preliminary design review - engineering.purdue.edu

Preliminary Design Review03/09/2021

PROJECT NEXTSTEP

AAE 450 - Spring 2021

ROLL CALL

2021 2022 2023 2024 2025

Brief StoryboardKashaf Jabbar

•Modified Apollo Lunar Module for lander

•10 Falcon Heavy

launches

•Two comms

satellites

launched

•Regolith

Collection,

Scouting,

Clearing rovers

sent

•Cargo only launched •Cargo and human transportation

•3 habitat, airlock

modules

launched

•4th comms

satellite launched

•Ground station

construction

started

•Start low-g

human health

research

•3 habitat, airlock

modules launched

•5th comms satellite

launched

•Geological

research

•Habitat at full

capacity (crew of

19)

•6th (last) comms

satellite launched

•Comms ground

station fully

constructed

•Observation

satellite launched

•Begin lunar

atmospheric

measurements

•Third comms

satellite launched

•Mining and

Construction

rovers start work

•Starship/Blue

Origin used for

cargo

Mission DesignAAE 450 - Spring 202103/09/2021

Preliminary Design Review

Team Lead: Cody Martin

Team Members: Cody Martin, Daniel Gochenaur, Mohit Pathak, Chad Blackwell, Stephen Grabowski, Matthew Popplewell, Dale Williams, Jarod Villeneuve-Utz, Alex Maietta, Eric Gooderham, Nikhil Gavini, Youssef Noureddine

Mission Design

❖ LEO orbit parameters➢ Launch from Cape Canaveral, Florida

■ Easily accessible launch location near to the equator.

➢ Altitude of 600km■ Allows for the establishment of a refueling depot that requires fewer corrective burns than at a lower orbit

■ Reduced “traffic” due to the high altitude

➢ Inclination of 28.5°, ~40° for specific launches

❖ Earth Ascent➢ Required deltaV of 9.5km/s

Lunar Ascent for Apollo Lunar Module

• Purpose: • Provide insight into what is involved with

landing/taking off from the lunar surface

• Key Assumptions:• Flat moon, constant thrust and mass flow rate,

2d, data from apollo lunar module

• Launch from Shackleton crater into 100km circular polar orbit• Approximate dV cost of about 2.0km/s

• Total ToF of 440 seconds

• From engine ignition the vehicle has traveled 277km along the lunar surface

Lunar Ascent Continued

Lunar Descent for Apollo Lunar Module

• Transfer from a 100km polar orbit into a 100x15.24km elliptical orbit

• Approximate dV cost of about 19.5 m/s

• ToF to periapsis of 1.9hrs

• At periapsis of this orbit powered descent begins• Vehicle touches down about 460 seconds later

• From engine ignition the vehicle has traveled about 425km along the lunar surface

• Approximate dV cost of about 2km/s

• To account for additional losses a total dV of 2.5km/s should be used

• Descent should start 194 degrees before target area

Lunar Descent Continued

Moon to Earth: Launch Dates

• What effect does the launch date have on ΔV from the Moon to the Earth?• Different launch dates will have different relative positions of the Moon and Earth

• Further distances require more ΔV

• Process was carried out in GMAT with:

• 4-body problem (Sun, Earth, Moon, Spacecraft)

• Solar Radiation Pressure

• Actual Lunar Orbit

Figure 1. Daily Distribution for All DeltaV Values for Moon to Earth Trajectories from October 16th, 2021 to November 30th, 2021

Figure 2. Monthly Distribution for All DeltaV Values for Moon to Earth Trajectories from October 31st, 2021 to October 16th, 2022

Moon to Earth: Launch Dates

Tabulated Data from GMAT

Figure 3. Daily Distribution for All DeltaV Values for Moon to Earth Trajectories

Tabulated Data from GMAT (cont.)

Figure 4. Monthly Distribution for All DeltaV Values for Moon to Earth Trajectories

Translunar Trajectory Specific Analysis• Main Idea: Get ΔV numbers for every specific launch on schedule

• Useful for Uncrewed Missions (mainly cargo w/o satellites)

• Key Assumptions:• Model uses Earth, Lunar, Solar Gravity and Solar Radiation Pressure

• Departure Time is always 12:00 from Earth Orbit

Earth Escape

TCM

Lunar Insertion

Sample Translunar Trajectory Breakdown

Translunar Trajectory Specific Analysis (cont.)• Years 1-6 currently analyzed

• Link to Spreadsheet with Delta V's

• Date Range: 10/8/2021 - 9/27/2027

• Minimum of 13 viable launch dates each year

• Average Time of Flight: 3.857 Days

• Average Total ΔV Cost: 3.829 km/s

• Next Steps: Complete Years 7-10 Analysis by CDR

Video of Translunar Trajectory to Moon

https://drive.google.com/file/d/1yR_z2oAA9pyO_BGraJRjeoSQExwn-5ZC/view?usp=sharing

http://drive.google.com/file/d/136mq00C4Cp1CUr_0sGUShAlpx8gXIYwE/view

Lunar Free Return• Free return trajectory designed for crewed missions

• GMAT used to simulate mission and calculate ΔV requirements• Model includes Earth, Lunar, and Solar gravity and Solar Radiation Pressure

• Displayed mission departs on October 10, 2024

• Simulation can be adapted to other departure dates

Earth-Moon Rotating View Earth Inertial ViewEarth Inertial View, Lunar Arrival

Lunar Free Return (Continued)

Moon Inertial Views of Free Return and Lunar Orbit Trajectories

Communication Constellation

• Walker Constellation configuration was chosen for:

• Stability

• Consistent contact with south pole

• Large surface coverage

• Potential for expansion

• Choose 6 satellites in 3 orbital planes

• Continuous and often redundant contact with south pole ground station

• Cross-link capabilities between satellites

• Stable - less than 100 m/s per year per satellite for stationkeeping

*Assumes ground station at center of Shackleton crater with 15° elevation angle

Element Value

a 3238.2 km

e 0

i 80°

RAAN 0°, 120°, 240°

Translunar Trajectories for Communication Satellites

• Launches will require correct phasing• Three launches will probably be required to set-up the constellation• Current research into optimal launch windows

Lunar Arrival and Corrections Animation

http://drive.google.com/file/d/11MjU3XBdWkKCQM-3vtdAD5Yfrc8y-MwJ/view


• General translunar trajectory and refinement maneuvers for communication satellite insertion have been developed

• Refinement maneuvering will generally consist of three energy reduction burns and two plane changes

Energy Reduction ManeuversRAAN Adjustment Inclination Adjustment


Representative ΔV Summary Representative Final Error

Observation Satellite

• Quasi-frozen lunar orbit at 50 km x 216 km chosen• Periapsis and apoapsis variations naturally bounded

• Won’t eventually impact the Moon like a circular orbit

• Modeled with Moon J20,20 harmonics, Earth, Sun, and SRP

• Requirement to stay within +/- 15 km of nominal altitude when taking images• Used as bounds for periapsis and apoapsis

• Burn targeting no more than once every 13.5 days (half of a lunar sidereal period)

• Yearly delta-v budget of 139 m/s (with impulsive maneuvers)• Not optimized, can be brought down further

Observation Satellite, Continued

- Varying transfer angles could provide more flexibility depending on the mission over the course of establishing the entire colony

- Transfer Angles Shown: 170-179 degrees

Analysis of Multiple Free Return Trajectories

Impact of Varying Transfer Angle

Translunar Injection Maneuver vs. Transfer Angle Time of Flight vs. Transfer Angle

- Different orbits could be used to lower delta-V requirements for station-keeping maneuvers or for communication satellites about the Lagrange points

- Lissajous Orbit- Halo Orbit

- Multiple references show the delta-V requirements could be as low as 4 m/s with a minimum of 6 maneuvers per year

Other Potential Types of Orbits

Earth-Moon Lagrange Point Locations

Orbits about Lagrange Point L1Linearized CR3BP

Lissajous Orbit about L1 Halo Orbit about L1 - In-plane and Out-of Plane frequencies are equal

Mission Launch Windows• 10 initial launch dates per year

• Launch dates fluctuate based on needs of mission

• Minimum launch window = 3 days by NASA

• Daily launch windows – 2 per day• Safe azimuths: 72° - 108° to avoid landmasses• Need to vehicle to meet lunar antipode

• Happens twice a day (over Indian Ocean and Atlantic Ocean)

• Atlantic Ocean antipode favorable for Southern Pole landing

Atlantic Antipode Pacific Antipode

Lunar Orbit and Launch Dates• Lunar anomalistic month =

27 d 13 h 18 m 33.2 s• Return to perigee

• 8.8505578 to return to same perigee

• Antipode month = 28 days• September 10, 2021 Antipode within safe

azimuths

Lunar Perigee Path

Launch Dates

Table : Launch Dates for Next 10 Years

Rendezvous Trajectories for LEO Depot•Able to get a spacecraft and a depot close within the same orbit with minimal ΔV cost

•Launch should be done in accordance with phasing angle to avoid prolonging mission length

http://drive.google.com/file/d/1WJ0thJd2O2uw0A95uXmPCMWMLdT5QFgU/view

Free Return Analysis in MATLAB• Trajectory Results

1. Slight variations in velocity caused considerable trajectory deviation

2. Likely will require trajectory corrections during transfer

CommunicationsAAE 450 - Spring 2021

9 March 2021Lead: Anna CismaruTeam members: Alec Leven, Bridget Cavanaugh

Overview

Objectives and Requirements

Main Purposes

- Ensure that reliable voice communication links exist between all entities that require them for mission execution for the full duration of the mission

- Ensure reliable and efficient data transmission between vehicles, Earth, the colony, and Mission Control for the full duration of the mission

Specific Requirements

- Ensure that main communication antennas meet size constraints of transport vehicles to be able to be shipped and constructed effectively

- Ensure that the maximum operating power of the antennas does not overburden the total power capacity of the system or subsystem it is connected to

AssumptionsAssumptions in Link Budget Calculations

- Values for pointing errors are assumed to be under the worst-case conditions- Half of the half power beamwidth of the antenna in question- Leads to consistent 3 dB loss due to pointing issues

- Rain degradation/attenuation is minimal for low-frequency bands (only apparent for Ka-band)

- Other losses due to transmitter construction amount to a maximum total of 4.5 dB (e.g. line loss, polarization loss, etc.)

