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Group One Presentations

January 31st, 2019

Preparing to Move to a Solution

Andrew LangPM

January 31, 2019

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Problem: Interdependent Lower Level Requirements

● Systems are all interdependent.● There is a large volume of work completed

that must be documented shared and integrated.

● Numbers and requirements paint a limited picture.

Solution: Get Data and Format Report

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Technical Progress● Gather “Big” data● Lower level requirements delivered: 70-80*● Tentative total mass sent from Earth: .585 Mg**● Total number of systems: 14

Management Progress● Sections of report proposed● Plan to use report as a “living document”● Final report will be composed in Latex● Automatic labeling and referencing

*As of 4:10 PM 01/30/2019**As of 5:35 PM 01/30/2019

BACKUP: Proposed Report Sections

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BACKUP: Overleaf Template

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BACKUP: Draft Project Metrics

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BACKUP: Current Action Items

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Landing Site Prep - Analysis of Methods to Prepare

Lunar Surface Andres Carrillo

CADMajor Launch Vehicle, Communication Satellites, Ice Harvesters,

Landing Site PrepJanuary 31, 2019

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Problem

Landing site prep involves a system to prepare the lunar surface for construction and safe landing. Complications or problems with the lunar surface:

- Dust/Lunar regolith particles- Meteorites- Not level/uneven land- Potential need to dig or build under surface of moon- No existing structures, road system, landing/launch pad, etc.

Excavation, Leveling, and Construction techniques/machines:- Explosives- Excavators- Bulldozers- Plows- Loaders

Solution

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Use a combination of excavation explosives and machines/vehicles to prepare the land site.● Explosive should be ANFO (Ammonium Nitrate/Fuel Oil)● Excavator Heavy Duty Bucket from CAT (3.737 Mg, 3.8 m3)● D11T Bulldozer from CAT (104.236 Mg, 172.13 m3)

BACKUP: ANFO Analysis

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ANFO - Ammonium Nitrate/Fuel Oil- Most used commercially for excavation, demolition, etc.- Explosive efficiency range from 50 - 80 percent of their calculated

energies - Cost-efficient- Great for low density material explosions

BACKUP: Heavy Machinery via CAT

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Retrieved from: https://www.cat.com/en_US/products/new/attachments/buckets-excavator/heavy-duty/2000000591.html

Retrieved from: https://www.cat.com/en_US/products/new/equipment/dozers/large-dozers/18332635.html

BACKUP: Particle velocity mapping code

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BACKUP: Sources

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1. Abbaspour, H., Drebenstedt, C., Badroddin, M., and Maghaminik, A., “Optimized design of drilling and blasting operations in open pit mines under technical and economic uncertainties by system dynamic modelling,” International Journal of Mining Science and Technology, vol. 28, 2018, pp. 839–848.

2. Athanasios Goulas, Ross J. Friel, (2016) "3D printing with moondust", Rapid Prototyping Journal, Vol. 22 Issue: 6, pp.864-870, https://doi.org/10.1108/RPJ-02-2015-0022

3. Gibney, E., “How to build a Moon base,” Nature News Available: https://www.nature.com/articles/d41586-018-07107-4.

4. Johnson, S. M., Bourque, R. F., Day, W. C., Gates, R. H., Gilson, D. R., Fraser, R. L., ... & Remboldt, A. L. (1971). Explosive excavation technology (No. NCG-TR-21). ARMY ENGINEER NUCLEAR CRATERING GROUP LIVERMORE CA.

5. Rodriguez, J., “This Is Must-Have Heavy Equipment for Construction Projects,” The Balance Small Business Available: https://www.thebalancesmb.com/must-have-earth-moving-construction-heavy-equipment-844586.

6. Tate, K., “Home On the Moon: How to Build a Lunar Colony (Infographic),” Space.com Available: https://www.space.com/21588-how-moon-base-lunar-colony-works-infographic.html.

7. Wall, M., “Back to the Moon: How New Lunar Bases Will Work,” Space.com Available: https://www.space.com/10634-moon-base-lunar-outpost-technology.html.

Scout Rover: Ice Detection Equipment

Matthew EustaceCAD

Scout Rovers January 31, 2019

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Problem

The Problem: Determine the required equipment needed by the scout rovers to detect and uncover ice from the regolith.Requirements: ● Visual instruments to allow scientists to gather data● Navigation and obstacle avoidance systems● Subsurface and soil water detection● Regolith drilling and composition detector

Assumptions● No pre-existent rover production

Need to Determine● Component sizing and design● Equipment MPV● Drilling technique● Regolith composition detection● Materials used in design

Adapted From Pavi Ravi’s presentation

Solution

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● Camera’s○ Useful for navigation and capturing data

● Soil water detection○ Analyze Rock and soil sample using a laser,

camera, and spectrometer● Subsurface water detection● Regolith Drilling

Conclusion:● CAD design is similar to estimates and industry

values● Look into drilling depth and component materials

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Backup

Rover Sensors and EquipmentMatthew EustaceCAD Team

The sensors and equipment needed for the mission were analyzed in order to better understand the rover’s mass, power, and volume requirements. The process is to look at what requirements the rover must accomplish. Research was then performed to find relevant sensors and equipment that meet these requirements. Mass, power, and volume calculations were then performed in order to better define rover sizing.

