Section III Presentations - engineering.purdue.edu · The Problem: Determine the Configuration of the Cycler Vehicle Requirements • The cycler must accommodate three docked taxi

Section III Presentations February 6, 2020

Eli SitchinFebruary 5, 2020

Discipline: CADVehicle and Systems Group: Cycler

2

The Problem: Determine the Configuration of the Cycler Vehicle

Requirements

• The cycler must accommodate three docked taxi vehicles without negatively affecting

stability.

• The cycler must rotate at a safe speed while providing > 0.38 g.

• All habitable portions of the cycler must be accessible from all other habitable portions.

Assumptions

• 65 m3 of habitable volume per person, with 70 total passengers

• 4 habitation modules

• 1 g at habitation module ceiling

Need to Determine

• Distance between the habitation module and axis of rotation

• Habitation module dimensions

• Taxi docking module placement

3

Preliminary Design

Considerations

• All structures currently modeled as

aluminum

• Mass to be estimated once

additional materials and

thicknesses are specified

Rotation Rate 0.1566 rad/s

Habitation Ceiling Radius 400 m

Floor Area1 2551 m2

Pressurized Volume1 17606 m3

Habitable Volume1 5451 m2

1 Habitation and Elevator-Habitation Interface only

Next Steps

• Determine solar panel requirements and

positioning

• Determine propulsion system positioning

• Create Habitation Module mockup

Habitation

Superstructure

SupportElevator

Elevator-

Habitation

Interface

4

Erick SmithFebruary 6th, 2020

CADLanding Pad(s)

Landing Pad Hub

Requirements:

• Must supply its own power

• Must be able to land three

taxis at a time

Solutions:

• Three landing pads

• Solar array ONLY for

landing pads and ATC

bunker

Landing Pad Itself

Requirements:

• Protects against the environment

• Move the rocket from vertical to

horizontal

• Load the taxi to the mass driver

system

• Refuel and resupply the taxi

Solutions:

• Underground gate

• Using a cradle help passengers

disembark and load taxi to mass

driver

Nicholas DeAngeloFebruary 6th, 2020

CommunicationsTaxi Vehicle (Phobos Tether and Mars Mass

Driver)

Problem

• Communication for Taxi Vehicle with Mars Mass Driver and Phobos

Tether - MPV• RF

• Ground satellites on Mars and Phobos

• Assumptions

• Allowed up to 1kW of power (allotted from Power and Thermal)• Laser Communication not effective through atmospheric barrier

• 9m satellites on Mars surface and Phobos surface• Same as Earth ground satellites

Taxi Vehicle Antenna

• Future Considerations:• Satellites in Mars orbit for continuous communication

• Account for further losses for more accurate/realistic data

• Minimize Mass, Power, Volume while staying within requrements

Parameter Value (unit)

Diameter 2 m

Mass [1] 20 kg

Power Needed 890.5 W

Frequency [3] 130 GHz

Adam WootenFebruary 6th, 2020

Communications Team LeadCycler – Laser/Optic

The Problem

Given Requirements

• Continuous Communication

• 1 Gbps HD Communication

Requirements Flowdown

• 4 dB Gain Margin

• Redundancy Cycler-Earth Comm

• Power < 25W

Goal

• Keep Same Antenna as SatellitesPhoto By: Adam Wooten

Max Cycler-Earth Distance: Jordan Mayer & Colin Miller

*Not To Scale*

Cycler Communication Optical Hardware

Mass (kg) 180

Power (W) 24

Volume (m3) 5.0

Antenna Diameter (m) 1.20

Gain Margin Min for

Relay (dB)

5.4903

Gain Margin Min for

Cycler-Earth (dB)

4.0860

Antenna Characteristics

Comm System Mass, Power, Volume

Photo By: Adam Wooten Photo By: Adam Wooten

*Not To Scale* *Not To Scale*

Cycler to Earth Max

DistanceCycler to Relay Max

Distance

Sidharth PrasadFebruary 6, 2020

Discipline: ControlsSystem: Taxi Vehicle

The Problem: Mars Entry

● Problem: Need to entry angle and necessary fuel to land

● Solution: Setup MATLAB file modelling dynamics

○ Too low angle exits atmosphere, too high can hurt

○ BC is an important parameter for determining good angles

● Going forward

○ Need more accurate taxi parameters, and can improve

atmospheric model and model accuracy

○ Need to calculate fuel needed to land and control for angle

○ Need to ensure that other metrics such as heat are not

surpassed

Graphs

BREAKResume at 2:10

Brady WalterFebruary 6, 2020

ControlsAnalysis of Communication Satellite Control

Problem

Requirements for Control

• Antenna pointing in correct direction

• Solar panels at optimal orientation

• Ion thruster oriented properly for course corrections

Considerations

• Continuous, accurate pointing required

• Gravitational and solar radiation apply torque

Solution

Reaction Wheels

• 3 wheels: pitch, yaw, roll

• Rockwell Collins RSI-12

Independent Solar Panel Control

• Panels rotate about one axis

• Maximize area in direct sunlight

Further Steps

• Develop feedback controller for solar panels to face sun

• Develop feedback controller for reaction wheels

Parameter (Per

Wheel)

Value

Mass 4.85 kg

Diameter .247 m

Height .085 m

Max Power

Consumption

191.25 W

Lifetime 15 yrs

Beverley K.W. YeoFebruary 6th, 2020

ControlsCycler – Stability Analysis

Problem: Spacecraft stability against perturbations

In space: no aerodynamic forces 🡪 stabilization via spin (gyroscopic effect)

What criterion to ensure cycler is stable?

