_S(:,1)));% figure(7)% plot(counter,P_rad_S(:,1),'r');% hold on;% plot(counter,P_rad_S(:,2),'k');% plot(counter, P_rad_S(:,3),'b');% plot(counter,P_rad_S(:,4),'g');% % title('Solar Radiation Force Refleced off of Saturn')

Code: Main Function

• % xlabel('Position')• % ylabel('Force Felt (N)')• % legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')• % • % counter =

linspace(0,length(P_rad_U(:,1)),length(P_rad_U(:,1)));• % figure(8)• % plot(counter,P_rad_U(:,1),'r');• % hold on;• % plot(counter,P_rad_U(:,2),'k');• % plot(counter, P_rad_U(:,3),'b');• % plot(counter,P_rad_U(:,4),'g');• % • % title('Solar Radiation Force Refleced off of Uranus')

% xlabel('Position')% ylabel('Force Felt (N)')% legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')% % counter = linspace(0,length(P_rad_N(:,1)),length(P_rad_N(:,1)));% figure(9)% plot(counter,P_rad_N(:,1),'r');% hold on;% plot(counter,P_rad_N(:,2),'k');% plot(counter, P_rad_N(:,3),'b');% plot(counter,P_rad_N(:,4),'g');% % title('Solar Radiation Force Refleced off of Neptune')

Code: Main Function

• % xlabel('Position')• % ylabel('Force Felt (N)')• % legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')• % • % counter =

linspace(0,length(P_rad_Mo(:,1)),length(P_rad_Mo(:,1)));• % figure(11)• % plot(counter,P_rad_Mo(:,1),'r');• % hold on;• % plot(counter,P_rad_Mo(:,2),'k');• % plot(counter, P_rad_Mo(:,3),'b');• % plot(counter,P_rad_Mo(:,4),'g');• % • % title('Solar Radiation Force Refleced off of Moon')

% xlabel('Position')% ylabel('Force Felt (N)')% legend('Vehicle1','Vehicle2','Vehicle3','Vehicle4')avg_Mo = sum(sum(P_rad_Mo))/4avg_Me = sum(sum(P_rad_Me))/4avg_Ma = sum(sum(P_rad_Ma))/4avg_V = sum(sum(P_rad_V))/4avg_E = sum(sum(P_rad_E))/4avg_J = sum(sum(P_rad_J))/4avg_S = sum(sum(P_rad_S))/4avg_U = sum(sum(P_rad_U))/4avg_N = sum(sum(P_rad_N))/4avg_Sun = sum(sum(Force_rad))/4avg_wind = sum(sum(Sun_wind))/4%Vehicle distanceoutput = Force_rad(1,:);output2 = sum(Force_rad');

Code: EMO2EME

• function output = EMO2EME(input, E)•• EMO2EME = [1,0,0;0,cosd(E),sind(E);0,-sind(E),cosd(E)];• x1 = [];• for count = 1:length(input(:,1))• converted = EMO2EME*input(count,:)';• x1 = [x1, converted];• end• output = x1;

Code: Distance

• function [udist, dist] = Distance(point1, point2)•• dist = [];• C1pos = point1.Vehicle1.position(:,1:3);• diff1 = C1pos-point2;• C2pos = point1.Vehicle2.position(1:13108,1:3);• L = length(C2pos(:,1));• L = length(point2(:,1));• diff2 = C2pos-point2;• diff3 = point1.Vehicle3.position(1:13108,1:3)-point2;• diff4 = point1.Vehicle4.position(1:13108,1:3)-point2;•• x1 = diff1(:,1);• y1 = diff1(:,2);

z1 = diff1(:,3);x2 = diff2(:,1);y2 = diff2(:,2);z2 = diff2(:,3);x3 = diff3(:,1);y3 = diff3(:,2);z3 = diff3(:,3);x4 = diff4(:,1);y4 = diff4(:,2);z4 = diff4(:,3);udist1 = [0,0,0,0];udist2 = [0,0,0,0];udist3 = [0,0,0,0];udist4 = [0,0,0,0];for count = 1:1:length(x1)

dist1 = sqrt((x1(count))^2+y1(count)^2+z1(count)^2);

dist2 = sqrt((x2(count))^2+y2(count)^2+z2(count)^2);

Code: Distance

• dist3 = sqrt((x3(count))^2+y3(count)^2+z3(count)^2);• dist4 = sqrt((x4(count))^2+y4(count)^2+z4(count)^2);• dist = [dist1; dist2; dist3; dist4];• udist1(count,1:3) = diff1(count,:)./dist1;• udist2(count,1:3) = diff2(count,:)./dist2;• udist3(count,1:3) = diff3(count,:)./dist3;• udist4(count,1:3) = diff4(count,:)./dist4;•• end•• udist = [udist1, udist2, udist3, udist4];

Code: Reflected Solar Radiation

• function sun_rad = reflective_solar_rad(A,a,Rp,f,c,ppos,cpos)• sun_rad = 2.*A.*a.*Rp.^2.*f./(3.*c.*(ppos).^2.*(cpos).^2);

Code: Planet Radiation

• function output = Planetradiation(A, T, Rp, c, rps)•• E = 5.67*10^(-8) * T^4;• output = A.*E.*Rp.^2./(c.*rps.^2*149597871);

Work Cited

Garcia, Mark. “International Space Station Facts and Figures.” NASA, NASA, 28 Apr. 2016, www.nasa.gov/feature/facts-and-figures.