Camera’sOne necessary requirement is to have camera’s to provide visual data to scientists during the Rover’s exploratory

mission.One possible source of equipment is similar to the Mastcam-Z. It is a camera mounted to the mast of the rover, that

allows colored panoramic imaging, 3D imaging, zoom capabilities, and high-definition video. The Mastcam-Z has a mass of about 0.004 Mg, requires about 17.4 watts, and has dimension of 0.11x0.12x0.26 m and therefore a volume of 0.0034 m2 .

NavigationAnother necessary requirement is navigational equipment. The Rover needs navigational capabilities so that it can

head towards predetermined locations looking for ice. The rover also needs to have the capability to adapt to the terrain and navigate obstacles not seen by ground control. While the GPS system would need to be different, the Mastcam could already provide the capabilities needed for getting data from the terrain ahead. An onboard computer could look at the specified direction of the rover while the images could be analyzed to determine the best route. 20

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Subsurface water detectionA necessary piece of equipment for this mission is a device that can look below the surface of the moon to look for

ice. Satellites can help target areas to deploy scout rovers, but once on the surface the rovers will need to define where to dig. It would take too long to successfully find subsurface ice if the scout rover was blindly digging holes.

One solution could be a sensor such as RIMFAX. This stands for Radar Imager for Mars’ Subsurface Experiment. This sensor uses radar waves to detect changes in the material underground. The instrument probes the ground beneath it and can find materials such as ice, rock, sand, and liquid water. RIMFAX has a mass of 0.003 Mg, power of 5-10 watts, and a volume of 0.0016 m2 or 0.196x0.120x0.066 m. It has the capability to see greater than 10 meters deep under the surface.

Soil Water DetectionAnother necessary piece of equipment is a tool that can analyze the components of soil. This sensor can then look at

soil samples around the rover and help analyze the best place to drill. Possible instruments could be the SuperCam or the ChemCam. Both instruments analyze rock and soil samples with

a laser, camera, and spectrometer. This tool can identify the chemical makeup of rocks, soil, and regolith from more than 7 meters away. They can look at the atomic and molecular makeup of the samples. The SuperCam has a mass of 0.0106 Mg, power of 17.9 watts, and volume of 0.017328 m2 or 0.38x0.24x0.19 m.

Regolith Drilling and Composition DetectorAnother requirement is a drill to dig into the soil once ice has been detected in the regolith. A device will also be

needed that can analyze the regolith sample and see what the composition is. The drill will have to reach a depth of at least 1 meter unless the science team requests a longer length. In order to reach 1 meter, the drill will have to be at least 1.5 meters in length based on various drills in production. So the volume would be about 0.09 m2, require about 100 watts of power, and have a mass of about 0.01 Mg. The Composition Detector is still being researched. 21

References

[1] Lecher, Colin, 2012. “The Ten Instruments that Mars Rover Curiosity willuse to Investigate the Red Planet,” Popular Science Available:https://www.popsci.com/technology/article/2012-07/tools-mars-rover-curosity-will-use-uncover-history-water-and-life

[2] “Rover,” NASA Available: https://mars.nasa.gov/mars2020/mission/rover/[3] Colaprete, Anthony, “Resource Prospector: Evaluating the ISRU Potential of the Lunar Poles,” NASA Available: https://www.hou.usra.edu/meetings/leag2016/presentations/Wednesday /Colaprete.pdf[4] “What are Science Instruments,” NASA Available: https://mars.nasa.gov/mars2020/mission/instruments/.

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Autonomous Lunar Landing Requirements

Kevin SheridanCommunication and Control

Landing/Deployment Vehicle(s) & Lunar Launch and Descent VehicleJanuary 31, 2019

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Problem: Autonomous Landing

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Waypoint Constraints

The Problem: Determine thruster requirements for landing, Determine compute/sensing package mass/power.

Requirements: • Landers must be able capable of correction maneuver.• Landing must be autonomous

Assumptions/Constraints: • Landing consists of 3 stages: zeroing of lateral

velocity, correction maneuver, and vertical landing. • Initial speed at 5 km is 1.69 km/s (from mission design)• Initial Mass: 5000 kg• 4 m X 2 m cylinder• Vacuum Isp of each thruster: 300 s• Maximum thrust: 71000 N

Solution: Thruster Requirements & Sensing/Compute Package

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Note: Thruster Characteristics, Trajectory parameters, Vehicle Mass/Inertia are configurable in Matlab code attached.

Note: Attitude and Position over time can be extracted from the code. This can be used for a visualization later.

Mass ~ 1 kg

Power 30 W

Sensing and Compute Package

Thruster Configuration: • 4 pairs of angled thrusters. (8 total)• Maximum Thrust Guess: 71 kN (per engine)• Vacuum Isp Guess: 300 s

1 m Camera

BACKUP: Minimum Time Trajectory Generation

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By choosing this thruster configuration, I was able to use the fact that forces can only be produced in the body frame z axis to compute the attitude of the spacecraft based on its acceleration.

Then, by using the higher order derivatives of the trajectory, the moments required can be computed at any point along the trajectory.