Perturbations:

• Energy dissipation

• Radiation forces (solar radiation, reflected solar radiation, magnetic, thermal)

Assumptions:

• Small-spin approximations

• Spin stabilization only

• Cycler vehicle is rigid

Model & Analysis: Spin stabilization

Euler’s Equations [1]:

Small perturbations i.e. small 🡪 :

LHS 🡪 torques, calculated from perturbation forces

Take Laplace transform 🡪 roots of characteristic equations 🡪 stability

[1] Wertz, J. R. (1978). Spacecraft attitude determination and control. Dordrecht: Kluwer Academic Publishers.

Pertubation Force

(Near Earth, flyby)

Magnitude (N)

Solar Radiation 1.308E-2

Reflected Solar Radiation 2.701E-4

Planet Thermal Radiation 1.706E-4

Spacecraft Radiation P/c

Solar Wind 4.41E-6

Meteoroids Negligible

Emily SchottFebruary 6, 2020

Human Factors: TaxiEmergency Preparedness

The Problem(s):

Common issues in commercial travel:

• Medical Issues [1]• Fainting

• Respiratory Symptoms

• Nausea/Vomiting

• Heart Attacks and Seizures

• Misc• Delays

• Weather

Potential issues in space travel:

• Medical Issues• Space Sickness

• Heart Attacks

• Fainting

• Vehicle Malfunctions• Fires

• Leaks

• Launch/Landing Issues

Assume 4 day maximum travel time,

need supplies for trip to cycler and back

The Solutions

Item Mass (kg) Volume (m^3)

2 days of extra food 119.52 0.12

2 days of extra water 960 0.96

FAA FAK, EMK, UPK [2] 1.6 each 0.007 each

AED [3] 1.5 0.01

Fire Extinguisher [4] 3.5 0.01

Sickness bags [5] 2.4 0.01

Totals 1,090.12 1.131

Food and Water estimate: Human Factors, Kait Hauber

Taxi model: Structures, Nicki Liu

5m

15m

5m

Allow for

passenger

cabin to

separate and

perform

vertical landing

Alexey ZeninFebruary 5th, 2020

Discipline: Human FactorsVehicle/Systems: Cycler

Topic: Bioregenerative Life Support System (BLSS)

Problem: Continuous Food Cultivation on Board

Requirements:• The BLSS requires minimum to no inputs

• Sufficient production of O2

• Provision of all necessary nutrients

• Sufficient reduction of CO2

• Minimize Human and Cultivational waste

Objectives:

• Determine BLSS’s output/input requirements

• Estimate mass, power, and area requirements

• Technological and operational Overview

Assumptions:

O2

0.84 kg/day

CO2

1 kg/day

Crew

Member

Human

Waste

Estimations

Air

Bio-fermented

Waste(Fertilizer)BLSS

Oxygen

3.2308 Mg

Inedible

solid mass

Food

22.2 Mg

Michael PorterFebruary 06, 2020

Discipline: Mission DesignVehicle & System: Asteroid Tether Sling

The Problem: Selecting and Designing an Orbit

Problem 1:

Select feasible orbit

Problem 2

Find suitable asteroid

Problem 3

Determine attitude stability

GMAT analysis produced using some initial conditions by Turner [5]

The Solution

Solution 1: Orbital Parameters [3]

Type DRO

X-Amplitude 69,500 km

Resonance 1:1:2

Period 13.785 days (Luna)

Stability Over 100 years

Solution 3: DRO Natural Stability [4]

**Adapted from

figure by Guzzetti

& Howell [4]

BREAKResume at 2:44

Valentin RichardFebruary 6, 2020

Mission Design Asteroid Tether Sling

The problem: How much using an asteroid in the Earth/moon system could be beneficial ?

Assumptions:

➣ The asteroid will be located on a DRO orbit around the moon

➣ The travel times are not taken into account

➣ Cost & feasibility of getting the asteroid on the DRO is not taken into account

Comparative study between different possible paths:

1 ➣ LEO 🡪 DRO 🡪 Earth/Moon Escape

2 ➣ LEO 🡪 Moon 🡪 DRO 🡪 Earth/Moon Escape

3 ➣ LEO 🡪 Cycler

The solution: ∆V estimates Path 1: LEO 🡪 DRO 🡪 Earth/Moon Escape

Maneuver Required ΔV (km/s)

LEO to DRO[1]

1. Intermediate L1-Lyapunov

2. Lunar Far-Side Insertion

3. Close Lunar Flyby

(ED Tether)

[ 3.576 ; 4.334 ]

[ 3.456 ; 3.785 ]

[ 3.365 ; 3.393 ]

DRO to Earth/Moon Escape[2] 1.01 (Asteroid Tether)

Total: [ 4.375 ; 5.344 ]

Path 3: LEO 🡪 Cycler

Maneuver Required ΔV (km/s)

Total[2]: 4.3 (ED Tether)

Path 2: LEO 🡪 Moon 🡪 DRO 🡪 Earth/Moon Escape

LEO to LLO (Hohmann)[3] 3.959 (ED Tether)

LLO to DRO[2] 0.86 (Lunar Tether)

DRO to Earth/Moon Escape[2] 1.01 (Asteroid Tether)

Total: 5.829

To be determined: ∆V to get to the cycler from DRO/Moon

and the Tether Sling’s influence on the asteroid orbital stability

Pierre VEZIN02/01/2020

Mission Design – ElectrDyn. Tether(DeltaV calc, Dynamic Simulation, Orbit degradation)

How will the use of the ED Tether affect the orbit ?