Longuski, James M., et al. “Survet of Nongravitational Forces and Space Enviromental Torques: Applied to the Galileo.”Journal of Guidance, Control, and Dynamics, 1992.

UCAR. “Calculating Planetary Energy Balance & Temperature.” UCAR Center for Science Education, 2015, scied.ucar.edu/planetary-energy-balance-temperature-calculate.

“Albedo.” Albedo of the Earth, hyperphysics.phy-astr.gsu.edu/hbase/phyopt/albedo.html.“Equatorial Radius of the Planets and the Sun.” Smart Conversion,

www.smartconversion.com/otherInfo/Equatorial_Radius_of_planets_and_the_sun.aspx.“Solar System Temperatures.” NASA, NASA, solarsystem.nasa.gov/resources/681/solar-system-temperatures/.Newburn, R. L., Jr., and Gulkis, G., "Planets and Satellites ofthe Outer Solar System, Asteroids and Comets,"

Foundations ofSpace Biology and Medicine, Vol. I, Space as a Habitat, edited by M.Calvin and O. G. Gazenko, Joint USA/USSR Publication, Scientificand Technical Information Office, NASA, Washington, DC, 1975,Chap. 5.

Wertz, J. R., Spacecraft Attitude Determination and Control, D.Reidel, Boston, MA, 1978, Appendix

Sarah CulpFebruary 13, 2020

Backup Slides

Fecal Matter Production Calculations• Mass: 128 g (0.128 kg) of feces is produced per person per day [6]• Volume: 300 mL is produced per person per day [3]• Dehydrating the feces leads to a mass reduction of 33% and volume reduction of 75% [6][17]• Assume: Cycler journey time 8 mos. • Assume: Taxi journey time is 4 days (Mission Design Estimate) [*]

CYCLER FECAL MATTER

Per Person Per Day Values # People # Days

Dehydration Reduction Totals

Mass (kg) 0.128 70 240 0.67 1440.768Volume (m^3) 0.0003 70 240 0.25 1.26

TAXI FECAL MATTER

Per Person Per Day Values # People # Days

Dehydration Reduction Totals

Mass (kg) 0.128 70 4 0.67 24.0128Volume (m^3) 0.0003 70 4 0.25 0.021

Assume feces water mass is sent back to water filtration system

[*] Number from Melissa Whitcomb

Taxi – Toilet Mass, Power, Volume Estimate• Assume toilet is 16 kg (based on ISS estimate) [10]• Assume toilet stalls are 1.67 m tall and 1x1 m length and width

(based on Node 3 WHC Kabin on the ISS) [2]• Assume two toilets are on the taxi, as a minimum in most aircraft

TOILETS Per Unit Value # Toilets Totals

Mass (kg) 16 2 32

Volume (m^3) 1.67 2 3.34

Power (kW) 0.2 2 0.4

Cycler – Toilet Mass, Power, Volume Estimate• Assume toilet is 5.89 kg (based on camping toilet – we have gravity, so we can reduce

mass) [15]• Assume toilet stalls are 1.67 m tall and 1x1 m length and width (based on Node 3

WHC Kabin on the ISS) [2]• Occupational Health and Safety Administration (OSHA) workplace standards require 4

toilets minimum per sex for a group of 56-80 people. [5]• Yet the ISS has 2 toilets per 6 people, and toilets have been known to break. This

equates to 23 toilets if this is metric is scaled up. [14]• Finding a happy medium: 16 toilets is the average of this max (ISS) and min (OSHA).

TOILETS Per Unit Value # Toilets TotalsMass (kg) 5.89 16 94.24Volume (m^3) 1.67 16 26.72Power (kW) N/A N/A N/A

• Considering corn as a baseline crop, it requires 2.5 lb of fertilizer (pre and post planting) per 100 𝑓𝑓𝑓𝑓2. This equates to 1.1 kg per 9.2 𝑚𝑚2. [4]

• Depending upon the square footage of the greenhouse, and assuming a 1-1 conversion rate of fecal matter to fertilizer (by mass), human waste may be overproduced for the size of the greenhouse, an eventuality that will be reached after 8 months (especially if the cycler journey was 2 years)

• Currently, Alexey Zenin has sized the greenhouse to be 7000 𝑚𝑚2, which means that 837 kg of feces would be required for use in the greenhouse, assuming a 1-1 conversion ratio. If the conversion ratio is changed to 0.5, 1674 kg are required. However, this 7000 𝑚𝑚2 is based on an system that is vastly overproducing food (by a factor of 4) and oxygen (by a factor of 2). Scaling this system back means that in a best case scenario, we are still left with excess fecal matter. [*]

• Thus, it is advisable that there be multiple methods of utilizing waste.