- Antennas are effectively modeled as parabolic shape

System Architecture

Mission Timeline

Year 1 Year 2 Year 3 Year 4 Year 5

Launch first 2 comm

satellites

Launch 3rd comm

satellite

Launch 4th comm

satellite

Launch 5th comm

satellite

Launch 6th comm

satellite

Launch observation

satellite

❋ humans arrive

Launch: materials needed to build

ground station, Wi-Fi and DMR equipment

Begin ground station construction

Build 1 DMR and internet tower close to the habitat

Continue to expand DMR and Wi-Fi networks as colony grows

Finish ground station construction

Satellite Support

Communication Satellite Link Budget

Figure 1: Maximized Link Margin for Lunar-orbiting satellite using Ka band

Table 1: Link Margin and other helpful values at a satellite power of 444 W and an antenna diameter of 0.970 m.

Wheatley Station: Satellite AntennaGeneral Information:- modeled after NEN ground stations:

receives in X band and S band, transmits in S band only- Is therefore fully compatible with

both Communications and Observation Satellites

- diameter: 10 m- transmitter power: 200 W- gain: 56 dBi

Table: X Band Communication Satellite Downlink to WSS Antenna

Ground Infrastructure

Wheatley Station: Earth Antenna

WSE Specifications:- Receives and transmits in S-band- Can handle up to 100 Mbps data rate - Diameter: 7.6 m- Transmitter Power: 362.7 W- Transmitter Gain: 40.4 dBi

Additional Information: - Looking at considering ways to minimize physical

size of antenna for transport purposes, but keep same effective transmitter diameter

- Calculating specific periods when connection will be strongest with NEN (for sending large data packages)

Top: Summarized Link Budget for WSE S-BandBottom: Transmitter Power vs. Diameter for Link Margin Optimization

Wi-Fi Network

Crew + Bath

Crew + Bath Crew + Bath

Crew + Bath

CAD Model by John Matuszewski

Lab/Gym

Food Prep/Dining

Farm

Used for recreational communications (live voice and video to/from Earth, email, etc. Will likely also be used to transmit operational data (system status, etc) to Mission Control. Communications components are shown in red.

laptop

laptop laptop

laptop

modem and router

Tower Modems, connections, and repeater boxes will be bought off the shelf

Digital Mobile Radio Voice Network

Colonists: Motorola DEP 450

4 W maximum

~0.7 lbs

5.0 x 2.4 x 1.7 inches

$200-300

Rovers: Motorola XPR 5550e

40 W maximum

~4 lbs

2.1 x 6.9 x 8.1 inches

$500-600

Next Steps

Action Items- Enable S or X band voice communication between comm sats and Earth

- Vehicles-satellites will present comms payload design for the comm sats. This already includes S

and X band patch antennas, as we are trying to stick to these

- Enable crosslinking between imaging sat and comm sats

- Wi-Fi network:

- choose modems and repeater units

- choose cable connection (fiber optic or other)

- DMR voice network:

- choose repeater units

- Finalize all satellite and ground system power and mass calculations

- Assess launch vehicle compatibility

- Add all antennas, transmitters, receivers, radios, etc to the MEL

- Do risk and failure analyses

Launch VehicleAAE 450 - Spring 202103/09/2021Launch Vehicle Team

Purpose, Requirements, and Assumptions

• Purpose

• Transport cargo and colonists to LEO, TLI, LLO

• Requirements

• Minimize cost per Mg to desired orbit

• Safely transport colonists to LLO starting 2023

• Transport colonists from lunar colony back to earth after designated habitation cycle

• Assumptions

• Starship will not be available until 2023 (Q3 or Q4)

• Adequate supply of in-service launch vehicles

Propulsion PDR

Pre 2023 Launch Vehicles

• Requirements• Starship unavailability requires other heavy lift vehicles• Transport cargo past LEO at lowest $/Mg

• Falcon Heavy (Expendable)• Initial cargo transports to LLO

• Falcon Heavy (Reusable)• Fuel depot to LEO

• Falcon 9 (Expendable)• Launch satellites into Lunar Orbit

Propulsion PDR

Post 2023 Launch Vehicles

• Requirements• Starship unavailability requires other heavy lift vehicles• Transport cargo past LEO at lowest $/Mg

• Starship• Colonist transport becomes available• Fairing size increases from 300m3 to 1100m3

• Falcon Heavy• Cheapest vehicle to transport required payloads to orbit• Approximately 1.4 million dollars per Mg to LEO, 50% less than any current

vehicle

Propulsion PDR

Launch Vehicle Statistics

Vehicle Cost per Launch Fairing Volume (m3) Payload to LLO (Mg) Cost per Mg to LLO

(M$ / Mg)

Falcon Heavy Reusable 90 million 295.2 4.00-5.00 18-22.5

Falcon Heavy

Expendable

150 million 295.2 14.1-15.3 9.80-10.64

Falcon 9 62 million 240.61 4.00-5.00 12.4-15.5

Starship (with LEO

refuel)

2-10 million 1100 87.1 - 90.7 .023 - .115

Propulsion PDR

Next Steps• Continue developing the launch

mission code• Develop launch schedule

Propulsion PDR

PropulsionLunar Descent ModuleAAE 450 - Spring 2021

Timeline

• Phase 1: 2021-2023• Proposed HLS Landers not available• Modified Apollo Lunar Module- Cargo Variant

• Phase 2: Late 2023-2024• Starship and Blue Origin Cargo Variants Become Available• More Significant Cargo Missions

• Phase 3: 2024 on• All HLS Systems Available• Cargo and Human Transport Missions

Apollo Lunar Module Cargo Variant

• Modifications• Remove Descent Stage• Remove Water Storage• Reduce ΔV Budget

• Increased Computing Power

• Lower Factor of Safety For Cargo

Lunar Lander Descent ModuleSource: NASA.gov

Apollo Lunar Module Cargo Variant

Original Mass: 15.301 Mg

Mass Removed: 5.17 Mg

Mass With Reductions: 10.131 Mg

Payload With Current Fuel Capacity: 5.82 Mg

Payload with Increased Fuel Capacity

(Falcon Heavy):7.45 Mg

Starship Lunar Variant

• Starship Lunar Variant expected to be completed early 2024

• Currently needs modifications to heat shielding, payload deliver, refueling, and reaction control systems

• Can deliver up to 100 people to lunar surface• SpaceX estimates 100+ Mg of payload

delivery to lunar surface• From propulsion team estimates, a

more precise payload number is 93.13 Mg

• In order to deliver this payload, 1090 Mg of propellant refuel in LEO is necessary (Trevor Pfeil)

• To meet this refueling capacity, 6 previous Starship launches must occur to provide propellant to refueling station

Starship Lunar Variant Source: Starship User Guide (SpaceX)

Human Return System • Human missions planned for 2024• Starship will be unable to ferry people in the initial stage due to lack of heat

shielding• Blue Origin’s HLS and Dynetics’ Dynetics HLS are planned to be operational by

2024 to deliver 2 man crew• Blue Origin’s HLS is the primary choice

• However, which ever proves operation first will be used due to minimal tradeoff between the two

Human Return System

• Blue Origin HLS planned to be competed in 2023• 2 Crew capacity and 4.5 Mg• Launched Via New Glenn or Vulcan Centaur

• Vulcan Centaur cost per launch $100-200 Million, first stage planned to be reusable

• New Glenn costs not published, first stage also planned to be reusable

• BE-7 throttleable and gimballed engine• 44.48 KN max thrust, throttleable to 0.89 KN• Hydrogen fuel

• Will shuttle to NASA’s Lunar Gateway to transfer to the Orion spacecraft to land back on Earth

Images obtained from:https://www.blueorigin.com/blue-moon/lunar-transport

Blue Origin National Team HLS

Cargo/Starship Back-up Plan

• Prior to human transportation, cargo must be carried to the lunar surface

• Two viable options

• Blue Moon (Blue Origin) • Crewed and cargo variants

• ALPACA (Dynetics)• Optimized for human transport • Still in conceptual development