The Problem: Momentum exchange

After SpaceCraft is released :

- S/C ends up with a Boost

- Tether gets slowed down

Requirements :

- Avoid ground at all cost

- Keep the Tether in a stable enough orbit

(high atmoshpere = drag)

Assumptions :

- separation triggered @ 1000 km

- S/C deltaV is prograde

- variables are Mass Ratio and S/C deltaV

Tether after

Release

Payload after

Release

EARTH

How will the use of the ED Tether affect the orbit ?

From a Circular LEO :

- Tethers w/ Mass ratio <= 15 will crash

- W/ tether length of up to 700km: not much margin !

From an elliptical LEO1 (energy stored through height):- Mass ratio can be reduced to 5 or 10 while remaining close to

LEO & w/ Perigee > 1000 km

[1] «Modeling and analyis of the Electrodynamic tether» J.Longuski, M.Mueterthies

M ratio : 10

15 20

M ratio : 10

15

2025

30

Melissa WhitcombFebruary 6, 2020

Mission Design

Orbital ∆v Calculations

∆v and Defining the OrbitsThe Main Objectives:

1. Define the orbits for all 4 cycler vehicles (and any other vehicles we create)

2. Find the ∆v for each path shift in the orbit

3. Find the total TOF for the human passengers on each orbit

The First Challenges:

1. Multiple versions of these calculations on the loose in 450

2. Asteroid characteristics?

3. How to use/open the cycler data?

4. Defining LEO, LLO, LMO, etc.

The Plan:

1. Get rough ∆v’s with Hohmann transfers, and refine via hyperbolic orbits and

Lambert arcs to get upper limits. (MATLAB → GMAT)

2. Condense (and verify) everyone’s ∆v calculations into one spreadsheet for all

teams’ use (Excel)

The Hohmann ∆v Values (minimum requirements)

Assumptions:

• Orbital velocities at LXO altitudes assumed to be

planar.

• LLO altitude taken from Apollo 11 and will need to be

updated with preferred stable orbit.

• Tether is assumed to be on the surface of the body

and not significantly affecting the length of body’s

radius.

• Asteroid is assumed to be of equal size and mass as

Phobos.

Coming Soon:

TOF’s (min & max) and ∆v (max)

Journey segment ∆v min

(km/s)

Earth surface to LEO 4.119

LEO to LLO/Asteroid 3.765

Moon surface to LLO 0.580

LLO to Cycler 1 3.640

Cycler to LMO 0.868

Mars surface to LMO 1.634

LMO to Phobos 1.911

Sources:

[1] NASA Planetary Fact Sheets

https://nssdc.gsfc.nasa.gov/planetary/planetfact.html

[2] ESA Mars Express https://sci.esa.int/web/mars-

express/-/31031-phobos

[3] Dr. Jekan Thanga, Purdue guest lecturer,

1/27/2020.

https://nssdc.gsfc.nasa.gov/planetary/planetfact.html

https://sci.esa.int/web/mars-express/-/31031-phobos

Peter Salek February 4, 2020

Discipline: Power and ThermalPowering the Mass Driver

1

Problem

Given Parameters

Moon:

Power required – 1.33 GW

Mars:

Power required – 2.66 GW

Mass Driver Efficiency – 90%

Requirements

• Investigate multiple methods of

power generation for the Mass

Driver

• Determine total mass of the system

Assumptions

• Multiple taxi vehicles will need to be

launched in succession

• Interval of 5 minutes between each

launch

2[1] AAE 450, Nicolas Martinez Cruces

[2] Davis, E., & Warp Drive Metrics LAS Vegas Nv. (2004). Advanced Propulsion Study.

[1]

[1]

[2]

Solution

Total Mass of

System

Moon Mars

Solar Panels 58,763 Mg 155,580 Mg

Nuclear Power 54,977 Mg 114,090 Mg

• Large Capacitor Banks to store energy for

each Launch

• Nuclear Reactors to Generate Power on

Both The Moon and Mars

• Power Generation of 10.2 MW on Mars

• Power Generation – 5 MW on The Moon

• Allows for launches regardless of sunlight

Future Power Generation Research

• Investigate using batteries to

store energy for launches

• Investigate Geothermal energy as

another power source on Mars

• Consider increasing time

between launches to decrease

mass

3

Josh SchmeidlerFebruary 6th, 2020

Power & ThermalThermal Analysis of Taxi Vehicle

Taxi Vehicle Heating on Mars Tether Sling

Problem

• Aerodynamic heating due to high speeds within atmosphere

Assumptions

• Tether situated on top of Olympus Mons

• Worst case scenario – 5 km/s flight speed

• Atmospheric entry model used

• Heating from convective and radiative heat transfer

Heating Rate

BREAKResume at 3:20

Joe TiberiFebruary 6th, 2020

Propulsion Team LeadED Tether

Escaping LEO with Electricity

• Investigating spinning tether in LEO

• Main Considerations investigated

• Astronaut accelerations

• Spin-Up/operating times

• Possible power consumptions

• Delta V required: 3.31 km/s

Physics Analysis• Current findings

• Large mass required to be in

LEO (3.4955e6 Kg)