Cycler - Fecal Matter Used by Greenhouse

[*] Greenhouse numbers from Alexey Zenin

Cycler – Turning Excess Fecal Matter into Useable Material (Plastic)

• A process uses engineered E.coli to produce polyhydroxybutyrate (PHB) plastic from human waste. (process displayed on main slide) [3]

• This PHB can be used in conjunction with a 3D printer to create tools that they may not have brought with them to Mars and can even be used as a secondary form of interior radiation shielding. [3]

• It is helpful to note that this process dehydrates feces and recycles excess water for crewed use

• A study on the effectiveness of dehydrated feces as radiation shielding proved it equivalent in effectiveness to hydrated feces. [6]

• TRL 4

ASTROPLASTIC INFRASTRUCTURE Per Unit Value # Toilets = # Units TotalsMass (kg) 174 16 2784Volume (m^3) 1.2 16 19.2Power (kW) 5.2 16 83.2

Fecal Waste System Totals (Infrastructure + Biomass)

This is how you’re solving the problem. CYCLER HUMAN WASTE TOTALS (8 mos) Infrastruture Biomass Total

Mass (kg) 2878.24 1440.768 4319.008

Volume (m^3) 45.92 1.26 47.18

Power (kW) 83.2 0 83.2

TAXI HUMAN WASTE TOTALS (4 Days) Infrastruture Biomass TotalMass (kg) 380 24.0128 404.0128Volume (m^3) 5.74 0.021 5.761Power (kW) 10.4 0 10.4

Trash Generation Calculations

• For a crew of 4 people 4.4 kg of trash is generated daily [9]• Per person, 1.1 kg/day• Per person, 1 ft3/day [1]

CYCLER TRASH

Per Person Per Day # People # Days Reduction Totals

Mass (kg) 1.1 70 240 0.25 4620

Volume (m^3) 0.0283168 70 240 0.1 47.57

A Comparison of Two Trash Disposal Solutions• Orbital Syngas/ Commodity

Augmentation Reactor (OSCAR)• The goal is to repurpose trash in such

a way that it can be vented off or used as fuel, air water, or spacecraft repair/construction [11]

• Chemical Reaction: [11]

• Let’s assume that resulting char and tar reduces waste the same amount a trash incinerator would

• 90% by volume and 75% by mass [12]• This is the reduction accounted for on

previous slide• TRL 4

• Heat Melt Compactor (HMC)• Compacts trash into 9 in square

tiles that take up less than 1/8 of original trash volume [16]

• Water is boiled off and recollected • Tiles can be used as radiation

shielding • Trash compacted tiles are 90%

effective for radiation shielding as compared to high density polyethylene (the standard) [9]

• Consume less than 500 W [9]• Volume is 0.236 𝑚𝑚3 for 4 people

[9]• TRL 7

Choosing OSCAR

• OSCAR was chosen as the final option due to the fact that it would have a greater volume reduction.

• Due to lack of strict mass, power, and volume information, I utilized the system values from the HMC to scale and estimate my OSCAR values. Thus, while not an accurate value, it is reasonable to assume that OSCAR would be filling a similar space as the HMC if they are fulfilling the same purpose. CYCLER OSCAR Per Unit Per Person # People TotalMass (kg) 21.75 70 1522.5Volume (m^3) 0.059 70 4.13Power (kW) 0.125 70 8.75

Trash Waste System Totals (Trash + Infrastructure)

* Assume taxi waste will be transferred to Cycler and requires no infrastructure

Cycler Trash Totals (8 months) Infrastructure Trash Produced Totals

Mass (kg) 1522.5 4620 6142.5

Volume (m^3) 4.13 47.572224 51.70222

Power (kW) 8.75 0 8.75

Taxi Trash Totals (4 days) Infrastructure Trash Produced Totals

Mass (kg) 0 77 77

Volume (m^3) 0 0.79287 0.79287

Power (kW) 0 0 0

Sources[1] Bacal, K. (n.d.) Waste Management. Retrieved from

https://www.google.com/url?sa=t&rct=j&q=&esrc=s&source=web&cd=2&ved=2ahUKEwiNnd3Q_srnAhXgB50JHQtNA1wQFjABegQIDBAE&url=https%3A%2F%2Fwww.faa.gov%2Fabout%2Foffice_org%2Fheadquarters_offices%2Favs%2Foffices%2Faam%2Fcami%2Flibrary%2Fonline_libraries%2Faerospace_medicine%2Ftutorial%2Fmedia%2FIII.3.2_Waste_Management.doc&usg=AOvVaw3iTKNhYEmddyDme7bMNLyP

[2] Borrego, M., Zaruba, Y. (2019, July 7-11). Exploration Toilet Integration Challenges on the International Space Station [PDF].Retrieved from https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20190027276.pdf

[3] Chen, X., et.al. (2018, February 23). Astroplastic: A start-to-finish process for polyhydroxybutyrate production from solid human waste using genetically engineered bacteria to address the challenges for future manned Mars missions [PDF]. Retrieved from https://www.biorxiv.org/content/10.1101/288746v1.full.pdf

[4] Chinn, L. (2018, December 17). What Kind of Fertilizer Is Best for Planting Sweet Corn? Retrieved from https://homeguides.sfgate.com/kind-fertilizer-planting-sweet-corn-71656.html