Blue Moon

• Cargo Variant

• Blue Moon with Ascent Vehicle• Since LOX/LH2 propellants, can add additional payload mass by slightly

reducing ascent propellants and manufacturing them on the lunar surface

• Designed for Southern Lunar Pole Landing Site

Images obtained from:https://www.blueorigin.com/blue-moon/lunar-transport

Payload Mass 4.5 Mg

Propulsion System BE-7 LOX/LH2 delivering 40 kN of

thrust

Propellant DepotsAAE 450 - Spring 20213/9/2021

Noah Niklaus, Trevor Pfeil, Theresa Sitter, Sarah Wagenaar

Purpose of Orbital Propellant Depots

• LEO Depot: Starship can land on the Moon

• Lunar Depot: Starship can return from the Moon to LEO

• No simultaneous launch requirement

• Increase flexibility in launch scheduling

Performance Parameters

• Maximum payload mass to 600 km Earth orbit = 100Mg

• GLOW of Super Heavy Booster = 5,000Mg

• Starship inert mass = 120Mg

• Super Heavy Booster inert mass = 180Mg

• Raptor engine vacuum ISP = 380 sec

• Raptor engine sea level ISP = 353 sec

Design Assumptions

• Zero / low g cryogenic fluid transfer is possible

• Starship tanker configuration = standard Starship with zero payload

• Delta V requirements for each leg are as specified by mission design

• 3.705Mg payload returned to Earth from lunar surface (crew return)

• LLO depot in a 100 km orbit

DV Values UsedMission Leg Delta V (km/s)

1 Earth Ascent → 600 km orbit 9.5

2 Rendezvous with LEO depot 0.01

3 TLI 3.047+0.1 = 3.147

4 FRD 0.4555+0.1 = 0.5555

5 LOI 0.6772 + 0.1 = 0.7772

6 LLO (100 km) → Lunar Surface 2.5

7 Lunar Surface → LLO 2.2

8 LLO → LEO 4.1124+0.1 = 4.2124

Depot Minimum Requirements

For a mission from Earth → Lunar Surface:

● With 1 full refueling stop in LEO, can take up to 93.13 Mg of payload

● LEO Depot min. required storage: 1090.182 Mg

○ (for 93.13 Mg of payload)

Earth Ascent

Full Refuel at LEO Depot

TLI +FRD +

LOI

Lunar Descent

Mission Progression

Depot Minimum Requirements

For a mission from Earth→ Moon → LEO (100 Mg of payload to the Moon):

● LEO Depot min. required storage: 411.449 Mg

● Lunar Depot min. required storage: 410.874 Mg + 273.972 Mg = 684.846 Mg

Earth Ascent

Partial Refuel at

LEO Depot

TLI +FRD +

LOI

Partial Refuel at

Lunar Depot

Lunar Descen

t

Lunar Ascent

Partial Refuel at

Lunar Depot

LLO to LEO

Mission Progression

Flight Counts

For a mission from Earth → Lunar Surface (93.13 Mg payload)

● 6 tankers to load LEO depot

● 1 Starship with payload to the lunar surface

For a mission from Earth→ Moon → LEO (100 Mg of payload to the Moon)

● 17 tankers to load both depots

● 1 Starship with payload to the lunar surface

Proposed Depot Design

● Docking interface with same fuel lines allowing connection for transfer

● Control thrusters for directing of fuel in and out of the storage space

● A storage tank or tanks depending on the amount of fuel needed

● Possible components for temperature control

○ Include multi-layered super isolation around tanks, structural

isolation of tank, multilayered sun-shield and solar powered cryo-

cooler

Next Steps

• Understanding structural requirements of cryogenic fluid storage in space

• Assess requirements for zero-g fluid transfer

• Determine thermal loads for long and short term propellant storage

• Set position of lunar fuel depot to minimize total propellant usage

• Examine feasibility and rates of lunar fuel production

• Justify specific amount of propellant that needs to be stored at each depot

• Estimate station masses and create a launch schedule for implementation

Systems - HabitatAAE 450 - Spring 202103/09/2021Team Lead: Sean AndersonTeam Members: Dylan Anderson, Sarah Baxter, Lara Cackovich, Seth Cantrell, Derek Carpenter, George Carroll, Wellington Froelich, Jason Kelly, Christian Mandrell, Hunter Mattingly, John Matuszewski, James Pannullo

Purpose & ProblemA permanent lunar surface base is a key part of colonizing the Moon. The base will protect the crew from radiation and micrometeorite debris while serving as a place for the crew to live and work on the lunar surface.

● Requirements○ Habitat modules must fit in Starship○ Habitat should be designed to be modular and easily expandable○ The modules should protect the crew from radiation and micrometeorite debris

● Need to Determine○ Type of habitat module○ Shape and sizing of the habitat modules○ Layout of the modules○ Mass, power usage, and volume of the habitat

● Assumptions○ Starship is ready by 2024 and capable of sending at least 45 Mg to lunar surface○ Starship cost $200M per launch

Systems - Habitat (Sean)

Human Factors

Aspects such as mental health, physical health, safety, etc. must be considered in the habitat design process. It is also important that the crew feels comfortable in the habitat for the long-duration mission.

● Requirements for Habitat Structure

○ Overall Volume/person = 103 m3/person

○ Bedroom Volume = 23.3 m3

○ Minimum ceiling height = 3.1 m

○ All rooms must include at least 1 monitor/tv to display images or videos to cope with homesickness

○ Bedrooms shall include: a bunkable bed, desk, chair, and wardrobe/dresser for storage

Systems - Habitat (Sean, Lara)

Solution: Lunar HabitatThe proposed lunar habitat is an inflatable/rigid hybrid structure comprised of 7 habitat modules and 6 airlock/joint modules. The design allows for modularity and scalability while meeting crew needs.

● Total Mass: 165.67 Mg

○ Habitat: 64.7 Mg

○ Power: 100.97 Mg● Total Power Needed: 200 kW● Total Volume: 4502 m3

○ Habitat: 4502 m3

○ Power: 326.4m3

● Crew Size: 19● Total Habitat Cost: $4.69B

○ Habitat: $1.41B○ Power: $3.28B

Systems - Habitat (Sean, Wellington)

CAD Models by John Matuszewski

Overview

● Construction & Materials○ Inflatable/Rigid Hybrid Pill-shaped

module○ External cage supports regolith

● Layout○ 4 crew modules, 1 farm, 1

kitchen/dining/misc., 1 lab○ Allows up to 6 egress points

● Modularity○ 3 Simple designs: habitat, airlock, and

connector○ Allows for scalability○ Sustainable growth

● Safety○ 2 exits per module○ Every module accessible from any

module even if one is compromised

Systems - Habitat (Derek)


Overview (Extra)Systems - Habitat (John)


http://drive.google.com/file/d/1Wwx00BsMQG_D-4Ohs_1XXlm42kqlerm7/view

Habitat Module

● Module functions○ Used WDM to determine base shape

and layout○ Ran optimization code to find

dimensions○ Flexible with the potential uses - i.e.

any module could be later optimized for any task

● Mass, Power, Volume Statistics Per Habitat

○ Mass: 7.758 Mg○ Power: 0 kW (for the structure)○ Volume: 540.2 m3

○ Cost Est.: $1,090,000 each



Habitat Module (Extra)Systems - Habitat (Derek, Jason)

Number of

People Per

Habitat

Inner

Radius

per

Module

(m)

Inner

Volume

per

Module

(m3)

Mass of

Habitat

per

Module

(Mg)

Packaged

Volume per

Module (m3)

Total Material

Cost per Module

(USD)

1.5X Material

Cost per Module

(USD)

Cost Per

Inner

Volume

(USD/m3)

Number of

Starships

(for all

required

modules)

4 3.7841 540.2 7.758 67.5 $726,395.88 $1,089,593.82 1344.67 7

Results of Jason Kelly and Derek Carpenters’s MATLAB script

Habitat Module (Extra)Systems - Habitat (Jason)

Optimization Code breakdown• Objective Function: Surface Area/Mass• Design Variables: Radius, Floor Height, Cylinder Length• Constraints: Minimum Ceiling Height, Pressurized Volume, Habitable Volume

Habitat Module Optimization (Extra)

Systems - Habitat (Jason, Derek)

Summary of habitat sizing options:

Habitat Module (Extra)

Summary of Sizing OptionsInner

Volume

[m3]

Mass

[Mg]

Inner

Inflated

Length [m]

Inner Radius

[m]

Packed

Length [m]

Packed

Diameter

[m]

Packed

Volume

[m3]

Total Cost

/ Module

[$]

Option 1 540.2 7.758 14.531 3.784 7.266 3.784 21.6

726,395.8

8

Option 2 767.01

10.36

6 19.150 3.835 9.575 3.835 28.8

963,738.4

0

Option 3 1209.1

15.56

3 28.730 3.841 14.365 3.841 43.3

1,445,479

.86

Systems - Habitat (Jason, Derek)