Finally, we can form a linear mapping between the thruster forces and moments and body frame Z force.

The trajectory generator implementation can be found here:https://github.com/AAE450/FutureMoon/tree/dev/LandingAnalysis

BACKUP: Wrapper function to trajectory generator

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Note: All code can be found here:https://github.com/AAE450/FutureMoon/tree/dev/LandingAnalysis

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Backup: Trajectory Generator

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Backup: Trajectory Generator

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Backup: Trajectory Generator

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Backup: Trajectory Generator

References

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Feetham, L., Aouf, N., Dubois-Matra, O., & Bourdarias, C. (2016, July). Image datasets for autonomous planetary landing algorithm development. In Mechanical and Aerospace Engineering (ICMAE), 2016 7th International Conference on(pp. 627-637). IEEE.

Sibley, G., Matthies, L., & Sukhatme, G. (2010). Sliding window filter with application to planetary landing. Journal of Field Robotics, 27(5), 587-608.

Cutler, M. J. (2012). Design and control of an autonomous variable-pitch quadrotor helicopter (Doctoral dissertation, Massachusetts Institute of Technology, Department of Aeronautics and Astronautics).

Data rate determination and link budget analysis

Minduli Wijayatunga Communications and Control

Safe Haven, Orbital Propellant Depot & Launch vehicles January 31, 2019

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Problem : Communication dropouts

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• Adjust orbits to improve Earth and North pole coverage

• Calculate downlink data rate. • Obtain an approximate link margin

for high definition video transmissions.

• Antenna resizing.

Previous weeks… • Communication satellite and Safe Haven orbits

were determined. • Orbits were optimised based on the % visibility

to the lunar south pole.

3 dropouts in visibility per day!

Possibly insufficient north pole coverage

Figure 1: Earth visibility and North pole visibility for a day (based on orbits suggested by Zach Trainer (COM) and Hamilton (MD)

Solution: Improved orbits, resultant data rate and link margin

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Spacecraft a(km) i(deg) e ⍵(deg) Ω(deg) 𝛉(deg)Safe Haven 4567.5 80 (90) 0.6 90 90(0) 270 Commsat 1 4900 70(90) 0.5 90 0(0) 0Commsat 2 4900 70(90) 0.5 90 90(0) 90 (120)Commsat 3 4900 70(90) 0.5 90 180(0) 180(240)

Table 1: Refined spacecraft orbits (Previous data from Zach Trainer (COMS) and Hamilton (MD)

Data rate = 14.26 Mbps(calculation on slide 6)

Figure 2: Earth visibility and North pole visibility of the redefined orbits for a day Table 2 : Antenna gain, diameter and beamwidth

Link margin = (0.69 ±0.21)dB(Link budget on slide 8)

X BAND Gain (dB) D(m) BW

Earth 50 5.02 0.49Commsat

(main) 25 0.28 8.74

Safe Haven/Commsat

(redundant) 20 0.16 15.55

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Backup : Total coverage of the refined orbits

Figure 4: Earth visibility, North pole visibility and south pole visibility of the redefined orbits for a day

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Backup : Method for calculating Earth Visibility

Figure 6: Steps for calculating Earth visibility

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Backup : Data rate determination

Smallest Earth visibility time = 2.81 h Total no of passes per day = 3 X 4 Definition of video = 1080 p

Required bitrate = 15 Mbps [1]

Data per day = 15 X 24 X 60 X 60 Mb

Total visibility = 12 X 2.81

Data rate = Data per day / Total visibility = 14.26 Mbps

Backup : Reasoning for frequency selection

BACKUP : Detailed link budget for Safe Haven

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Communications satellite > Parks Radio Telescope Parameter Value Value,dB Notes

RF frequency 8.5 X band communications Distance to ground station 3.84E+08

Data rate (Mbps) 14.2574 before error correction coding Phase modulation index

(rad/pk) 1.2[2] Transmit power 3.16227766 5 TWTA required for deep space

Transmit Passive loss 0.630957344 -2 Transmitting antenna gain 100 20 (from Joshua Yeggy’s presentation)

EIRP 199.5262315 23 Pt Lt Gt Path loss 1.87E+20 -2.03E+02 (4 pi d f / c)^2

Atmospheric loss 0.794328235 -1 Ground antenna gain 1258925.412 50 ParksTotal Recieced power 1.06517E-12 -119.7258183

Data to total power 0.868696858 -0.611317494 Data to noise power 1.104E-21 -209.5703093 Assume noise temperature of 80 k

Received Eb/ No 67.67196143 17.69277015 Pr + Dr -N - data rate(dB)Required Eb/No 2.511886432 4

Receiver system loss 0.630957344 -2 Link Margin 0.69277015

BACKUP : Detailed link budget for Communication Satellite

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Communications satellite > Parks Radio Telescope Parameter Value Value,dB Notes

RF frequency 8.5 X band communications Distance to ground station 3.84E+08

Data rate (Mbps) 14.2574 before error correction coding Phase modulation index

(rad/pk) 1.2[2] Transmit power 3.16227766 5 TWTA required for deep space

Transmit Passive loss 0.630957344 -2 Transmitting antenna gain 100 25 (from Joshua Yeggy’s presentation)