[5] Department of Labor logo UNITED STATESDEPARTMENT OF LABOR. (n.d.). Retrieved from https://www.osha.gov/laws-regs/regulations/standardnumber/1915/1915.88

[6] Falck, N. (n.d.). Radiation Exposure During Space Travel: Using ... Retrieved from https://digitalcommons.calpoly.edu/cgi/viewcontent.cgi?article=1027&context=symposium

[7] Heiney, A. (2018, December 21). Rocket Ranch - Episode 7: Turning Space Trash into Gas. Retrieved from https://www.nasa.gov/mediacast/episode-7-turning-space-trash-into-gas

[8] Herridge, L. (2018, October 10). Waste Handling in a Microgravity Environment Challenge. Retrieved from https://www.nasa.gov/feature/recycling-in-space-waste-handling-in-a-microgravity-environment-challenge

Sources

[9] Lee, J., Fisher, J., & Pace, G. (2017, July 16-20). Heat Melt Compactor Development Progress [PDF]. Retrieved fromhttps://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170011317.pdf

[10] Malik, T. (2008, October 10). Space Station Toilet Breaks Again. Retrieved from https://www.space.com/5958-space-station-toilet-breaks.html

[11] Meier, A., Shah, M., Medina, J. (2019, July 7-11). Microgravity Experimentation of Long Duration Space Mission Waste Conversion[PDF]. Retrieved from https://ttu-ir.tdl.org/bitstream/handle/2346/84889/ICES-2019-129.pdf?sequence=1&isAllowed=y

[12] Municipal Waste Combustors. (n.d.). Retrieved from https://www.mass.gov/guides/municipal-waste-combustors [13] Penn, R., Ward, B., et.al. (2017, 30 December). Review of synthetic faeces and faecal sludge for sanitation and wastewater

research [PDF]. Retrieved from https://www.eawag.ch/fileadmin/Domain1/Abteilungen/sandec/publikationen/EWM/FS_Quantification_Characterisation/synthetic_human_faeces.pdf

[14] Rage against the latrine: BOTH toilets on ISS reportedly broken down. (n.d.). Retrieved from https://www.rt.com/news/474442-iss-toilet-outage-space-diapers/

[15] SereneLife 5.3 Gal. Portable Outdoor and Travel Toilet-SLCATL320. (n.d.). Retrieved from https://www.homedepot.com[16] Tabor, A. (2018, August 16). What is NASA's Heat Melt Compactor? Retrieved from https://www.nasa.gov/ames/heat-melt-

compactor [17] The Editors of Encyclopedia Britannica. (2015, January 8). Feces. Retrieved from https://www.britannica.com/science/feces

David FoxFeb. 13, 2020

Backup slidesDiscipline: Human Factors

Vehicle: Cycler

1

Energy Star television power requirements

A = tv surface area, [in2]P = Max power while on [W]

For 50” tv, A = 1068 in2

P = 78.5 * tanh(0.0005 * (1068 – 140) + 0.038) + 14

P = 50.4 W (while on)

2

References

[1] Simon, M., et al. Factors Impacting Habitable Volume Requirements: Results from the 2011 Habitable Volume Workshop. NASA, Dec. 2011, ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20110023287.pdf.

[2] Akin, David L. Introduction to Habitability. University of Maryland, 2013, spacecraft.ssl.umd.edu/academics/697S13/697S13L10.habitability2x.pdf.

[3] “Televisions Key Product Criteria.” Products | ENERGY STAR, 1 Mar. 2019, www.energystar.gov/products/electronics/televisions/key_product_criteria.

3

Backup Slides Jennifer Bergeson

Eclipse Time Equation Generation

• Angle swept out by a body was calculated using θ = 2 ∗ atan 𝑝𝑝𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑎𝑎𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏

• Where rbody is the radius of the body and abody is the semi-major axis of that body’s orbit around the sun

• Circular orbits of the planets were assumed here.

• Mean motion of Earth and Mars was calculated using 𝑛𝑛 = μ𝑠𝑠𝑢𝑢𝑠𝑠𝑎𝑎𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏3

.5

• Where abody is the semi-major axis of either Earth or Mars around the sun, and μsunis the gravitational parameter of the sun

• Eclipse time was computed using Δ𝑓𝑓𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏 = θ𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑛𝑛𝑚𝑚𝑚𝑚𝑚𝑚𝑠𝑠∗60

• The only exception to this was Δtearth. Addressed on next slide.

Eclipse Time Equation Generation Continued

• In the case of Δtearth the value of nmars was replaced with the value of nearth. The reason for this was that the eclipse time due to Mars was calculated at Mars, since this is the location in the cycler orbit when the Martian eclipse would be relevant. The eclipse time due to Mercury and Venus was also calculated at Mars, because there is no Mercury or Venus flybys, and eclipse time is the longest at the farthest point in the orbit from the sun. However, the eclipse time for Earth was calculated at the Earth flyby, because this is the point in the orbit when that time would actually be relevant.