Habitat Module (Extra)Systems - Habitat (Derek)

WDM of Options Scale: 1-3

Weight

Option

1

Option 1

Weighted Option 2

Option 2

Weighted Option 3

Option 3

Weighted

Cost 0.1 3 0.3 2 0.2 1 0.1

Mass 0.4 3 1.2 2 0.8 1 0.4

Inner

Volume 0.2 1 0.2 2 0.4 3 0.6

Packed

Volume 0.3 3 0.9 2 0.6 1 0.3

Total 2.6 2 1.4

Airlock Module

● Safety: Fire and Depressurization Event evacuation

● Maximum Capacity of 12 persons in airlock○ Sets maximum module capacity

● Modular connection, multiple escape routes, ease of movement between modules

● Mass, Power, Volume○ Mass: 508 kg + 4 (305 kg) = 1.73 Mg○ Power: 0 kW (structural)○ Volume: 35.3 + 4 (21.2) = 120.1 m3

● Possible mass and volume reductions through design refinement

Systems - Habitat (Wyatt, George, Sarah)


Radiation Shielding● Requirements

• < 50 msv a year• 200 g/ cm3 of regolith

● Assumptions• Low levels of water ice• Density of 1.71 g/ cm3

● Regolith Layer• 1.165 m deep• Volume 634.59 m3

• Mass 1,085 Mg

Systems - Habitat (Chris, Seth)


Radiation Shielding SupportSystems - Habitat (Chris, Seth)

Cross-section of regolith support structure by Christian Mandrell

● Requirements○ Structure is independent of

inflatable habitat○ Withstand the load of 1.165m thick

regolith layer○ Safety Factor of 1.65○ Fits in Starship’s Payload

● Payload Requirements for Support

Mass 205 kg

Volume 0.412 m3

Colony LayoutSystems - Habitat (Lara)


Crew + Bath

Lab/Gym Farm

Food Prep/Dining

Crew + Bath

Crew + Bath

Crew + Bath

Payload Packing

● Starship payload model○ Habitats shall fit inside of Starship

payload bay○ Transport missions should fit 3x habitat

modules○ Transport missions should fit 3x airlock

modules and up to 12 connectors● Assumptions

○ Habitats compress by 50% in both radius and length for transport


CAD Model by Blake Gossage

TimelineSystems - Habitat (Sean)

Years 1-2 Year 5Year 3 Years 6-10Year 4

SMALL STEP GIANT LEAP NEXT STEP

Finish Site Prep

Initial Habitat

Habitat Complete

Full Crew

Start Site Prep

Habitat BuildupExpansion

All regolith is piled and ready to be bagged

by the crew

No Crew

3 Starship launches total for 3 habitat modules, 3

airlocks modules, and rigidization/furnishings

Crew of 5

1 Starship launch total for 1 habitat

module and rigidization/furnishings

Crew of 13

Habitat is now fully crewed

Crew of 19

Rovers will clear regolith and make piles that are to by bagged and

used for radiation shielding

No Crew

3 Starship launches total for 3 habitat modules, 3

airlock modules, and rigidization/furnishings

Crew of 9

Expand the habitat, capabilities, and crew as

needed

Crew of 19+

Phase BreakdownSystems - Habitat (Sean)

NEXT STEP

SMALL STEP

GIANT LEAP

● 1 Crew Module

● Farm Module

● Kitchen/Dining/Gym/Misc. Module

● 3 Airlocks

● 12 Connectors

● Rigidization Materials and Furnishings

● 2 Crew Modules

● 3 Airlocks

● 12 Connectors

● Rigidization Materials and Furnishings

● 1 Crew Module

● Rigidizations Materials and Furnishings

Structures

● Background○ Designed for 19 colonists○ Compared multiple initial options for structural habitat components

■ Had to reduce mass - Only inflatable not durable enough■ Hybrid structures were the “best of both” approach from rigid and inflatable

options

● Requirements○ The habitat modules must be transportable by Starship○ The habitat system must be sustainable for 10 years with regular

maintenance○ The habitat (both individually and as a system) must be able to support

regolith thicknesses defined by the Human Factors team○ The habitat must have 3 or more modules


Structures (Extra)

● Design○ Layer design - Liner (Nomex), Restraint (Vectran), Bladder (Mylar), Thermal

(Aluminized Polyester)

○ Regolith Cage - Semicircular arcs that cover the length of the habitat

○ (Mass power and volume covered previously)

○ Mass and Cost breakdown by material type:

Systems - Habitat (Derek, Jason)

Liner

Mass

(Mg)

Bladder

Mass

(Mg)

Restraint

Mass (Mg)

Thermal

Mass

(Mg)

Liner Cost

(USD)

Bladder

Cost (USD)

Restraint

Cost (USD)

Thermal

Cost (USD)

0.3317 0.02534 4.873 2.565 $21,265.29 $491.60 $670,037.50 $33,601.50

Power Generation and Storage

● Background

○ The system is designed as a solar array that will be supplemented by battery banks that will provide power to the colony

○ Preliminary power will be 100kW for roughly 10 colonists that will be expanded as the mission progresses to roughly 200kW for 19 colonists

● Requirements

○ The system shall provide power sufficient for habitat operation at all times.

○ The system shall maintain voltage levels within 5% of desired levels

○ The system shall store enough power to sustain habitat through lunar night

○ The system shall be modular and easily expandable

Systems - Habitat (Dylan Anderson & Noah Stockwell)

Power System Diagram● Power provided from solar

panels● DC Converter lowers voltage

to operating voltage (120 V)● DC Switching Unit handles

channel routing and fault tolerance between batteries and habitat

● Each module obtains power through relay and breaker switch

Systems - Habitat (James Pannullo)

Power Usage and Cost

● Power Cost with launch: $3.275B● Power Mass: 100.97 Mg● Power Volume: 326.4 m3

Systems - Habitat (Dylan Anderson & Noah Stockwell)

No Starship Case

In the event that Starshis is not available, the habitat design, crew size, and cost would be drastically affected due to the need to use smaller landers and launch vehicles


Category Value Change

Total Mass [Mg] 40 38.2% Reduction

Total Volume [m3] 1410 68.7% Reduction

Total Power [kW] 100 50.0% Reduction

Crew Size 7 63.2% Reduction

Total Habitat Cost [$ Billions] 13.5 957.0% Increase

Cost per Person [$ Billions] 1.93 2601.0% Increase

No Starship Case (Extra)

Same numbers as before on a per-habitat basis:


Category Value Change

Total Mass [Mg] per module 4 48.4% Reduction

Total Volume [m3] per module 141 73.9% Reduction

Total Power [kW] per module 10 65.0% Reduction

Crew Size (Total) 7 63.2% Reduction

Total Habitat Cost [$ Billions] per

module

1.35 675.0% Increase

Action Items

• Refine dimensions, mass, power, volume numbers for all modules• Consider raising the floor height and increasing port diameter so port is flush with the

floor• Explore redesign of airlock to reduce volume and mass, regolith• Develop floor plans for interior layout• Create CAD models of the interior layout• Integrate CAD models of life support, power generation and storage, and other

subsystems into habitat model• Finalize number of modules needed• Finalize timeline• Develop more animations and videos of the habitat and its functions• Model the support legs on the exterior of the modules• Model how crew will get from airlocks to lunar surface• FEA of the external habitat cage


Systems - Life SupportAAE 450 - Spring 2021

03/10/2021Team Lead: Marcus FullerTeam Members: Dylan Anderson, Aaron Baum, Wilson Barce, Blerton Ferati, Blake Gossage, Ray Huang, Riley Harwood, Sanjidah Hossain, Hunter Mattingly, Reid Trafelet, Marcus Fuller

Problem: Environmental control

• Background• The moon is an inhospitable environment, with extreme conditions not

suited for human habitation. Therefore, a highly controlled habitat environment is necessary.

• Constraints:• The system must maintain a breathable atmosphere• The design should close the mass loop to minimize recurring costs

• Assumptions:• NASA estimates are accurate• Life support requirements scale linearly with population

Systems - Life Support

Solution: ECLSS

The Environmental Control and Life Support System (ECLSS) is the umbrella term for the subsystems employed to maintain the atmosphere within the colony, as well as recycle other critical resources.


System Architecture

• Key Subsystems:• Oxygen Regeneration• Atmospheric Control• Waste Management• Water Management• Fire Suppression


Oxygen Regeneration Subsystem

Background:• Need a method to convert CO2 into a H2O

Solution:• Bosch Reactor• Bosch reactors have a closed mass loop, turning CO2 and H2 into H2O

and C• Other alternatives have a hydrogen deficit, resulting in mass losses over time.

• Mass: 1025 kg initially, 12.9 kg/day for filters• Volume: 1.07 M3


Pressure Storage Subsystem

Background• The pressure storage subsystem is responsible for maintaining air

reserves for both normal and contingency use for atmospheric control within a habitat or rover.

Constraints• System must contain the amount of air needed at any specific time

which is determined the number of crew, oxygen consumption per person, duration between oxygen resupply, and oxygen regeneration.

• System must satisfy any mass, volume, or reliability constraints required for a corresponding subsystem.


