Preliminary Design Review - engineering.purdue.edu
Transcript of 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 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
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
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
Translunar Trajectories for Communication Satellites
• 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
Translunar Trajectories for Communication Satellites
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
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
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)
CAD Model by John Matuszewski
Overview (Extra)Systems - Habitat (John)
CAD Model by John Matuszewski
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
Systems - Habitat (Derek)
CAD Models by John Matuszewski
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)
CAD Models by John Matuszewski
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)
CAD Models by John Matuszewski
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)
CAD Model by John Matuszewski
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
Systems - Habitat (Derek)
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
Systems - Habitat (Derek)
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: $3.275B● Power Mass: 100.97 Mg● Power Volume: 326.4 m3
Systems - Habitat (Dylan Anderson & Noah Stockwell)
No Starship Case
In the event that Starshis is not available, the habitat design, crew size, and cost would be drastically affected due to the need to use smaller landers and launch vehicles
Systems - Habitat (Sean)
Category Value Change
Total Mass [Mg] 40 38.2% Reduction
Total Volume [m3] 1410 68.7% Reduction
Total Power [kW] 100 50.0% Reduction
Crew Size 7 63.2% Reduction
Total Habitat Cost [$ 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:
Systems - Habitat (Derek)
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 - Habitat (Sean)
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.
Systems - Life Support
System Architecture
• Key Subsystems:• Oxygen Regeneration• Atmospheric Control• Waste Management• Water Management• Fire Suppression
Systems - Life Support
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
Systems - Life Support
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.
Systems - Life Support
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
Systems - Life Support
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.
Systems - Life Support
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.
Systems - Life Support
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
Systems - Life Support
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
Systems - Life Support
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)
Systems - Life Support
Problem: Thermal Control
• Background• Needed to absorb, transfer, and remove internal heat load from cabin air
and internal components• Driving Requirements:
Systems - Life Support
Solution: ATCS
Systems - Life Support
ATCS Specs
Systems - Life Support
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
Systems - Life Support
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
Systems - Life Support
• 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
Systems - Life Support
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.
Systems - Life Support
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)
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
Vehicles - Satellites
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
Vehicles - Satellites
• 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
Vehicles - Satellites
Communication System - Communications
Vehicles - Satellites
Communication System - Observation
Vehicles - Satellites
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
Vehicles - Satellites
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
Vehicles - Satellites
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
Vehicles - Satellites
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
Vehicles - Satellites
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
Vehicles - Satellites
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
Vehicles - Satellites
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
Vehicles - Satellites
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
Vehicles - Satellites
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)
•Rocker Bogey Suspension
•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)
•Rocker Bogey Suspension
•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