Inner Planetary Power Loss Calculation

• Solar flux at a particular location was calculated using 𝐹𝐹 = 𝐹𝐹𝑠𝑠𝑝𝑝𝑛𝑛 ∗𝑝𝑝𝑠𝑠𝑢𝑢𝑠𝑠2

𝑎𝑎𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏2

• Where Fsun is the flux emitted at the surface of the Sun, rsun is the radius of the Sun, and abody is the semi-major axis of the body’s orbit around the Sun

• Reduced solar flux at a particular location due to an eclipsing body was calculated using 𝐹𝐹𝑝𝑝𝑟𝑟𝑏𝑏𝑝𝑝𝑟𝑟𝑟𝑟𝑏𝑏 =𝐹𝐹𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏 ∗ 1 − 𝐴𝐴𝑒𝑒𝑒𝑒𝑒𝑒𝑒𝑒𝑒𝑒𝑠𝑠𝑒𝑒𝑠𝑠𝑒𝑒 𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏

𝐴𝐴𝑠𝑠𝑢𝑢𝑠𝑠• Where Freduced is the reduced solar flux during the eclipse, Fbody is the original solar flux without

the eclipse, Aeclipsing body is the projected surface area of the eclipsing body, and Asun is the projected surface area of the Sun.

• Power loss due to eclipsing inner planet was calculated using 𝑃𝑃𝑙𝑙𝑏𝑏𝑠𝑠𝑠𝑠= 1 − 𝐹𝐹𝑚𝑚𝑒𝑒𝑏𝑏𝑢𝑢𝑒𝑒𝑒𝑒𝑏𝑏𝐹𝐹𝑏𝑏𝑏𝑏𝑏𝑏𝑏𝑏

∗ 100

• Where Ploss is the power lost to the eclipse.

Inner Planetary Power Loss Calculation Continued

• In the calculation of the inner planetary power loss, it was assumed that the Sun and inner planets were sufficiently far away that the inner planets could be modeled as a cylinder superimposed on the emissive sphere from the sun.

• The code for the eclipse time and interplanetary power calculation is very long and is not included in these slides. It can be obtained by contacting jberges@purdue.edu.

Maximum Burn Time Calculations

• The equations themselves for the maximum burn time calculations are fairly straightforward. The complicated portion was extracting the information from the provided cycler profile data. Due to the lengthy and intensive nature of the code required, it has not been attached here. (It is over 500 lines long.) If this code is desired, please contact jberges@purdue.edu.

• The provided cycler information used less than full-power burns. Power and thermal needed to know the equivalent full power burn time between refueling.

• To get the equivalent full-power burn time, the equation Δ𝑓𝑓𝑟𝑟𝑒𝑒 = Δtactual ∗ p was used.

• Δteq is the equivalent full-burn time• Δtactual is the actual burn time• p is the burn power on a 0-1 scale where 1 is a full power burn

Maximum Burn Time Calculations Continued

• The total maximum burn time between refueling was computed using the following equation: Δ𝑓𝑓 = ∫𝑟𝑟𝑛𝑛𝑟𝑟𝑏𝑏𝑝𝑝𝑛𝑛𝑛𝑛𝑟𝑟𝑝𝑝 1

𝑟𝑟𝑛𝑛𝑟𝑟𝑏𝑏𝑝𝑝𝑛𝑛𝑛𝑛𝑟𝑟𝑝𝑝 2Δ𝑓𝑓𝑟𝑟𝑒𝑒𝑝𝑝𝑒𝑒𝑎𝑎𝑎𝑎𝑙𝑙𝑟𝑟𝑛𝑛𝑛𝑛• Where encounters 1 and 2 are refueling encounters with Earth

Cycler 1 Total Burn Times Per Refueling

Cycler 2 Total Burn Times Per Refueling

Cycler 3 Total Burn Times Per Refueling

Cycler 4 Total Burn Times Per Refueling

References

[1] “Energy from the Sun”, American Chemical Society, retrieved on February 11, 2020. https://www.acs.org/content/acs/en/climatescience/energybalance/energyfromsun.html

Appendix – Mars Angular Velocity At Specific Latitude

• Assumption of linear velocity distribution between equator and pole• Used MatLab to develop the following numerical solution:

𝑅𝑅𝑘𝑘𝑎𝑎𝑝𝑝𝑠𝑠

𝑅𝑅𝑘𝑘𝑎𝑎𝑝𝑝𝑠𝑠

𝑉𝑉𝑟𝑟𝑒𝑒𝑝𝑝𝑎𝑎𝑛𝑛𝑒𝑒𝑏𝑏𝑝𝑝𝑒𝑒𝑎𝑎𝑙𝑙

𝑉𝑉𝑝𝑝𝑏𝑏𝑙𝑙𝑟𝑟𝑠𝑠

𝜃𝜃

𝜃𝜃 = 18.4°𝑉𝑉𝑝𝑝𝑏𝑏𝑛𝑛 = 𝑉𝑉𝑟𝑟𝑒𝑒𝑝𝑝𝑎𝑎𝑛𝑛𝑏𝑏𝑝𝑝𝑒𝑒𝑎𝑎𝑙𝑙 � cos 𝜃𝜃

𝑉𝑉𝑟𝑟𝑀𝑀𝑒𝑒𝑛𝑛 = 𝑉𝑉𝑙𝑙𝑎𝑎𝑝𝑝𝑛𝑛𝑟𝑟𝑙 − 𝑉𝑉𝑝𝑝𝑏𝑏𝑛𝑛

Appendix – Differential Equation Problem

• The following differential equations were solved when propogating:

𝑉𝑉𝑀𝑀 = ̇𝑟𝑟𝑀𝑀

𝑉𝑉𝑏𝑏 = ̇𝑟𝑟𝑏𝑏

𝑉𝑉𝑧𝑧 = ̇𝑟𝑟𝑧𝑧

𝑟𝑟 = 𝑟𝑟𝑀𝑀2 + 𝑟𝑟𝑏𝑏2 + 𝑟𝑟𝑧𝑧2

𝑎𝑎𝑀𝑀 = −𝜇𝜇𝑟𝑟𝑀𝑀𝑟𝑟3

𝑎𝑎𝑏𝑏 = −𝜇𝜇𝑟𝑟𝑏𝑏𝑟𝑟3

𝑎𝑎𝑧𝑧 = −𝜇𝜇𝑟𝑟𝑧𝑧𝑟𝑟3

Appendix – MatLab Listing

Works Cited1. The Topography of Mars. (2009, January 12). Retrieved January 2020, from https://www.asc-

csa.gc.ca/eng/astronomy/mars/topography.asp2. Propagator adapted from AAE 590 Dr. Frueh’s class work

Backup Slides

Passive Thermal ControlPassive Thermal Control includes insulation around the entire spacecraft as well as Louvers which are shutters for the Sun to radiate onto the surface and close when insulation is necessary. Heaters are necessary when the spacecraft is in low power mode and needs some thermal energy to regulate the temperature of the cabin. Heat pipes provide a low energy solution for moving heat around by using concentric pipes that have fluid on the outside and vapor on the inside.

Passive Thermal Control

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

MLI 119.067 ~0 396.9

Heaters 0.885 ~0 ~0

Louvers 873.153 ~0 3968.9

Heat Pipes 0.13 0 260

Heat AcquisitionHeat Acquisition is made up of Heat Exchangers and Coldplates. Coldplates are metal plates that have channels for coolant to flow through. These plates are exposed to cold environments to cool the coolant quickly. However, this device can be used to heat coolant as well if the surrounding temperature is warmer than the coolant. This is a very commonly used device for acquiring Heat into the system. The Heat exchangers are necessary for transfering heat from one coolant system into another which would be used in conjunction with the coldplate in this case.

Heat Acquisition

Component Mass (Mg)

Power (kW)

Volume (m3)

Fixed Radiators

289.630

~0 1092.9

Coldplate 15.176

0 49818

Heat RejectionHeat Rejection is necessary to accommodate the amount of heat generated by the spacecraft. This is done through simple radiators that are positioned on the outside of the spacecraft and use convection, conduction and radiation to dispel the heat generated by the electronics and other systems on board.

Heat Rejection

Component Mass (Mg)

Power (kW)

Volume (m3)

Fixed Radiators 289.63 ~0 1092.9

Heat TransportThe heat transport system is made up of all the pumps, tubes and fluid that transports the heat around to the various systems that either dispel heat or acquire heat. This sizing is based off the existing size of the components since the overall size of the spacecraft determines how long the Heat Transport system has to be.

Heat Transport

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

Pumps 0.48 2300 1.7

Plumbing Add a factor of

15%

~0 ~0

Controls Add a factor of

5%

~0 ~0

Fluids Add a factor of

5%

0 ~0

Heat pumps 10.117 135.7 ~0

MATLAB Code

MATLAB Code

SourcesHanford, Anthony J. and Michael K. Ewert. “Advanced Active Thermal Control Systems Architecture Study.” (1996).

Larson, Wiley J., and Linda K. Pranke. Human Spaceflight: Mission Analysis and Design. McGraw-Hill, 2007.

Larson, Wiley J., and James R. Wertz. Space Mission Analysis and Design. Kluwer Academic Publishers, 2005.

Kunihiro, Kazuaki, et al. “High Efficiency Power Amplifiers for Mobile Base Stations: Recent Trends and Future Prospects for 5G.” IEICE Transactions on Fundamentals of Electronics, Communications and Computer Sciences, vol. E101.A, no. 2, 2018, pp. 374–384., doi:10.1587/transfun.e101.a.374.

Gilmore, David G., and Martin Donabedian. Spacecraft Thermal Control Handbook. American Institute of Aeronautics and Astronautics, 2003.

Juhasz, A.j. “An Analysis and Procedure for Determining Space Environmental Sink Temperatures with Selected Computational Results.” Collection of Technical Papers. 35th Intersociety Energy Conversion Engineering Conference and Exhibit (IECEC) (Cat. No.00CH37022), doi:10.1109/iecec.2000.870928.