• Very Long tether required

• Maximum Power Required

• ~170 MW

• Next steps

• Investigate traversing the

tether

• Adding in ED force

• Momentum bank size

requirements and lifespan

Shuting YangFebruary 06, 2020

PropulsionPhobos Tether SlingPropellant Analysis

Slide 1 of 3

Problem: Propellant Analysis

Requirement:

• No chemical propellants

Assumptions:

• A tapered tether system is used.

• The safety factor of the tether is 10.

• The maximum acceleration is 3g.

• The spin-up time at 3g is 24 hours.

Current Proposal:

• Use solar panels and batteries to convert and store solar energy from host star

to electric energy.

Analysis Procedures:

L: Tether Length

M: Tether Mass

A: Unit Solar Array Area

g: Gravitational

Acceleration on

Earth’s Surface

Slide 2 of 3

Solution:

23.7339

14.0531

13.3041

11.7481

10.2315

0 5 10 15 20 25

LOX/LH2

NTO/MMH

LOX/RP-1

NTO/Aerozine 50

HTPB

Pro

pella

nt

Type

Mass of Tether / Mass of Propellants

7.8468

4.6462

4.3986

3.8841

3.3827

0 2 4 6 8 10

LOX/LH2

NTO/MMH

LOX/RP-1

NTO/Aerozine 50

HTPB

A (m2)

Pro

pella

nt

Type

Solar Array Area

Slide 3 of 3

Natasha Yarlagadda February 6, 2020

Propulsion TeamTaxi - Mars Landing

56

Problem

Assumptions:- Constant gravity (overestimate)

- Cylindrical Aluminum Lithium alloy propellant tanks

- RP-1 (kerosene) is CH1.97 for chemical balances

- OMS/RCS system help taxi exit Mars orbit and align vertically

Mars Landing Site

vtank

mtank

moxidizer

mfuel

Types of propellant

Power generated

Requirements: Propulsion:

- Chemical propulsion system to decelerate taxi to safe landing speed

- Fthrust, required= 7686000 N[*]

Structures:

- Taxi needs to land vertically on Mars

- Propellant tanks fit within given volume, V = 211.17 m3

Human Factors:

- Reaction by-product provides reusable water

57*See backup slides for required thrust calculation

Solution: LH2 / LOX Propellant System

Propellants Selected LOX (liquid oxygen) LH2 (liquid hydrogen)

Propellant Masses[1] moxidizer= 64.80 Mg mfuel= 10.83 Mg

Tank Volumes[2] vox, tank= 59.63 m3 vfuel, tank= 160.6 m3

Power Generated[3] Fthrust, 1 engine = 2184076 N

Total Specs mtank+prop+engines= 181.9 Mg[4][5] vtank= 255.15 m3 Fthrust, 4 engines= 8736307 N

(Adapted from Space Shuttle RS-25)[6]

LOX LH2

References cited in backup slides58

Rachel RothFebruary 6, 2020

StructuresCommunication Satellites

Slide 1 of 3

Problem – Material Selection and Initial Sizing/Layout

System Requirements

Slide 2 of 3

Power and Thermal

• Solar energy collection and storage

• Thermal control system

Propulsion

• Ion propulsion system

Communications

• Earth relay satellites in GEO

• L4/L5 Lagrange points relay satellites

• Mars relay satellites

Tensile Strength

Thermal Performance

Corrosion Protection

Material Selection Criteria

Solution – Material Selection and Initial Sizing/Layout

Slide 3 of 3

[1] AAE 450 Communications Team

[2] AAE 450 Propulsion Team

[3] AAE 450 Power and Thermal Team

Communications Satellite Initial Sizing

solar panel

*GEO Satellite: 3 RF Antennas + OCS GEO to L4/L5

L4/L5 Satellite: OCS GEO to L4/L5 + OCS L4/L5 to Mars

Mars Satellite: 3 RF Antennas + OCS L4/L5 to Mars [1]

L4/L5

Top-down view Sun

solar panel

GEO/Mars

Top-down viewEarth/

Mars

RF antenna

gimbal

Al 7075-T7351

OCS

OCS

5.5 m

3 m

3 m

Eli Sitchin Backup SlidesFebruary 5, 2020

Discipline: CADVehicle and Systems Group: Cycler

62

Backup Slide: Superstructure Design

Considerations

• Taxis will dock parallel to the axis of rotation, to reduce their

effect on the moment of inertia.

• Taxi docking mechanism not yet designed

• Docking port radius preliminarily set to 2 m [1]

Next Steps

• Determine interior layout (controls, crew transport, etc.)