Pressure Storage Subsystem

Solution• ISS High Pressure Gas Tank

• Large tank, shielded for orbital debris• Mass: 544 kg Volume: 427.6 liters Pressure: 23.44 MPa• Recommended for initial habitat pressurization and contingency for failed

resupply.

• Orbital ATK Composite Overwrapped Pressure Vessel• Smaller tank, high reliability• Mass: 16.8 kg Volume: 87.0 liters Pressure: 31.00 MPa• Recommended for rover oxygen supply and contingency for depressurization

event


Trace Contaminant Control Subsystem

Solid contaminants will be captured using a two phase filtration system. Trace gas contaminants will be removed by the Oxygen Regeneration System, and additional filters.

Mass: 76.8 kgVolume: .5935 m3

All mass and volume calculations for contaminant control are per habitat.Trace gas filters must be replaced after 1-3 months, necessitating a 13.6-4.5 kg/month mass requirement.


Waste Management SubsystemBackground

• Due to the low gravity and hostile environment present on the moon, waste management systems will have to act differently from existing systems on Earth and in microgravities such as on the ISS.

Constraints• Low gravity and materials needed restrict the possible of traditional

plumbing and waste management systems.• System has to be hygienic to minimize spread of germs and pathogens• Created waste has to be removed from colony to prevent it from

creating a negative effect on the colonist and the colony as a whole.


Waste Management SubsystemSolution• Waste Digester

• Waste Digester can convert 0.035 liters of methane per dry gram of human waste

• Waste Digester requires ~0.0394 MW for daily operation

• Urine Recycling• Urine recycler can process 18.13 kg of

waste for a 24 hour period• Urine Recycler requires 424 W when

processing and 108 W when in standby mode


Water Recovery Subsystem

Background

• As more water that can be recovered, less water takes up launch mass and volume also reducing risk and increasing colony sustainability.

Constraints

• Water for life support, drinking, food, and hygiene needs to be pure enough for safe consumption.

Solution

• Mass: 350 kg• Power: 2.1 kW• Volume: 1.5 m3


Water Recovery SubsystemSystems - Life Support

Next Steps:● Generate consolidated

water model to project water recovered as a function of colony population.

● Estimate cost/budget● Design waste storage

facility● Develop plan for human

health research of recycled water and colonist urine.

Yes

No

Fire Detection and Suppression Subsystem

• Dual fire detection system - photoelectric sensors in ducts scanning for smoke particles and radiant heat sensors in ceiling to detect anomalous high temperatures in cabin (ISS Heritage)

• 3 stage fire suppression system based on aerospace industry suppression systems

1. Handheld Suppression with Halon 1211/Pressurized CO2 canisters (M/V TBD)

2. Halon 1301 Deluge system (M: 132kg/module, V: 0.126m^3/module)3. Gas Purge and cycling system (M/V TBD)


Problem: Thermal Control

• Background• Needed to absorb, transfer, and remove internal heat load from cabin air

and internal components• Driving Requirements:


Solution: ATCS


ATCS Specs


Calculations per module:• Total Mass: 0.25 Mg• Total Power: 41.76 W• Total Heat Load Rejection: ~16.5 kW• Faulted Loop Heat Load Rejection:

~8.2 kW

Problem: Physical Health

• Background• Due to the poor conditions for human habitation on the Moon, the

physical health of colonists is likely to deteriorate over the duration of a long term mission.

• Requirements:• Maintain physical health of colonists to the degree they are able to

perform tasks around colony• Solution is lightweight and compact

• Assumptions:• ISS human spaceflight measures apply to lunar colony


Solution: Crew Standard Measures

• Collection of physical, physiological, and mental measurements that encompass human spaceflight and lunar risks

● Samples to analyze: blood, urine, stool, saliva, body swabs, questionnaire responses

● Measurements taken preflight, postflight, and during the mission● Full equipment listed in Master spreadsheet


• Sleep questionnaires• Actigraphy• Metabolic panels• Cognition tests

• Immune system response• Microbiome measurements• Heart ultrasounds• Sensorimotor tests

Exercise Subsystem

• The exercise equipment available to colonists includes a treadmill, squat rack, resistance machine, and cycling machine.

• Equipment is modeled after ISS exercise equipment to be effective in lower gravity.

• Recreational sports to be developed further: VR 360 treadmill, TVs for gym, small manual rovers for racing, crater skiing


CAD Models courtesy of Sam Conkle

Problem: Medical Equipment

• Background• The trip to the lunar surface as well as life on the colony can make the

crew more susceptible to medical issues that are more difficult to treat away from Earth.

• Constraints:• Medical items such as pills and instruments shall be present on trips

between Earth and the lunar colony.• There shall be an adequate stockpile of medical supplies on the lunar

colony to address any medical issues that may arise for the colonists.• Assumptions:

• The medical supply list on the ISS is scalable to the size of the colony.


Solution: Medical Equipment MasterList

• Full Medical Equipment MasterList can be found under the master equipment list in the drive (over 130 medical items listed)

• MASS AND VOLUME CALCULATIONS (cost to be evaluated next)

Systems-AgricultureAAE 450 - Spring 2021

03/09/2021Team Lead: Nicky MassoKyle Alvarez, Jaxon Connolly, Wellington Froelich, Reece Hansen, Monica Viz

Planned Design• Aeroponics• 24 towers/module• 105 potatoes per week per module• 256 other plants per week per module• Lights on for 16 hr/off for 8 hr• Control sensors use negligible power• Bring crystallized 10-10-10 NKP salt mix (0.4 kg/month)

Systems - Agriculture (Nicky Masso, Kyle Alvarez, Wellington Froelich)

Subsystem Mass of System Power (kWh / day) Volume (m3)

Watering 2.271 Mg 70 0.82

Lights 0.612 Mg 32 0.243

Why Not Just Bring FoodSystems - Agriculture (Nicky Masso, Kyle Alvarez)

AeroponicsSystems - Agriculture (Wellington Froelich)

[1]

[2]

[1] & [2] Tower Garden Seedlings | Aeroponic Seedlings | Tower Farm

(truegarden.com)

https://truegarden.com/

What Crop Should we Grow?

• Goal to get as many calories grown per week• Choose crops with high Kcalories/gram

• Beans, wheat, potatoes

• Choose potatoes due to prior research on yields rates• 119 grams per plant [1]

• 118 calories per plant• The rest are for personal use/research

• no significant impact on calories harvested

Systems - Agriculture (Kyle Alvarez)

[1] Farran, I., & Mingo-Castel, A. M. (2006, January). Potato minituber production using aeroponics:

Effect of plant density and harvesting intervals. https://link.springer.com/article/10.1007/

BF02869609

Production Schedule 5 Week Ex.Systems - Agriculture (Wellington Froelich)

Legumes Leafy Greens Squash Vine Produce Fruits

Black beans, green

beans

lettuce, kale, chard butternut, zucchini tomatoes, peppers strawberries, currents

Estimation of Potato Yields

• 12390 calories per week per agriculture unit• 105 potato plants in total

• About 77 % of one person’s weekly caloric needs• Enough for about 13.5 dinner meals per week

• About 10.18 kg mass savings per week per Agriculture unit• Volume savings of about 0.099 m3

Systems - Agriculture (Kyle Alvarez)

Data TablesName and Group/Section #

Grams per Plant Calories per plant Plants harvest

per week per

tower

Calories

harvested per

week per tower

Calories per

dinner estimation

Mass per dinner

estimation

119 118 11.1 538.7 920 0.756

Total volume of

packaged food for

one person for 10

days m3

Average volume for

one day m3

Assuming dinner

counts for a third of

volume, volume of

a average dinner m3

Total dinners saved Volume savings per

week m3

0.219 0.0219 0.0073 13.5 0.098

Structure and Layout• Nicky talks about our planned system’s hardware• dimensioned top/side/iso views in solidworks• closeup of plant holder

Systems - Agriculture (Nicky Masso)

Required Supplieslights, pumps, hoses, seeding area+supplies, tools

Systems - Agriculture (Nicky Masso)

Control SystemWith the agricultural system, the control system must include:• A temperature regulator - keeping the temperature at roughly 70 degrees F or

21 degrees C

• A humidity collector - collects the excess water vapor and converts it back to usable water for the aeroponics system

• A timed spraying schedule - for every 5 min, there will be a 15 second spray of nutrients on to the plant

Systems - Agriculture (Reece Hansen, Monica Viz)

Control System• Basic Nozzle Spray Simulink

Systems - Agriculture (Reece Hansen, Monica Viz)

• Basic Temperature Simulink

Systems Research FacilitiesAAE 450 - Spring 202103/09/2021Preliminary Design ReviewTeam Lead: Emily Anderson

Team Members: Joong Hyun Pyo, Cody Martin, Roberto Cardano, Chad Blackwell, Riley Harwood, Blerton Ferati, Jarod Utz

Science Objectives: Lunar Environment• Determine global density, composition, and time variability of lunar

atmosphere before it is perturbed by further human activity

• Solution: Observation satellite mass spectrometer

• Determine the composition of the lunar surface specifically looking for water and other useful natural resources