Backup Slides: Vince Bartels - Motor SelectionCopied from my document on the drive here (Power & Thermal > Tether Sling > Research Notepad): https://drive.google.com/open?id=1xwEgCk81S1fbwE0LUsVekPugNKLAJAiuyN9StFuHinU

Backup Slides: Vince Bartels - Motor SelectionFinal Selection: DC HTS Motor

Motor was selected due to its ridiculously high efficiency (peaking at 99.3%) and good fit for the task at hand. Source 2 indicated that these motors mainly become viable at industrial sizes of over 1000hp, a minimum total which our system easily reaches.

Given that HTS motors are brushless, the DC configuration retains the advantages of a typical BLDC motor:

- High torque- Low thermal emission- Vacuum compatible- Little degradability

System will need to dissipate more heat to keep material at superconducting temperatures (~93K [3]), however, Phobos’ low surface temperature will help to offset that as long as motor is not in direct sunlight.

Backup Slides: Vince Bartels - Motor SelectionSources:

[1] Marshall Space Flight Center. (n.d.). Selection Of Electric Motors For Aerospace Applications. Retrieved from http://www.klabs.org/DEI/References/design_guidelines/design_series/1229msfc.pdf

[2] Shoykhet et al. (2008, May). Development of Ultra-Efficient Electric Motors . Retrieved February 12, 2020, from https://www.osti.gov/servlets/purl/973932

[3] Vepa, R. (2018). Modeling and Dynamics of HTS Motors for Aircraft Electric Propulsion. Aerospace. doi: 10.3390/aerospace5010021

[4] Gamble, B. B., & Snitchler, G. L. Superconducting synchronous motor construction.U.S. Patent No. 5777420A

[5] Phobos In Depth. (2019, December 19). Retrieved February 12, 2020, from https://solarsystem.nasa.gov/moons/mars-moons/phobos/in-depth/

Backup Slides: Vince Bartels - Solar Panel SizingBeginning with the chart on the left, I created the below table of potential power requirements and their respective required spin times. This allowed me to generate the required GWh for future battery applications which, due to the inverse relationship between power and spin time, is the same regardless of power input.

Power Reqs from MD for Phobos [6]

GW Spin time req (days) GWh

20 2.89 1386.552

40 1.44 1386.552

60 0.96 1386.552

80 0.72 1386.552

100 0.58 1386.552

How motor power affects spin-up timeCreated by Grace Ness (MD)

Backup Slides: Vince Bartels - Solar Panel SizingThe next task was to determine radiation intensity on Phobos (and other bodies, for future reference) to complete my ultimate task of coming up with a reliable W/m2 number. This was achieved by use of the following equation [5]:

Radiation Intensity W/m^2 [5]

Phobos/Mars 588.5706216

Moon 1367

where IPhobos is the solar radiation intensity at Phobos, I1AU is the radiation intensity at a distance from the sun of 1AU, and d2

Phobos is the distance of Phobos from the sun. Assuming that the distance of Phobos is approximately equal to the distance of Mars, I came up with the table on the left

Backup Slides: Vince Bartels - Solar Panel Sizing588.57 W/m2 would be a great number, however, it fails to take into account the efficiency of the solar panels.

As such, the next major part of my research was determining the most efficient solar panels for the AM0, or zero atmosphere, condition. Source 1 provided me with ample information regarding cutting edge solar cell efficiencies, and I came to find that the highest efficiency solar cells from three major companies (Azurspace [2], Solaerotech [3], and SpectroLab [4]) provided ~32% efficiency in the AM0 environment.

Solar Array Efficiency [1]Top Production [2][3][4] 0.32SBT [1] 0.35Silicon (traditional) 0.169

SBT, or Semiconductor Wafer Bonding technology, showed the highest reported AM0 efficiency at 35%, however, it is not in production.

To finish off the below chart, I also included a traditional Silicon based solar cell with AM0 efficiency of 16.9% for comparison

Backup Slides: Vince Bartels - Solar Panel SizingCombining all of this information, I was able to make the final table seen below.

This table takes each potential solar cell, multiplies their efficiency by the radiation intensity seen at Phobos’ surface, and accounts for a lack of solar tracking (50-75% loss in power generation [5]).

Then, using all 5 potential power requirements, I was able to generate the table on the right which indicates the required solar array area for each type of solar cell.

Due to the general uncertainty of the final power required, these numbers do not take the motor efficiency into account, however, they easily can once total mass of the system is determined.

PHOBOS Array Size (km²) based on GW reqSolar Array Efficiency [1] Sun Tracking (W/m^2) No Sun Tracking (W/m^2) 20 40 60 80 100

Top Production [2][3][4] 0.32 188.3425989 117.7141243 169.90 339.81 509.71 679.61 849.52SBT [1] 0.35 205.9997176 128.7498235 155.34 310.68 466.02 621.36 776.70Silicon (traditional) 0.169 99.46843505 62.1677719 321.71 643.42 965.13 1286.84 1608.55

Backup Slides: Vince Bartels - Solar Panel SizingSources:[1] Surampudi, S., et al. (2017). Solar Power Technologies for Future Planetary Science Missions. Jet Propulsion Laboratory, Pasadena.