• Finalize docking port dimensions Docking Port

Radius 10 m

Height 45 m

Docking Port Location (from Bottom of Image) 32.5 m

63

Backup Slide: Habitation Module Design

Considerations

• Minimum safe ceiling radius of 100 m multiplied by SF = 4

• Interior height accommodates adult men of height +6σ [2]

• Interior width accommodates passenger quarters and one

hallway [3]

Next Steps

• Determine interior layout (quarters, galley, exercise area, etc.)

• Research life support requirements

Ceiling Radius 400 m

Interior Height 2.5 m

Interior Width 6 m

Interior-Exterior Distance 1.2 m Roof

Floor

Pressurized Volume 3220 m3

Habitable Volume 1138 m3

Floor Area 456.4 m2

Arc 11.14º

4 Modules Total

64

Backup Slide: Elevator-Habitation Interface Design

Considerations

• Arc designed to ensure sufficient room for elevator and

passage between adjacent habitation modules

• Volume requirements may change if life support/propulsion

equipment stored in the interface

Next Steps

• Determine interior layout

• Research life support/propulsion requirements

Pressurized Volume 2363 m3

Habitable Volume 900.9 m3

Floor Area 362.7 m2

Arc 1.776º

Ceiling Radius 400 m

Interior Height 2.5 m

Interior Length 38 m

Interior-Exterior Distance 1.2 m

2 Modules Total

65

Backup Slide: Elevator/Support Design

Considerations

• Elevator radii based on ISS truss [4] and ADA

elevator guidelines [5].

Next Steps

• Determine structural loads

Support Cross-Sectional Area 1 m2

Elevator Outer Radius 2.5 m

Elevator Inner Radius 2 m

Truss/Elevator Length 388.8 m

8 Supports Total

2 Elevators

Total

66

Backup Slide: MATLAB Code for Minimum Radius

67

Backup Slide: MATLAB Code for Minimum Radius

68

Backup Slide: MATLAB Code for Habitation Dimensions

69


70


71

References

[1] “Reference Guide to the International Space Station,” National Aeronautics

and Space Administration, Sep. 2015.

[2] Rosen, M., Appel, C., and Ritchie, H., “Human Height,” Our World in Data,

2019.

[3] “Shreve Room Layout,” Housing at Purdue University Available:

https://www.housing.purdue.edu/Housing/Residences/Shreve/layout.html.

[4] Petty, J. I., “STS-113 Payloads,” National Aeronautics and Space

Administration, Oct. 2002.

[5] “Part 36 - Nondiscrimination on the Basis of Disability by Public

Accommodations and in Commercial Facilities,” Americans With Disabilities

Act, Jul. 1991.

72

Erick Smith: Backup Slides February 6th, 2020

CADLanding Pad(s)

Landing Pad Hub

• Initial concept of the design

Landing Pad

• Initial Concept of the

Landing pad itself, including

unofficial materials used for

the design.

Nicholas DeAngeloFebruary 6th, 2020

Backup CommunicationsTaxi Vehicle (Phobos Tether and Mars Mass

Driver)

Backup – Link Budget MATLAB Script

• [2] [3] [4] Script for Link Budget to find Power required

Backup – Link Budget MATLAB Workspace

• Workspace for Link

Budget MATLAB Script

References

• [1] – AAE 450 Spring 2015 Project Aldrin-Purdue – Report pg. 50-51

• [2] – “Frequency Range Chart,” Amphenol RF Available:

https://www.amphenolrf.com/frequency-range-chart/.

• [3] – “Frequency Range Chart,” Amphenol RF Available:

https://www.amphenolrf.com/frequency-range-chart/.

• [4] – Free Space Path Loss Calculator Available:

https://www.pasternack.com/t-calculator-fspl.aspx.

https://www.amphenolrf.com/frequency-range-chart/

https://www.amphenolrf.com/frequency-range-chart/

https://www.pasternack.com/t-calculator-fspl.aspx

Sidharth PrasadFebruary 6, 2020

Backup ControlsTaxi Vehicle

Supplementary Slide: Matlab CodeVinf = 5*10^3; % m/s

mu_p = 4.282837*10^13; %m^3/s^2 grav param mars

re = 3389.5*10^3+135*10^3;

Vei = sqrt(Vinf^2+2*mu_p/re);

tspan = [0 1000];

y01 = [re,0,0,Vei,deg2rad(-25),0,0];

y02 = [re,0,0,Vei,deg2rad(-19),0,0];

y03 = [re,0,0,Vei,deg2rad(-14),0,0];

y04 = [re,0,0,Vei,deg2rad(-30),0,0];

[t1,y1] = ode45(@entrydynamics,tspan,y01);




Supplementary Slide: Matlab Code

figure()

plot(y4(:,4)/1000,(y4(:,1)- 3389.5*10^3)/1000)

hold on;

plot(y1(:,4)/1000,(y1(:,1)- 3389.5*10^3)/1000)

plot(y2(:,4)/1000,(y2(:,1)- 3389.5*10^3)/1000)

plot(y3(:,4)/1000,(y3(:,1)- 3389.5*10^3)/1000)

xlabel('Velocity (km/s)');

ylabel('Altitude (km)');

legend('Gamma = -30','Gamma = -25','Gamma = -19', 'Gamma = -14');

grid on;

title('Velocity vs. Altitude for Taxi (BC 150)');

% figure()

% plot(t,(y(:,1)- 3389.5*10^3)/1000);

% xlabel('Time (s)');