EIRP 199.5262315 23 Pt Lt Gt Path loss 1.87E+20 -2.03E+02 (4 pi d f / c)^2

Atmospheric loss 0.794328235 -1 Ground antenna gain 1258925.412 50 ParksTotal Recieced power 1.06517E-12 -119.7258183

Data to total power 0.868696858 -0.611317494 Data to noise power 1.104E-21 -209.5703093 Assume noise temperature of 80 k

Received Eb/ No 67.67196143 17.69277015 Pr + Dr -N - data rate(dB)Required Eb/No 2.511886432 4

Receiver system loss 0.630957344 -2 Link Margin 5.70

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Backup : Antenna diameter and beamwidth calculation

Antenna Diameter – antenna gain equation[4]

G = Gain in dBK = Efficiency (taken as 50%)D = Antenna diameter (m)λ = Wavelength (m)

Antenna Diameter – antenna beamwidth equation [4]

D = Antenna diameter (m)λ = Wavelength (m)

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Backup : Matlab code

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Backup : Matlab code

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Backup : Matlab code

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Backup : Matlab code

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Backup : Matlab code

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Backup : Matlab code

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Backup : Matlab code

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Backup : Matlab code

References

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[1] Understanding video bitrates, encoding.com

[2] Space mission analysis and design, James R. Wertz

[3] European Space Agency, https://www.esa.int

[4] Parabolic Reflector Antenna Gain,https://www.electronics-notes.com

Human Factors Supplies and Storage Requirements

Sandra BonillaHuman Factors

Resource StorageJanuary 31, 2019

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Problem

Problem: Personal supplies and basic items necessary for human survival.

Requirements:• The crew must have enough supplies for a 250-day mission. • Sufficient resources for 500 days is needed in case of

emergencies.

Need to Determine:• Necessary supplies and in what quantity.• Total mass and volume of supplies.• Whether a shower is feasible or if other options should be used.

Solution

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Mass (Mg) Volume (m3)

Food 2 12

Water (2 week supply) 5 5

Beds .36 46.1

*calculated by Miles Hokanson

Possible shower design:• 1.34 m3

• .2 Mg• 10.5 kW

Mass and Volume for Clothing and Toiletries

Shower Rinseless Shower

Total Mass (Mg) 0.894 1.39

Total volume (m3) 13.4 20.9

• Water will not run and drain continuously – less water will be needed.

• Water must be collected after each shower. • Total of 45 m3

resource storage needed.

• 2 week water supply located in the habitat.

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Moon Base Oxygen Generation

Scott SachaHuman FactorsGroup Leader

January 31, 2019

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Problem: Oxygen Supplies-Moon Base

• Requirements: 6.1732 kg Oxygen/day (O2)

• Discover• Pros: Awesome we just need to capture it!• Cons: No air/Not possible

• Bring• Pros: We can breathe!• Cons: Storage uses space and mass on

launches, limited time usage.

• Create• Pros: Long lasting survivable method.• Cons: Need materials and energy.

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Solution: Moon Base Electrolysis System

(255 Days)Power

(kW/day)Mass (kg)

Volume (m3)

Cost (USD) Notes

Water N/A 3570 3.57 0Assume water =

pure

KOH N/A 544.3108 0.454249 3072 N/A

Solar 3 UNKNOWN UNKNOWN 0Use disc. rover

power sys.

Battery 3 216 0.0943 14001.25V, 100 Amp

(parallel)

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[1] Human Research Program Education and Outreach, “Oxygen Generator System,” MATH AND SCIENCE @ WORK. [Online]. Available: https://www.nasa.gov/pdf/570242main_OxygenGen_CHEM_ED.pdf. [Accessed: 16-Jan-2019].

[2] CR Scientific LLC, “Electrolysis Experiments,” CR Scientific: Electrolysis Experiments - Introduction. [Online]. Available: http://www.crscientific.com/electrolysis.html. [Accessed: 16-Jan-2019].

[3] W. C. Adams, “Measurement of Breathing Rate and Volume in Routinely Performed Activites,” California Environmental Protection Agency, Aug-1994. [Online]. Available: https://www.arb.ca.gov/research/resnotes/notes/94-11.htm. [Accessed: 16-Jan-2019].

[4] martin.chaplin@btinternet.com, m. (2019). Electrolysis of water. [online] Www1.lsbu.ac.uk. Available at: http://www1.lsbu.ac.uk/water/electrolysis.html [Accessed 30 Jan. 2019].