• Solution: Shortwave Infrared Laser Reflectometer

• Determine compositional state and distribution of the volatile component in lunar polar regions

• Solution: Sample collection, drilling, Mini-RF & satellite imaging

• Determine the characteristics of the lunar dust environment and assess effects on lunar exploration and astronomy

• Solution: Sample collection, drilling, Lunar Dust Analyzer

Scientific Objectives: Human Health

• Understand the short term and long term biological and physical effects of the lunar environment on the human body

• Solution: Vital sign monitoring, crew medical sample collection

• Understand the effects of the lunar gravity on human performance

• Solution: Bone densitometry, muscle mass measurements

• Understand the mental and cognitive effects of long-duration missions

• Solution: Crew Standard Measures

Timeline

Years 1-2 Year 5Year 3 Year 4

Lunar Atmosphere

MeasurementsOnce observation satellite is built, mass

spectrometer will begin atmospheric

measurements prior to human arrival

No Crew

Begin Crew Standard

MeasuresAs humans arrive, scientific objectives are

focused on gathering health data and

beginning lunar gravity research

Crew: 5 people

Research Lab HabitatOnce the lab habitat is finished, begin

geological sample analysis and lunar

surface exploration missions. Continue and

refine human health research with

additional lab space

Crew: 13 people

Science Objective Adjustment

and RefinementLong-term health effects of crew, development of

new science objectives, refining previous goals

based on new data

Crew: 19 people

Instruments: Satellites

● Mass spectrometer○ Mass: 8kg Power: 10W Volume: 0.002 m3

● Mini-RF○ Mass: 14kg Power: 25-50W Volume: 0.0396 m3

● Camera○ Mass: 8.2 kg Power: 6.4 W Volume: .189 m3

● Shortwave Infrared Laser Reflectometer○ Mass: 14kg Power: 15-73 W Volume: 2U

Instruments: Rovers

● Mass Spectrometer○ Mass: 13kg Power: 10.4 W Volume: 0.043m3

● Drill for Sample collection○ Mass: 41kg ○ Power: 30-40Watt avg. Max 200W for longer 5-10 minute drill duration○ Volume: 0.25m3 for 1 meter drill

● Neutron Spectrometer (Determines distribution of H)○ Mass: 1.4 kg Power: 5.2W Volume: 0.00129m3

● Surface Imaging○ Mass: 1.1 kg Power: 11 W Volume: 0.00383m3

Instruments: Habitat● Bone Densitometer

○ Mass: 10kg Power: 66W Volume:0.03255m3

● Centrifuge○ Mass: 8kg Power: 50W Volume: 0.0219m3

● Cardiac Ultrasound○ Mass: 7.3kg Power: 265-636W Volume: 0.01264m3

● Actigraphy Watch○ Mass: 0.031kg Power: 9.3W Volume: 90cm3

● Blood sample equipment○ Mass: 31.73kg Volume: 0.1213m3

● High speed camera○ Mass: 4.5kg Volume: 0.004997m3

● Computer (ThinkPad laptop)○ Mass: 1.75kg Power: 57W Volume: 0.00173m3

Science Traceability Matrix

Next Tasks

• Floor plan modeling

• Refining scientific objectives

• CAD designs

• Cost & Budgeting

Vehicles - SatellitesAAE 450 - Spring 2021

10 March 2021Lead: Sammy DickmannTeam Members and Contributors:

Lorin NugentJaxon ConnollyMonica VizStephen GrabowskiMatt Popplewell

Greg FrettiAlan GelmanNicky MassoEthan WollerAnna Cismaru

Dale WilliamsRoberto CardanoSanjana Singh

Objectives• Communications satellites:

• Enable communication between the Moon and the Earth• Enable communication among the Lunar colony• Enable communication between the observation satellite and the Earth

• Observation satellite:• Take optical images of the colony and other areas of interest• Complete research objectives

• Collect RF data to map lunar volatiles for research • Collect and analyze atmospheric samples

Vehicles - Satellites

RequirementsVehicles - Satellites

Requirement ID Requirement Description Measurement

CTRL-1 The satellite optical camera shall have a nominal GSD of 0.5 m. GSD

CTRL-2, COM-1

The observation satellite shall be capable of downlinking 4 2.5 km x 2.5 km images

per pass. Bps

CTRL-3 The optical camera shall have a swath of 2.5 km. Pixels*GSD

CTRL-4

The observation satellite altitude variation while taking images shall not exceed +/-

15 km. km

CTRL-5 The pointing accuracy for the observation satellite shall be 60 arcsec. arcsec

CTRL-6 All satellites shall have a minimum expected lifetime of 15 years. years

CTRL-7

The MiniRF shall probe lunar regolith on S and X Band frequencies at 2380 Mhz (±10

Mhz) and 7140 Mhz (±10 MHz), respectively. MHz

CTRL-8 The MiniRF shall have a swath of 6 km for S band and 4 km for X band km

CTRL-9 The MiniRF shall have a zoom capability of 15 m x 30 m m

CTRL-10 The mass spectrometer shall be capable of storing 100,000 ions. ions

CTRL-11 The mass spectrometer shall extract ions at a rate of 2kHz kHz

CTRL-12

The mass spectrometer shall have a mass resolving power of m/Δm = 46,000

(FWHM) Full width at half maximum (FWHM)

+additional linked research requirements

Assumptions• Possible to develop and manufacture all satellites simultaneously• Accelerated development, manufacture, and test timeline of 2 years• Largest design assumption: components are compatible

• No interference• No field-of-view keep-out zones• Same communications protocol• Thruster impingement negligible

• Operations costs during launch year are not prorated


Concept of OperationsVehicles - Satellites

ScheduleVehicles - Satellites

Years 1-2 Years 4-10Year 3

Design, develop, manufacture, and test all satellites● Year 1: Design and

development● Year 2: Manufacture

and test

Launch and commission satellites● 2 communications

satellites per launch● Observation satellite

launched after at least one comms launch

Nominal operationsDesigned satellite lifetime: 15 years

Budget and CostsVehicles - Satellites

• Costs related to satellites change throughout mission timeline• Development, manufacture, and test costs vs operations costs

Years 1-2 Years 4-10Year 3

Communications: $12 million + launch costs $12 million per year$150 million per year

Observation: $30 million per year Launch costs Included

Total: $180 million per year $12 million + launch costs $12 million per year

Dimensions and SizingVehicles - Satellites

Communications Observation

Bus (m) 3.1 x 2.4 x 1.5 1.7 x 1.5 x 1.1

Solar Panels (m) 1.22 x 3.66 1.96 x 5.87

Communications Observation

Current Best Estimate Max Expected Value (20%) Current Best Estimate Max Expected Value (20%)

Dry Mass (kg) 577 692 287.1 344.6

Wet Mass (kg) 1107 1329 371.0 445.2

Power (W) 973 1168 2980 3576

CAD Model - CommunicationsVehicles - Satellites

Ka-Band antenna

Tanks

RWAs Sun sensor

X-Band

S-Band

ACS thrusters

CAD Model - ObservationVehicles - Satellites

Mission Design -Communication Constellation


• Walker Constellation configuration was chosen for:• Stability• Consistent contact with south pole• Large surface coverage• Potential for expansion

• Choose 6 satellites in 3 orbital planes• Continuous and often redundant contact with south pole ground station• Cross-link capabilities between satellites• Stable - less than 100 m/s per year per satellite for stationkeeping

Element Value

a 3238.2 km

e 0.0

i 80°

RAAN 0°, 120°, 240°

Mission Design -Observation Satellite

• Quasi-frozen lunar orbit at 50 km x 216 km chosen• Periapsis and apoapsis variations naturally

bounded• Won’t eventually impact the Moon like a

circular orbit• Modeled with Moon J20,20 harmonics, Earth,

Sun, and SRP

• Requirement to stay within +/- 15 km of nominal altitude when taking images• Used as bounds for periapsis and apoapsis• Burn targeting no more than once every 13.5

days (half of a lunar sidereal period)

• Yearly delta-v budget of 139 m/s (with impulsive maneuvers)• Not optimized, can be brought down further

Communication Infrastructure• 3 Frequency bands enabled

• Ka Band: • For transmission of “dead” data (unchanging) to Earth at the White Sands NEN Ground

Station• 0.97 m dish antenna, 444 W transmitter power, 32 Mbps max data rate

• S band• For crosslink with Observation satellite, uplink from Earth, uplink from Moon• patch antenna on the satellite

• X band• For downlink to Earth, downlink to the Moon• patch antenna on satellite