[2] Azurspace. (2019). 32% Quadruple Junction GaAs Solar Cell Type: Qj Solar Cell 4G32C - Advanced. Retrieved from http://www.azurspace.com/images/0005979-01-00_DB_4G32C_Advanced.pdf

[3] SolAero. (2018). Imm-α Space Solar Cell. Retrieved from https://solaerotech.com/wp-content/uploads/2018/04/IMM-alpha-Preliminary-Datasheet-April-2018-v.1.pdf

[4] Spectrolab. (n.d.). Xte-Lilt (Low Intensity Low Temperature) Space Qualified Triple Junction Solar Cell. Retrieved from https://www.spectrolab.com/photovoltaics/XTE-LILT Data Sheet _12.23.19.pdf

[5] Carufel, G. de, Li, Z. Q., Crues, E. Z., & Bielski, P. (2016). Lighting Condition Analysis for Mars’ Moon Phobos. Retrieved February 12, 2020, from https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150019633.pdf

Yashowardhan GuptaFebruary 13, 2020

Backup Slides

Discipline: Power and ThermalVehicle & Systems: Communication Satellites (Thermal)

Code for the absorptivity analysis

Mass Power Estimates

Bibliography

1. Yang, L., Li, Q., Kong, L., Gu, S., and Zhang, L., “Quasi-All-Passive Thermal Control System Design and On-Orbit Validation of Luojia 1-01 Satellite,” Sensors, vol. 19, 2019, p. 827.

2. Gilmore, D. G., and Donabedian, M., Spacecraft thermal control handbook, Reston, VA: American Institute of Aeronautics and Astronautics, 2003.

3. Rachel Roth (AAE 450)4. Price, M. K., Mass and Power Modeling of Communication Satellites, Cleveland, OH: NASA, 1991.

Kristen FleherFebruary 13, 2020

Back Up Slides

Parameter Value

Initial Orbit 300 km

Final Orbit 1,000 km

Delta V 375 m/s

Mass Bank Weight 4,900 Mg

Tether Mass 2,259 Mg

Power Required 8.02 MW

Current Required 770 A

Voltage Required 10.4 kW

ED Tethers Volume 252 m^3

Average Acceleration .0015 m/s^2

Time to Boost Orbit 2.85 days

MATLAB Script

MATLAB Script

Backup Slides – Debris shielding calculation

Source: Christiansen, E. L. (1993). Design and performance equations for advanced meteoroid and debris shields. International Journal of Impact Engineering, 14(1-4), 145–156. doi: 10.1016/0734-743x(93)90016-z

Backup Slides – Mass and volume calculations

10 m

Outer ring:𝐴𝐴𝑏𝑏 = 𝜋𝜋 𝑟𝑟𝑏𝑏2 − 𝑟𝑟𝑏𝑏 − 𝑓𝑓𝑏𝑏 2

= 𝜋𝜋 52 − 5 − 0.0031 2

= 0.0974 𝑚𝑚2

Inner ring:𝐴𝐴𝑒𝑒 = 𝜋𝜋 𝑟𝑟𝑒𝑒 + 𝑓𝑓𝑤𝑤 2 − 𝑟𝑟𝑒𝑒2

= 𝜋𝜋( 5− 0.0031− 0.3 2 − 5− 0.0031 − 0.3 − 0.0422 2)= 1.240 𝑚𝑚2

Total mass:𝑚𝑚 = 𝜌𝜌 [ 𝐴𝐴𝑏𝑏 ℎ + 𝐴𝐴𝑒𝑒ℎ + 2𝜋𝜋 𝑟𝑟𝑏𝑏2𝑓𝑓]

= 2840[ 0.0974 ∗ 50 + 1.240 ∗ 50 + 2𝜋𝜋 ∗ 52 ∗ 0.0422]= 208.7 𝑀𝑀𝑀𝑀

Note: t is the thickness on the ends of the cylinder, taken as 4.22 cm

Total volume:𝑉𝑉 = 𝜋𝜋 𝑟𝑟02 ℎ

= 𝜋𝜋 ∗ 52 ∗ 50= 3927 𝑚𝑚3

Cycler Vehicle Structures: Backup Slide

Aluminum 6061-T6 Properties [ASM]Density [kg/m3] 2700Yield Strength [MPa] 276

Cycler Vehicle Structures: Backup Slide

Cycler Tether Mass and Volume

Tether Radius [m] 400Mass [Mg] 58.5

Volume [m3] 60Spectra 2000 Material Properties [2]

Density [kg/m3] 970Yield Strength [MPa] 3325

Connective tether material was chosen as Spectra 2000. MV have been updated in the drive.

Cycler Vehicle Structures: Backup SlideSources:ASM. (n.d.). 6000 Series Aluminum Alloy. Aerospace Specification Metals

Inc. Retrieved from: http://asm.matweb.com/search/SpecificMaterial.asp?bassnum=MA6061T6.

Longuski, J.M. & Jokic, M.D. (2004). Design of Tether Sling for Human Transportation Systems Between Earth and Mars. Journal of Spacecraft and Rockets. Vol. 41, No. 6. Retrieved from: https://arc.aiaa.org/doi/pdf/10.2514/1.2413

NASA. (Oct 28, 2019). International Space Station Facts and Figures. Retrieved from: https://www.nasa.gov/feature/facts-and-figures.