Delta-V Budget for Ice Shipment on Lunar/Depot Round Trip

Cody HawkinsMission Design Group Leader

Martian Transport / Habitat / Landing & Deployment / Resource StorageJanuary 31, 2019

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Problem

- Requirements- Reusable vehicle must ferry ice

from surface to prop. depot- Ice delivery must be net positive

- Assumptions- Impulsive maneuvers- 2% reserve for ΔV- 6.68° inclination from Earth-Moon

plane to Moon equator

- Determine- ΔV requirements by site- Best site from ΔV perspective

Site Latitude (°)

Shackleton 89.9° S

Sverdrup 88.5° S

Rozhdestvenskiy 82.2° N

Solution

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- Found ΔV requirements varied little over landing sites

- All latitudes are extremely polar

- Recommend Shackleton Crater as mining/habitat site

- Based on ice field density and negligible ΔV difference

- ΔV listed here is the for cost for entire trajectory

- Assume the vehicle will be refueled at the OPD and from a small depot on the Moon

- Thus it only needs to carry ~½ of what is listed here

Site Latitude (°) ΔV (km/s)

Shackleton 89.9° S 4.978

Sverdrup 88.5° S 4.974

Rozhdestvenskiy 82.2° N 4.959

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Moon axis

Moon equator

Earth axis

Earth equator

Ecliptic plane

Earth Moon plane

6.68° obliquity

23.44° obliquity

Conclusion: need to reach 6.68° inclination relative to lunar equatorial plane for intercept with L1

- Change plane as far out as possible

L1

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Step Description Avg. ΔV (km/s)

1 Launch to circular LLO 1.63

2 Transfer from circular LLO to L1

0.64

3 Plane change at L1/apoapsis 0.095

4 Stop at L1 0.072

5 Transfer from L1 to LLO 0.072

6 Reverse plane change at L1/apoapsis

0.095

7 Circular LLO capture 0.64

8 Enter descent arc 0.021

9 Propulsive landing 1.69

TOTAL 4.955

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CODEclear all, close all, clc;

%% Important notes:% this code calculates the dV required to send ice from the lunar surface% to L1 and back. The ship may actually refuel at times, but this gives% a total accounting of the cost

% units of dV all in km/s% assume impulsive maneuvers everywhere% neglect L1 instability in this particular analysismu_Moon = 4902.801076; % gravitational parameter of the moon, km^3/s^2R_Moon = 1737.4; % lunar equatorial radius (km)w_Moon = 2*pi/(27.3217*86400); % angular velocity of the moon (rad/s)latitude_sites = [89.9, 88.5, 82.2]; % shackleton, sverdrup, rozhdestvenskiy latsr_park = 90 + R_Moon; % LLO parking orbit (km)inc_final = 6.68; % inclination of Moon/Earth orbit relative to lunar equator (deg)L1_rel_Moon = 58010; % distance from Moon to L1 (km)%% launch to LLOdV_extra = -R_Moon*w_Moon*cosd(latitude_sites); % dV gained from Moon rotationdV_LLO = sqrt(mu_Moon./r_park) + dV_extra; % dV required to reach LLO from launch sites%% hohmann tranfer from LLO to L1 r_p_hmnn = r_park; % periapsis of transfer at parking orbit (km)r_a_hmnn = L1_rel_Moon; % apoapsis of transfer at L1 (km)a_hmnn = 0.5*(r_p_hmnn + r_a_hmnn); % SMA of transfer (km)dV_hmnn_1 = sqrt(mu_Moon*(2/r_p_hmnn - 1/a_hmnn)) - dV_LLO; % dV to enter transfer%% first inclination change at L1 arrivalv_a_hmnn = sqrt(mu_Moon*(2/r_a_hmnn - 1/a_hmnn)); % velocity at apoapsis of transfer (km/s)dV_plane_1 = 2*v_a_hmnn*sind((latitude_sites-inc_final)./2); % dV to change plane at apoapsis%% stop at L1dV_L1_arrive = sqrt(mu_Moon*(2/r_a_hmnn - 1/a_hmnn)); % dV to kill remaining velocity at apoapsis/L1%% refuel

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%% depart L1dV_L1_depart = dV_L1_arrive; % same maneuver, going other way%% second inclination change at L1 departuredV_plane_2 = dV_plane_1; % same maneuver, going other way%% circularize at moondV_hmnn_2 = dV_hmnn_1; % same maneuver, going other way%% wait for alignment%% descent burnr_a_descent = r_park; % set apopasis at parking orbit altitude (km)r_p_descent = R_Moon; % set periapsis at surface (km)a_descent = 0.5*(r_a_descent + r_p_descent); % construct SMA (km)dV_beginDescent = sqrt(mu_Moon/r_a_descent) - sqrt(mu_Moon*(2/r_a_descent - 1/a_descent));%% land burnr_burn = 5 + R_Moon; % altitude at which landing burn is applied (km)dV_land = sqrt(mu_Moon*(2/r_burn - 1/a_descent)) + dV_extra; % dV to brake%% summary & plotdV_fromMoonToL1 = dV_LLO + dV_hmnn_1 + dV_plane_1 + dV_L1_arrive;dV_fromL1ToMoon = dV_L1_depart + dV_plane_2 + dV_hmnn_2 + dV_beginDescent + dV_land;dV_Total = dV_fromMoonToL1 + dV_fromL1ToMoon;

labels = categorical({'Shackleton', 'Sverdrup', 'Rozhdestvenskiy'});bar(labels, dV_Total);title('Hawkins: |\DeltaV| Required for Ice Shipment on Moon/OPD Round Trip')

xlabel('Launch/Base Site');ylabel('|\DeltaV| Required (km/s)');ylim([min(dV_Total)*.99 max(dV_Total)*1.01])