• Allowing the communication satellites to have S and X band capability enables voice communication with Earth before ground station is set up• After the ground station is set up, live voice and video will be relegated to that,

and satellites will only be used for voice and video comms in emergencies


Communication System - Communications


Communication System - Observation


Attitude Control System• Reaction Wheel Assembly comprised of 4 reaction wheels• Mass/Power/Volume is based upon the required torque,

momentum storage, and power requirements


Based on Collins’ RDR 23-0 RWA

Based on LRO’s RWA designed by NASA Goddard

Communications Satellites

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

27.8 360 0.0110

Observation Satellites

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

52.8 500 0.1001

ACS Thruster Layout - Communications


Torque Axis

Thruster #1 Thruster #2

+X 3 4

-X 7 8

+Y 2 6

-Y 1 5

+Z 5 6

-Z 1 2

ACS Thruster Layout - Observation


Torque Axis

Thruster #1 Thruster #2

+X 1 2

-X 7 8

+Y 3 4

-Y 5 6

+Z 4 6

-Z 3 5

Observation Payload• Camera

• Use: observation and documentation of colony health and progress• Near Angle Camera with 0.5 m resolution

• MiniRF• Science objective: map lunar topology for exploration and mining• Technology demonstration on LRO (failed relatively soon into lifetime)

• Mass Spectrometer• Science objective: determine atmospheric composition and effect of human

habitation• Based off MASPEX on Europa Clipper


Camera MiniRF Mass Spectrometer

Mass (kg)

15 13.8 8

Power (W)

20 25 25

Power Systems• Solar arrays used are DSS Roll-Out Solar Arrays (ROSA)

• Following a 3:1 ratio of length to width• BOL power requirements were calculated based degradation and correction

factor • Correction factor of 2.1 used for eclipse• 0.5% performance degradation per year• 7.24% reduction in performance over 15 years


Satellite EOL Power Requirement (kW) BOL Power Requirement (kW) Dimensions of Solar Arrays (m)

Communications 2.4906 2.6851 1.221 x 3.664

Observation 6.3882 6.887 1.953 x 5.868

Thermal Systems• Satellite will primarily utilize a passive thermal system

• Silver teflon coating on exterior and interior of satellite for protection from solar radiation

• Excess heat radiated from plate coated with SG121FD white paint• Excess heat conducted to passive radiator (plate) via thermal straps

• Some components may require targeted systems• Sensors need to be cool -> radiators• Hydrazine can’t freeze -> heaters


Propulsion Systems• Communications satellite will use full chemical propulsion for stationkeeping,

attitude control and momentum unloading• Hydrazine monopropellant and Helium as pressurant• 2 propellant storage tanks of 0.88 m dia are to be used• The pressurant tank is composite overwrapped pressure vessel (COPV)

as used on LRO• 636 kg of propellant (includes 20% extra) will be carried

• The observation satellite will use electric propulsion for orbit changes and chemical propulsion for attitude control/momentum wheel dumping• Electric propulsion uses a system based on the NSTAR xenon thruster,

used on the Dawn and Deep Space 1 space probes.• Hydrazine used for attitude control• 20.2 kg Xenon supercritical liquid, 63.8 kg Hydrazine


Action Items• Launch vehicle selection for both types of satellite (+ launch schedule)• Find solution to excessive ACS propellant mass on the observation satellite• Engine and nozzle sizing and CAD model on satellite• Reduce contingency to 0% and recalculate for CDR• More detailed thermal control system, especially regarding hydrazine storage

and instrument requirements• Dynamic model of the satellite attitude


PDR - Rover Systems

Team Lead: Vishank Battar

Team Members: Cody Zrelak, Ajay Chandra, Alan Gelman, Nick Johnson, Omtej Vallabhapuram,

Joseph Seiter, Chase Neff

Problem: Area Preparation

Description:

We need to have the rough surface of the moon prepared for colonial activities. Examples of areas that need to be prepared are:

• 90 m2 per habitat

• Habitat

• Solar Field

• Launching/Landing Pad

• Recreational Area

Requirements

• System ready before mission start date

• Area Clearing completed by a rover that is ready before humans arrive

• Autonomous

• Designed to flatten large areas for habitats

• Solar panel powered

• Can function without satellite based GPS

• Can clear 90 m2 area per habitat

• Fit into 1U

• Last for 10 years

Solution: Clearing Rover

•Regolith Clearing Rover

•Total Mass: 543 kg

•Total Volume: 10 m^3 (Fits in “1U”)

•Max power: 1.073 kW

•Clearing Rate

•Assuming a rover speed of 0.045m/s

•Area Clearing Rate: 486 m^2/hr

•Features•Mostly parts from the Collection Rover

•Auger for flattening land

•Plow for moving and piling

•Rocker Bogey Suspension

•3m Foldable Auger Arm

General Clearing Rover Design

Part Mass and Material Chart

Next Steps

• Calculate the Power needed• Time to complete a task on a single charge

• Calculate Cost Estimation

Problem: Gathering Regolith for Habitat Shielding

Description:

Regolith is needed for the habitats group for shielding and needs gathered before humans arrive to set up the habitats quickly. Regolith is also needed for in-situ water extraction.

• 174 m3 of regolith per habitat for shielding

• 0.81 m3 of regolith per person per week

Requirements

• System ready before mission start date

• Autonomous

• Can gather regolith at needed collection rates

• Can bring all needed shielding regolith to landing site

• 174 m3 per habitat

• Can function without satellite based GPS

• Fit into 1U

• Lasts for 10 years

Solution: Collection Rover

•Regolith Collection Rover•Total Mass: 490 kg

•Total Volume: 10 m^3 (Fits in “1U”)

•Max power: 936.5 W

•Maximum Bay Size: 5 m^3

•Collection Rate•Assumption: Half the bay can be filled

•2.5 m^3 regolith per trip

•Using 174 m^3 regolith per habitat

•70 trips per habitat

•Features•Bottom Bay Doors (Dump Regolith)


•Pneumatic Elevator Collector

General Collection Rover Design

Part Mass and Material Chart

Next Steps

• Calculate the Power needed• Time to complete a task on a single charge

• Calculate Cost Estimation

Problem: Resource Collection

Description:

We need to process raw materials into usable materials for:

• construction of habitats

• water extraction

• agricultural purposes

• maintenance and repair

Requirements

• Autonomous

• Large storage capacity for transporting mined extracts back to base.

• Storage bay must be able to hold any kind of material and be thermally isolated.

• Must be compatible with modular attachments (robotic arm, drill, etc.)

• Must be stable with fully loaded bay on steep inclines (Shackleton crater angle)

• Must be stable with fully loaded bay on uneven surfaces

• Must be operable using minimal solar power.

• Must keep dust out of moving parts

• Must provide for redundancy of mining drills.

Solution: Mining Rover

• An autonomous mining rover that is able to successfully maneuver the lunar surface around shackleton crater and mine for raw materials such as basalt and ice. It will complete this objective with the use of of a robotic arms for beacon placement, a fluted drill for resource extraction, an archimedes screw for resource intake, and a storage bay for resource storage that prioritize rover stability and space efficiency

• The payload bay is leak-proofed and thermally isolated to prevent mined ice from possibly melting and hampering electronics performance.

Mining Rover Drill Weighted Decision Matrix

Conclusion: A core drill is the best choice for the drilling system on the mining rover

Sizing the Drilling System

• The upper limit for the drill area will be based on the maximum power available to the drill• The overall power available for the rover is 1 kW and we have designated 75% for mining operations (0.75

kW)• Height of the drill needs to be less than 2.5m b/c this is largest dimension of payload bay

Calculations for Drill Sizing

Mining Rover Intake System

Due to the high rate of collection needed for practical use basalt mining, the capabilities of the fluted drill and archimedes screw designs to provide sufficient basalt are limited. Hence, despite the advantages of a fluted drill in sample maneuverability, intake and storage, it is impractical for use in our mining rover intake system.

Conclusion: Utilizing the robotic arm attached to the mining rover to collect basalt cores mined from the ground is the intake system that will be required to make the more collection efficient core drilling method effective.

Mining Rover Next Steps

• Further solidify power requirements based on numbers for robotic arm• Begin cost estimation• Consider solar panel placement/power availability• Finalize drill/core collection sizing

• Easy sizing for collection and storage vs. collection rate

Problem: Location Searching/Observation

Description:

It will be necessary to have a vehicle that can look for and find locations for resource collection and other science objectives:

• Find high concentration Lunar Basalt Sites

• Find potential lava tube locations

• If radio communication is lost with a mission group, can quickly find them

• Scout for autonomous mapping purposes

Solution: Scouting Rover•Total Mass: 502.9 kg

•Total Volume: 5 m^3

•Max power: 783.6 W

•Features•Camera Imaging system (lunar regolith composition)

•Ground-penetrating Radar (water ice detection)


•Rotary Percussive Corer Drill (sampling)

Problem: Need a method to construct habitats and other structures

Description:

Colonists on the moon will be constructing habitats, solar farms, fuel depots, agricultural hubs, etc. They will need to conduct various construction maneuvers that will require specialized tools

Solution: Construction Rover

Activity Tool

Dig Backhoe

Reach high up Crane/ hooks/ attachments

Fine placement of material Robotic arm

Drill Drill

Demolish Wrecking ball attachment

Extra - CAD (Crane)

Extra - CAD (Arm)

Extra - CAD (Modular Rover w/ Crane)

Problem: Handling external objects/interchanging tools

All kinds of rovers will need to maneuver external objects and perform different actions such as:

• Drilling

• Sampling

• Lifting

• In situ analysis

Solution: External Robotic Appendage (ERA)

• Scalable in Design

• 6 Rotational D.O.F.