% ylabel('Altitude');

% grid on;

% title('Altitude vs. Time for Taxi (BC 300, Gamma -16)');

Supplementary Slide: Matlab Codefunction rho = density(r)

r_mars = 3389.5*10^3; % in m

alt = r-r_mars; %altitude

rho = 0.013*exp(-alt/11000); %(1/0.013)*e^alt; % density in kg/m^3

end

function dydt=entrydynamics(t,y)

%y(r,theta,phi,V,gamma,psi,s)

%dydt (rdot,thetadot,psidot,Vdot,gammadot,psidotsd0t)

sigma = 0;% bank angle

q = 1/2*density(y(1))*y(4)^2;%dynamic pressure

m = 22633.00; % vehicle mass

CD = 1.7;%coeff drag

A = 50;%Ref Area approx of space shuttle lander

LD = 0; % 0.8 % Lift/Drag approximated from space shuttle

Beta = 200; %m / (CD*A)%Ballistic Coefficient

mu_p = 4.282837*10^13; %m^3/s^2 grav param mars

gr = -mu_p/y(1)^2;%radial grav component

Supplementary Slide: Matlab Code

dydt(1)= y(4)*sin(y(5));

dydt(2)= y(4)*cos(y(5))*cos(y(6))/(y(1)*cos(y(3)));

dydt(3)= y(4)*cos(y(5))*cos(y(6))/(y(1));

dydt(4)= -q/Beta + gr*sin(y(5));

dydt(5)= (q*LD)/(y(4)*Beta)*cos(sigma) + 1/y(4)*(gr*cos(5)) +

(y(4)*cos(y(5))/y(1));

dydt(6)= (q*LD)/(y(4)*Beta)*sin(sigma)/cos(y(4)) -

y(4)/y(1)*cos(y(5))*cos(y(6))*tan(y(3));

dydt(7)= y(4)*cos(y(5));

dydt = dydt';

end

Brady WalterFebruary 6, 2020

ControlsAnalysis of Communication Satellite Control:

Backup Slides

Sources

Barret, C. “Spacecraft Flight Control System Design Selection Process for a

Geostationary Communication Satellite”, Marshall Space Flight Center, Alabama:

Sept. 1992. Accessed via

https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930001814.pdf

“RSI 12 Momentum and Reaction Wheels”, Collins Aerospace, 2020. Accessed

5 Feb 2020. Accessed via https://www.rockwellcollins.com/Products-and-

Services/Defense/Platforms/Space/RSI-12-Momentum-and-Reaction-

Wheels.aspx

https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930001814.pdf

https://www.rockwellcollins.com/Products-and-Services/Defense/Platforms/Space/RSI-12-Momentum-and-Reaction-Wheels.aspx

Beverley K.W. YeoFebruary 6th, 2020

Backup - ControlsCycler – Stability Analysis

Appendix: Perturbation Force Calculation

Perturbation Force Values based on Near-Earth flyby (N) [2] Magnitude (N)

Solar Radiation 9.2E-5 1.308E-2

Reflected Solar Radiation 1.9E-5 2.701E-4

Planet Thermal Radiation 1.2E-5 1.706E-4

Spacecraft Radiation 2.7E-7 P/c

Solar Wind 3.1E-8 4.41E-6

Meteoroids 4.9E-10 Negligible

• Scaled according to area ratio , cycler area based on L =

120m, W = 15.65m (from CAD)

[2] Longuski, J. M., Todd, R. E., & Konig, W. W. (1992). Survey of nongravitational forces and space environmental

torques - Applied to the Galileo. Journal of Guidance, Control, and Dynamics, 15(3), 545–553. doi: 10.2514/3.20874

Alexey ZeninFebruary 5th, 2020

Backup Slides

Backup Slides

Backup Slides

Backup Slides

Backup Slides

Backup Slides

Backup Slides

Backup Slides

References:

Alamaro M. 2007. Large-scale greenhouses attached to power plants for the productive use of

waste heat and CO2 emissions. Copyright#Moshe Alamaro 2007

alamaro.home.comcast.net/alamaro/Greenhouse Concept.pdf

Barlow P. 1996. An introduction to gravity perception in plants and fungia multiplicity of mechanisms.

Adv Space Res. 17:6972.

Bartsev SI, Gitelson JI, Lisovsky GM, Mezhevikin VV, Okhonin VA. 1996. Perspectives of

different type biological life support systems (BLSS) usage in space missions. Acta

Astronautica. 39(8):617622.

Tako Y, Masuda T, Tani T, Shinohara M, Abe K, Arai R, Suzuki M, Tsuga S, Komotsubara O, Nozoe S,

Aibe Y. 2008. On a Japanese bioregenerative life support programThe Mini-earth

projectoverview of the Closed Ecology Experiment Facilities (CEEF), some results from

experiments including closed human habitation and circulation of air and water, and future

plan of experiments using the CEEF. Presented at Workshop on Bioregenerative

Environmental Control State-of-the-art and Trends; 1820 December; Torino, Italy.