Cislunar Vehicle Departures and Arrivals

Matthew WardaMission Design

Cislunar & Landing Vehicle Trajectory DesignJanuary 31, 2019

Problem

● Requirements○ Vehicle must navigate between the Earth and Moon○ Must be able to access Lagrangian point L1 or L2 from low

lunar orbit (LLO)○ Minimal ΔV necessities for transfer orbits○ Preferably low time-of-flight (TOF) values

● Assumptions○ Impulsive maneuvers○ Centrobaric and symmetrical bodies○ Isolated system consisting of the Earth, Moon, and Sun

● Need to Determine○ Departure and arrival locations from the Earth and Moon○ ΔV and TOF comparisons between transfer options

Solution

Transfer Angle ΔV (km/s) ΔV + 2% margin of error (km/s) TOF (Earth days)

15° 13.320 13.586 4.885

30° 12.397 12.645 4.892

60° 10.445 10.654 4.917

90° 8.418 8.586 4.950

120° 6.426 6.555 4.983

150° 4.714 4.808 5.006

180° 3.930 4.009 5.015

Lagrangian Point ΔV (km/s) ΔV + 5% margin of error (km/s) TOF (Earth days)

L1 0.7147 0.7504 2.687

L2 0.7111 0.7466 3.136

Transfers from 300 km Earth orbit to 90 km Moon

orbit

Transfers from 90 km Moon orbit

to Lagrangian

points

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● References

○ Cornish, N. J. (2019, January 14). What is a Lagrange Point? | Solar System Exploration: NASA Science. Retrieved January 26, 2019, from https://solarsystem.nasa.gov/resources/754/what-is-a-lagrange-point/

○ Howell, E. (2017, August 21). Lagrange Points: Parking Places in Space. Retrieved January 27, 2019, from https://www.space.com/30302-lagrange-points.html

○ Williams, D. R. (2018, July 18). Planetary Fact Sheet. Retrieved January 27, 2019, from https://nssdc.gsfc.nasa.gov/planetary/factsheet/

Scout Rover Travel Power Required

Myles HomanPower and Thermal

Scout RoversJanuary 31, 2019

ProblemRequirements:

- The scout rover needs to be able to traverse inclines up to 30 degrees at a constant speed

- This constant speed needs to be 1.064 km/h or 0.2956 m/s- The scout rover will be around 168.2 kg or 0.168 Mg

Assumptions- The mass of the motor, wheels, and chassis is 50 kg- Aerodynamic resistance on the Moon is negligible- The speed of the scout rover is constant- The rover has 4 wheels with a coefficient of rolling resistance of 0.1

Need to Determine- The power required to overcome rolling resistance- The power required to travel at an incline- From these 2 powers, the total power required for travel

SolutionEquation: Ptotal = Pri+Prr+Pra -Pri is the power required to travel at an inclination, Prr is the power required to overcome rolling resistance, and Pra is the power to overcome aerodynamic resistance

-After calculating Ptotal, average electric engine efficiency and a safety factor of 30 % are taken into account

Chassis/motor mass

0.05 Mg (50 kg)

Chassis/motor power

0.1196 kW (119.6 W)

Chassis/motor volume

0.5 m3

Scout rover mass

0.168 Mg (168.2 kg)

Scout rover power

0.2115 kW (211.5 W)

Scout rover volume

1.131 m3

Backup

Backup

Backup

References

[1] “Electrical Motor Efficiency” Retrieved from https://www.engineeringtoolbox.com/electrical-motor-efficiency-d_655.html

[2] “Power to Move a Person” Retrieved from https://www.xootr.com/power-to-move-a-person.html

[3] Apostolopoulos, Dimitrios & Wagner, Michael & Leger, Chris & Jones, Jack. (2005). Experimental Characterization of a Robotic Inflatable Wheel.

Communication Satellites and Safe Haven

Alycia McEachenPower and Thermal Systems

January 17, 2019

85

Problem

86

• The communication satellites must remain in contact with Earth at all times.

• The systems in the safe haven must have ample power

• The temperature must be regulated effectively

Communication Satellite Requirements

Safe Haven Requirements

• The power source and how to store excess and reserves

Determine Determine• Insulation type and

amount• Power required to

regulate temperature

Solution

87

Communication Satellite

*Per Satellite

Safe Haven

Power* 3500 W

Area* 7.6 m2

Mass* 0.112 Mg

Lithium Ion

Volume 0.1035 m3

Mass 0.287 Mg

Habitat Volume 30m3

Insulation

Type RSI per meter (m*K/w)

Silica Aerogel 69

Polyurethane rigid panel 48

BACKUP

• Energy density of solar radiation at moon is similar to energy density on earth- 1368w/m2

88

Assumptions

BACKUP

Screen Shot 2019-01-31 at 12.08.33 AM

89

BACKUP

90

“Common Sizes of Solar Panels,” Brightstar SolarAvailable: http://brightstarsolar.net/common-sizes-of-solar-panels/.

“From the Lab to the Marketplace,” LBNL - KEVLAR--CLUE #1Available: https://www2.lbl.gov/Science-Articles/Archive/aerogel-insulation.html.

“Insulation,” Department of EnergyAvailable: https://www.energy.gov/energysaver/weatherize/insulation.

“Lithium Ion,” PanasonicAvailable: https://na.industrial.panasonic.com/sites/default/pidsa/files/ncr18650b.pdf.