• “Hand” will consist of swappable attachments (rover dependent)

• Modularly attachable.

• For large scale design, must support power lifting.

• Generalized control software, must be compatible with all rovers.

• Must be autonomous and allow for manual override.

Required sensors include gyroscopic sensors, IR

sensors, imaging. More on the hand depending on

rover.

Extra - Control

Rover Trajectory Control Simulation

- Autonomous control of lunar rovers will require a robust control system to accurately follow predetermined paths and commands

- Start by analyzing rover wheel dynamics and vehicle kinematics, input desired constant forward velocity and track changing heading angle (testing with following an arbitrary circle ie. clearing rover function)

- Lunar surface parameters as well as vehicle mass, power, volume are variable in code, higher fidelity model will come from adding slip ratio estimation to feedback loop

- Working on introducing basic terrain complexity (slope/incline)

% Rover Trajectory Control Testing

% The goal of this program is to formulate basic dynamics and kinematics

% governing the motion of a rover on the lunar surface, as well as to test

% basic control strategies for commanding desired motion

clc;

%% Initialization of Parameters

% Vehicle Parameters

% (Assuming every wheel is stiff, and the rover itself is a rigid body

% these parameters will specify the physical characteristics of the rover)

m = 500; % total mass of body (kg)

m0 = m / 2;

n = 6; % number of wheels

l_r = 9.4; % rover length (m)

w_r = 5.5; % rover width (m)

h_r = 4.3; % rover height (m)

l_w = 6; % rover wheel base (m)

% Mars Spring Tire 2 parameters

r_t1 = 0.2667; % outer tire radius (m)

r_t2 = 0.2159; % inner tire radius (m)

w_t = 0.2032; % tire width (m)

m_t = 5.41; % tire mass (kg)

I_t = 0.5 * m_t * (r_t1^2 + r_t2^2); % tire moment of inertia (kg m^2)

p = 0.25 * 745.7; % individual drive motor power (Nm/s)

vmax = 13 * 1000 / 3600; % maximum forward velocity

Tmax = p / (vmax / r_t1); % maximum available applied torque (Nm)

% Lunar Parameters

g_m = 1.6; % Moon gravitational acceleration (m/s^2)

mu0 = 0.60; % Lunar regolith static friction coefficient

mu = 0.55; % Lunar regolith dynamic friction coefficient

R = 360 / 2; % desired circular path radius (m)

%% Wheel Dynamics / Rover Kinematics Derivations

Fr = mu * m_t * g_m; % wheel motion resistance force (N)

Tr = r_t1 * Fr; % rolling resistance torque (Nm)

% desired wheel angle

%del = atan((l_w/2) - R*cos(alpha.data)) ./ (R*sin(alpha.data) + (w_r/2));

%% Plotting

figure(1)

plot(Xtrack.data/R,Ytrack.data/R,'b')

hold on

plot(X.data/R,Y.data/R,'--r')

title('Lunar Rover Circular Trajectory Tracking')

legend({'Desired','Actual'},'Location','northeast')

xlabel('X/m')

ylabel('Y/m')

axis([-1.2 1.2 -1.2 1.2])

hold off

Extra - Power & Thermal

• An active thermal system will be used to generate heat in cabin for crewed rover

• A passive system will bring excess heat to cabin from external components

• Solar arrays used will be DSS Roll-Out Solar Arrays (ROSA)

• Sizing follows a 3:1 length to width ratio

• BOL Power requirements are calculated based on degradation over a period of ~15 years

• 0.5% performance degradation per year

• 7.24% total degradation over 15 years

• Only one side of rover will be illuminated at a time, so three arrays are necessary

• One on each side, and one on top with tilting capabilities

• Two solar arrays must be able to generate enough power for rover to function

Rovers EOL Power Generation (W) BOL Power Generation (W) Dimensions of Solar Arrays (m)

Clearing & Collection 1000 1078 0.77 x 2.32

EVA vs. Rover Analysis

• Due to the inherent risk and complexity of EVAs, it is important to minimize EVA time for the inhabitants.

• Risks include dust contamination, suit failure, personal injury and exhaustion.

• For each potential task to be completed, several aspects must be considered. Assuming the task must be completed, the question is whether the potential increase in rover complexity/programming is the worth the added risk for an EVA mission.

• Due to increases in rover ability and the limitations of EVA suits, there are few tasks that can be completed by a human that cannot be completed by a rover equipped with a complex robotic arm.

• For some tasks, a human would be required. These tasks include fixing rover failure, reaching areas that the rovers are not able to and rover maintenance. However, the rover related tasks could potentially be fixed via a pressurized rover garage.

• While EVA suits are clearly still necessary due to potential emergency use, it is recommended that rovers are used for nearly every task, as the increased complexities of the rovers do not outweigh the added complications of using humans. This analysis is shown on the next slide.

EVA vs. Rover Analysis

Vehicles - InteriorAAE 450 - Spring 2021 03/10/2021Preliminary Design Review

Team Member: Joong Hyun Pyo, Sanjidah Hossain, Mackenzie Marks

Purpose

After the initial phase of colonizing the Moon, we want to provide the colonist with addition non-human support from research/medical purposes to trivial maintenance of the habitat for an extended period of time

• Requirements• Heavy load lifting

• Average load bearing: 2.5-3x Human body weight• Precision Control

• Medical Procedure • Replacing Components, Electrical work

• Additional Habitat Control/Maintenance• Autonomous Experiments• Cleaning

• Assumptions• Fully established colony with proper power, life support, and temperature

control

Vehicles - Interior

Problem: Heavy Load Maneuvering

• Background• Despite low gravity conditions, heavy loads will be difficult to move

without the assistance of a machine• Space conservation requires stacking, thus the ability to reach heavy

resources stacked on top of each other safely is necessary• Constraints

• Limited space for operation within habitat• Limited power available

• Assumptions• Resource storage facilities concentrated in one section of the habitat• Resources organized and labeled in a systematic manner to allow for

automation

Vehicles - Interior

Solution:Heavyload Robotic Arm

• Mass = 350 kg (Body: 100 kg, Arm: 250 kg)• Power = ~3kWh• Volume = Body: 1 m x 1 m x 2 m (BxWxH), Arm: 0.5 m X 1 m (RxL)• Speed = 0.5-0.8 m/s• Features

• About 8 degrees of freedom on arm, additional 3 degrees of freedom on wheels and 2 degrees of freedom on forklift features

• Automated storage and retrieval• Capable to lifting up to 230 kg, able to reach max height of 3.5 meters• Capable of maintaining and updating storage inventory• Potentially provide maintenance assistance for interior subsystems

Vehicles - Interior

Heavyload Robot Initial Concept

Figure 1: Top View

Vehicles - Interior

Figure 2: Side View Figure 3: Arm Side View

Problem: Mobility & Dexterity

• Background• Given the size of the habitat, the inside will be crowded with human

activity and materials. Therefore, the robot requires minimal space while being able to perform both trivial and complicated tasks.

• Constraints:• Limited space in interior of habitat• Unexpected activity/collisions during maneuvers between stations.

• Assumptions• Habitat has reflective tape/lines along the floor• A.I. Capable Technology

• Object identification and basic problem solving• All areas of materials are organized

Vehicles - Interior

Solution: Assistant Robot

• Mass: ~204kg• Power: ~2kWh (60 Minute Battery Operational Time)• Volume: 1.31 m3

• Speed: 1.5 m/s

• Features• Medical Procedures• Autonomous Experiments• Habitat Maintenance

Vehicles - Interior

CAD ModelVehicles - Interior

Features:● cameras in the ears and eyes● legs lift● arms rotate about spheres● torso bits rotate about cylinders

so the vehicle can get taller and shorter

● hands work like human hands with full wrist and finger mobility

● head rotates about the neck piece

Model By: Mackenzie Marks (CAD Team)

General Design Schematic

Camera

SensorsTriple camera setup(IR,2

Stereo) on the head of

Robot to guide and

analyzes objects.

A.I. Capable Determination

The Robot will make a decision based

on preset knowledge and determine if

human operation is required.

Task Assignment

Operation Team

determines then

assigns the task to

Robot

Human Operation

Manual take over of

Robot will allow a

team of experts to

perform the specific

task.

Autonomous

OperationIdentifies workspace and

required materials for set

task.

Materials collection

phase from

workstations.

Collection

Docking/Chargin

g Station

Robot lingers idol.

Executes task based on

programmed knowledge

and material identification

from sensors.

Execution

Manual

ExecutionEither on-site operational

team can perform, or off-

site experts.

Vehicles - Interior

Timeline

• Given the low lunar gravity, we expect not much use of internal vehicles/robots in the early phases• HeavyLoad Robotic Arm will come to play once objects with more mass

are brought in-doors• Requires larger habitats

• Use of robots for automation and maintenance is a luxury item/task• Initial phases of the colonization will require extensive care from

crewmates and can’t be relied on robots• Requires sustainable power, room for error

• Recommendation• Post 10 year project for Lunar Colonization

preliminary design review - engineering.purdue.edu

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