Michael Porter – BackupFebruary 06, 2020

Discipline: Mission DesignVehicle & System: Asteroid Tether Sling

DRO vs. Halo Analysis

***All analysis comes from notes taken speaking to Brian McCarthy, Masters Thesis on Halo orbits

DRO Detailed Analysis• DRO’s in the 60,000-70,000 km amplitude have

shown immense stability during 30 year sims. [1]

• A DRO of 61,500 km has shown stability for over 500

years (no station keeping required to maintain orbit)

while accounting for perturbing forces via solar,

Earth, Venus, Jupiter, solar radiation pressure and

asymmetric lunar gravity field [1]. Not chosen due to

unstable natural attitude. [4]

• Analysis run via GMAT has confirmed the current

DRO stability for a 100 year span accounting for

asymmetric lunar gravity field, solar, Saturn, Jupiter,

Earth gravity fields, and SRP.

• 1:1:2 period means the Moon passes the Earth once

every time the spacecraft passes the Earth once, but

the spacecraft will cross between the Moon and

Earth twice in this time .

Solution 1: Orbit Parameters [3]

Type DRO

X-Amplitude 69,500 km

Resonance 1:1:2

Period 13.785 days (Luna)

Stability Over 100 years

433 Eros (A898 PA) Orbital Elements

***All Orbital Elements are taken from JPL [2]

and recreated in GMAT

References[1] Bezrouk, C. J., Parker, J., “Long Duration Stability of Distant Retrograde Orbits,” AIAA, published

online 01 Aug. 2014

https://doi.org/10.2514/6.2014-4424

[2] Chodas, P., “NEO Earth Close Approaches,” JPL Close Approach Database, retrieved 03 Feb. 2020

https://cneos.jpl.nasa.gov/ca/

[3] Conte, D., Di Carlo, M., Ho, K., Spencer, D., Vasile, M., “Earth-Mars transfers through Moon Distant

Retrograde Orbits,” Acta Astronautica, Vol. 143, Feb. 2018, pp. 372-379

https://doi.org/10.1016/j.actaastro.2017.12.007

[4] Guzzetti, D., Howell, K., “Natural periodic orbit-attitude behaviors for rigid bodies in three-body

periodic orbits,” Acta Astronautica, Vol. 130, Jan.-Feb. 2017, pp. 97-130

https://doi.org/10.1016/j.actaastro.2016.06.025

[5] Turner, G., “Results of Long-Duration Simulation of Distant Retrograde Orbits,” Journal of

Aerospace, published online 08 Nov. 2016

https://pdfs.semanticscholar.org/e6ae/ca6ab0ddccb718c3d6ba24b57fcf5f4a9a57.pdf

GMAT Code DRO modeling

%General Mission Analysis Tool(GMAT) Script

%Created: 2020-01-29 21:48:50

%----------------------------------------

%---------- Spacecraft

%----------------------------------------

Create Spacecraft DefaultSC;

GMAT DefaultSC.DateFormat = UTCGregorian;

GMAT DefaultSC.Epoch = '07 Jan 2016 11:59:28.000';

GMAT DefaultSC.CoordinateSystem = moon;

GMAT DefaultSC.DisplayStateType = Cartesian;

GMAT DefaultSC.X = -61499.99999999993;

GMAT DefaultSC.Y = -4.092726157978177e-012;

GMAT DefaultSC.Z = 7.275957614183426e-012;

GMAT DefaultSC.VX = -1.908195823574488e-016;

GMAT DefaultSC.VY = 0.4999999999999994;

GMAT DefaultSC.VZ = -1.734723475976807e-017;

GMAT DefaultSC.DryMass = 687000000000000;

GMAT DefaultSC.Cd = 2.2;

GMAT DefaultSC.Cr = 1.8;

GMAT DefaultSC.DragArea = 15;

GMAT DefaultSC.SRPArea = 1;

GMAT DefaultSC.NAIFId = -10001001;

GMAT DefaultSC.NAIFIdReferenceFrame = -9001001;

GMAT DefaultSC.OrbitColor = Red;

GMAT DefaultSC.TargetColor = Teal;

GMAT DefaultSC.OrbitErrorCovariance = [ 1e+070 0 0 0 0 0 ; 0

1e+070 0 0 0 0 ; 0 0 1e+070 0 0 0 ; 0 0 0 1e+070 0 0 ; 0 0 0 0

1e+070 0 ; 0 0 0 0 0 1e+070 ];

GMAT DefaultSC.CdSigma = 1e+070;

GMAT DefaultSC.CrSigma = 1e+070;

GMAT DefaultSC.Id = 'SatId';

GMAT DefaultSC.Attitude = CoordinateSystemFixed;

GMAT DefaultSC.SPADSRPScaleFactor = 1;

GMAT DefaultSC.ModelFile = 'aura.3ds';

GMAT DefaultSC.ModelOffsetX = 0;


GMAT DefaultSC.ModelOffsetY = 0;

GMAT DefaultSC.ModelOffsetZ = 0;

GMAT DefaultSC.ModelRotationX = 0;

GMAT DefaultSC.ModelRotationY = 0;

GMAT DefaultSC.ModelRotationZ = 0;

GMAT DefaultSC.ModelScale = 1;

GMAT DefaultSC.AttitudeDisplayStateType = 'Quaternion';

GMAT DefaultSC.AttitudeRateDisplayStateType =

'AngularVelocity';

GMAT DefaultSC.AttitudeCoordinateSystem =

EarthMJ2000Eq;