Simone, “Is my roof suitable for solar panels (and what is the weight of a solar panel)?,” SunmetrixAvailable: https://sunmetrix.com/is-my-roof-suitable-for-solar-panels-and-what-is-the-weight-of-a-solar-panel/.

“Staying Cool on the ISS,” NASAAvailable: https://science.nasa.gov/science-news/science-at-nasa/2001/ast21mar_1.

References

BACKUP

91

Lunar Ascent/Descent Vehicle Feasibility

Ryan DelahuntyPropulsion

Lunar Ascent/Descent VehicleJanuary 31, 2019

92

Problem:

93

Mass [Mg] Volume [m^3] Power [kW]

Structure 10.00 3.690 -

Engine 2.500 100.0 -

Total Propellant 185.4 2265 -2,127,000

-Liquid Oxygen 43.11 20.28 -13,290

-Liquid Hydrogen 158.9 2245 -1,525,000

Payload 100 108.7 -460,100

Total Values 246.3 1846 -2,587,000

Refer to Cody Hawkins’ Presentation for Mission Details

Fuel Transport Mission – Sizing

Main Conclusion:

Total Mission delta-V: 4.955 km/s

Premise fails under ideal conditions -100 Mg payload per trip -Ideal Combustion -Light Structure -Steering loss not accounted for

Deliver 400 Mg of water, from the lunar surface to L1, over the course of a year, with a reusable bipropellant spacecraft.

The mission was tested via a 1DOF numerical method and verified by the Tsiolkovsky rocket equation. (See backup)

Premise

Not Feasible

Recommendation: Lunar Fuel Depot

94

Lunar Fuel Depot

Production Increase

-Adds design challenges to Mars Transport Vehicle-Simplifies overall engineering challenges

-Also requires a production output increase

Two Potential Solutions

-Increase would need to be by orders of magnitude-Inherently inefficient to deliver water to orbit with

traditional rocket propulsion

95

BACKUP

96

BACKUP

97

BACKUP

98

BACKUP

99

BACKUP

100

Backup

CisLunar Transport Vehicle

Stephane Peltereau-VilleneuvePropulsion

Cislunar Transport Vehicle | Orbital Fuel DepotJanuary 31, 2019

101

102

Problem

Requirement:

- Transfer from LEO (Low Earth Orbit) to LLO (Low Lunar Orbit)

- Delta V = 13.586 km/s (2% Error Margin)

Assumptions:

- 1 Mg Payload- 2 Mg Inert Mass

Need To Determine:

- Propellant Mass Required to carry out mission

- Volume of Fuel/Oxidizer Tanks- Tank Arrangement

Solution

103

Tank Specifications:

- LO2 (Liquid Oxygen) Volume: 0.6928 m3 - LH2 (Liquid Hydrogen) Volume: 1.860 m3

- Concentric Tank Design- Cryogenic Tanks

Propulsion System

- Refuelable LH2, LO2 transfer ports - Would require electric LH2 & LO2 electric

pumps- Four engines on gimbal- Pintle Injector design, fuel centered

BackUp

Backup

References

106

Sutton, G. P., and Biblarz, O., Rocket propulsion elements, Hoboken (N.Y.): Wiley., 2017.

Heister, S. D., Anderson, W. E., Pourpoint, T., and Cassady, J., Rocket Propulsion., 2019.

Safe Haven Specific Science Payload Specifications

Harry LlamasScience

Safe HavenJanuary 31, 2019

107

108

Problem

Chief Scientific Factors Affecting Safe Haven Specifications:• Radiation Shielding• Emergency Food and Water Supplies

Adequate radiation shielding will be required to ensure Astronauts do not receive a dangerous level of radiation whilst seeking shelter within the safe haven. Since this is an emergency shelter it will not necessarily have the same level of radiation shielding as within the habitat

Whilst they are within the safe haven they will have to have an adequate food supply to survive for short lengths of time and possible evacuation back to Earth

H

r

Solution

109

Mass (Mg) Power (Watts) Volume (Cubic Meters)

Total Food Consumption (Safety Factor of 2)

0.05712 N/A 0.336

Total Water Needs (Safety Factor of 2)

0.317975 N/A 0.317975

Radiation Shielding 10.26e-6 /cm² N/A 0.038 /cm²

Evacuation to Earth would take around 3 days¹ depending on Orbital path. Therefore, planning for periods of time of a week gives a conservative estimate for the amount of resources needed. Then for food and water estimates another safety factor of 2 is applied to ensure the astronauts are sustained whilst within the safe haven.

BACKUP

110

Total water needs of 6 man crew is around 6 gallons per day in an emergency situation² assuming astronauts will be drinking half a gallon per day and using another half gallon for personal hygiene. This would then mean for a weeks stay aboard the safe haven 42 Gallons would be required and then 84 Gallons giving a safety factor of 2 should more water be used or the stay be longer than expected. This equates to 317.975 litres.

Each Astronaut still needs to consume at least 2000 calories³ per day at a net mass of 3.6e-4 kg/calorie that would require at least 57.12 kg⁴ of food for a 2 week stay and at 2e-6 cubic metre per calorie that would give 0.336 cubic metres for a two week stay

The maximum allowable radiation exposure of a person as denoted by the united states is 50 mSv and this applies for lunar missions of up to 180 days, so for the 7 days they might be living within the safe haven it would be important to ensure that they receive less than 1.94 mSv⁵ of radiation Estimates for the worst case during a solar particle event put the total radiation received in the lunar environment at around 1 Sv per year⁶ so approximately 19.16495551 mSv for the week long duration. This would have to be mitigated by radiation shielding. Commonly aluminium is used as the casing of spacecrafts⁷, this can only absorb half the radiation that hits it, this would still mean the astronauts are receiving a very high dose during this stay.

111

Backup

Calculations made using thicknesses of Apollo command module⁹ walls and density of aluminium⁸

112

References

1. https://nssdc.gsfc.nasa.gov/nmc/spacecraft/display.action?id=1969-059A

2. https://www.beprepared.com/blog/18680/much-water-store/3. https://www.nasa.gov/vision/earth/everydaylife/jamestown-ne

eds-fs.html.4. https://airandspace.si.edu/exhibitions/apollo-to-the-moon/onli

ne/astronaut-life/food-in-space.cfm5. https://en.wikipedia.org/wiki/Spaceflight_radiation_carcinoge

nesis#Current_permissible_exposure_limits6. http://adsabs.harvard.edu/abs/2012P%26SS...74...78R7. https://www.nasa.gov/audience/foreducators/5-8/features/F_

Shielding_Space_Rays.html8. https://www.google.com/search?q=density+of+aluminium&o

q=density+of+aluminium&aqs=chrome..69i57j0l5.5335j1j7&sourceid=chrome&ie=UTF-8

9. https://en.wikipedia.org/wiki/Apollo_command_and_service_module#Command_module_(CM)

Preliminary Science Laboratory Sizing and Equipment

Theo SorgGroup Lead, Science

Habitat, Fuel-Depot, Landing Site PrepJanuary 31, 2019

113

Problem: Habitat Laboratory

114

Requirements: - The laboratory must not exceed 40 cubic meters of space- Equipment must not exceed 100 kg- Equipment plus tables/shelves must not exceed 1000 kg- Minimum of 8 cubic meters of free space to move about- Must be capable of analyzing geological samples

***This could change pending input from EAPS counterparts

Assumptions- All custom equipment/shelving made from Aluminum- Some equipment purchased as Commercial Off The Shelf (COTS) parts, modified for lunar use

Need to Determine- Contents of laboratory- Approximate sizing

Totals- Mass (excluding surrounding structure): 740 kg- Power required for equipment: 720 Watts- Volume: 30 cubic meters

Equipment- Scales (2)- Hammers (2)- Chisels (full set)- Polarizing light microscope- Scanning Electron Microscope

Future Iterations- Adjustments in equipment per science objectives- Layout adjustments per human factors input- More accurate volume estimation and CAD

Solution

115

4 meters

3 meters

2.5 meters

BACKUP: Equipment

116

Equipment Mass (Kg) Power (Watts)

Hammers 2 0

SEM [1] 50 700

Polarizing Light Microscope [2]

9 20

Chisels 5 0

Total 66 kg 720 W[1]https://www.nikonmetrology.com/en-gb/product/jcm-6000-plus-neoscope[2]https://microscopecentral.com/collections/petrology-geology-microscopes/products/amscope-binocular-polarizing-microscope-40x-640x

BACKUP: Tables and Shelves

117

Mass Estimate of Scout Rovers/Demo Systems

Jack GreenStructures

Scout Rovers/Demonstration SystemsJanuary 31, 2019

118

Problem: Determine Mass of Scout Rovers

119

Requirements: - Safely handle a capacity of between 0.079 Mg and 0.104 Mg (based on Science Team estimates)- Successfully deal with launch, landing, and deployment

- Landing may require additional structural support- Radiation Protection

Previous Calculation Estimates:- Science payload mass : ~0.103 Mg- Science payload volume : ~1.64 m3

- Numbers from Science Team

Need to Determine...- Architectural mass including

- Radiation Shield- Structure and support

- Volume of rover material

Solution: Material Type and Shape

120

Radiation Protection- Polyethylene Shield

- Density: 0.96 g/cm3

- Thickness: 5 cm- Has about half the mass of equivalent

aluminum shield- Excellent at radiation absorption due to its high

hydrogen contentShield Shape

- Dome optimal for space and mass minimization

Rover Structure- Additional Mass of ~0.03 Mg

Future Iterations- CAD Design- More detailed power analysis

- (with help from power team)

Results:

Backup - Material Mass Comparison

121

- For Polyethylene

- For Kevlar

Backup - References

122

[1] https://science.nasa.gov/science-news/science-at-nasa/2005/08sep_radioactivemoon [2] https://www.space.com/24731-mars-radiation-curiosity-rover.html [3] https://www.space.com/21353-space-radiation-mars-mission-threat.html [4] https://www.thomasnet.com/articles/custom-manufacturing-fabricating/radiation-shielding-materials [5] https://www.nasa.gov/vision/space/travelinginspace/radiation_shielding.html [6] https://www.nature.com/articles/s41598-017-01707-2.pdf [7] https://mars.nasa.gov/mer/technology/bb_power.html