GMAT DefaultSC.EulerAngleSequence = '321’;

%----------------------------------------

%---------- ForceModels

%----------------------------------------

Create ForceModel Propagator1_ForceModel;

GMAT Propagator1_ForceModel.CentralBody = Earth;

GMAT Propagator1_ForceModel.PrimaryBodies = {Earth};

GMAT Propagator1_ForceModel.PointMasses = {Luna, Sun};

GMAT Propagator1_ForceModel.Drag = None;

GMAT Propagator1_ForceModel.SRP = Off;

GMAT Propagator1_ForceModel.RelativisticCorrection = Off;

GMAT Propagator1_ForceModel.ErrorControl = RSSStep;

GMAT Propagator1_ForceModel.GravityField.Earth.Degree =

4;

GMAT Propagator1_ForceModel.GravityField.Earth.Order = 4;

GMAT Propagator1_ForceModel.GravityField.Earth.StmLimit =

100;

GMAT

Propagator1_ForceModel.GravityField.Earth.PotentialFile =

'JGM2.cof';

GMAT Propagator1_ForceModel.GravityField.Earth.TideModel

= 'None';

%----------------------------------------

%---------- ForceModels

%----------------------------------------

Create ForceModel DefaultProp_ForceModel;


GMAT DefaultProp_ForceModel.CentralBody = Luna;

GMAT DefaultProp_ForceModel.PrimaryBodies = {Luna};

GMAT DefaultProp_ForceModel.PointMasses = {Earth, Sun};

GMAT DefaultProp_ForceModel.Drag = None;

GMAT DefaultProp_ForceModel.SRP = Off;

GMAT DefaultProp_ForceModel.RelativisticCorrection = Off;

GMAT DefaultProp_ForceModel.ErrorControl = RSSStep;

GMAT DefaultProp_ForceModel.GravityField.Luna.Degree = 4;

GMAT DefaultProp_ForceModel.GravityField.Luna.Order = 4;

GMAT DefaultProp_ForceModel.GravityField.Luna.StmLimit =

100;

GMAT

DefaultProp_ForceModel.GravityField.Luna.PotentialFile =

'LP165P.cof';

GMAT DefaultProp_ForceModel.GravityField.Luna.TideModel

= 'None';

%----------------------------------------

%---------- Propagators

%----------------------------------------

Create Propagator DefaultProp;

GMAT DefaultProp.FM = DefaultProp_ForceModel;

GMAT DefaultProp.Type = RungeKutta89;

GMAT DefaultProp.InitialStepSize = 60;

GMAT DefaultProp.Accuracy = 9.999999999999999e-012;

GMAT DefaultProp.MinStep = 0.001;

GMAT DefaultProp.MaxStep = 2700;

GMAT DefaultProp.MaxStepAttempts = 50;

GMAT DefaultProp.StopIfAccuracyIsViolated = true;

%----------------------------------------

%---------- Coordinate Systems

%----------------------------------------

Create CoordinateSystem moon;

GMAT moon.Origin = Luna;


GMAT moon.Axes = ObjectReferenced;

GMAT moon.XAxis = R;

GMAT moon.ZAxis = N;

GMAT moon.Primary = Earth;

GMAT moon.Secondary = Luna;

Create CoordinateSystem moon_fixed;

GMAT moon_fixed.Origin = Luna;

GMAT moon_fixed.Axes = BodyFixed;

Create CoordinateSystem sun_fix;

GMAT sun_fix.Origin = Sun;

GMAT sun_fix.Axes = BodyFixed;

Create CoordinateSystem asteroid_frame;

GMAT asteroid_frame.Origin = DefaultSC;

GMAT asteroid_frame.Axes = BodyFixed;

%----------------------------------------

%---------- Subscribers

%----------------------------------------

Create OrbitView DefaultOrbitView;

GMAT DefaultOrbitView.SolverIterations = Current;

GMAT DefaultOrbitView.UpperLeft = [ 0.004705882352941176

0.002534854245880862 ];

GMAT DefaultOrbitView.Size = [ 0.4994117647058823

0.4486692015209126 ];

GMAT DefaultOrbitView.RelativeZOrder = 898;

GMAT DefaultOrbitView.Maximized = false;

GMAT DefaultOrbitView.Add = {DefaultSC, Earth, Luna};

GMAT DefaultOrbitView.CoordinateSystem = moon;

GMAT DefaultOrbitView.DrawObject = [ true true true ];

GMAT DefaultOrbitView.DataCollectFrequency = 1;

GMAT DefaultOrbitView.UpdatePlotFrequency = 50;

GMAT DefaultOrbitView.NumPointsToRedraw = 0;

GMAT DefaultOrbitView.ShowPlot = true;

GMAT DefaultOrbitView.MaxPlotPoints = 20000;


GMAT DefaultOrbitView.ShowLabels = true;

GMAT DefaultOrbitView.ViewPointReference = Luna;

GMAT DefaultOrbitView.ViewPointVector = [ 0 0 950000 ];

GMAT DefaultOrbitView.ViewDirection = Luna;

GMAT DefaultOrbitView.ViewScaleFactor = 1;

GMAT DefaultOrbitView.ViewUpCoordinateSystem =

EarthMJ2000Eq;

GMAT DefaultOrbitView.ViewUpAxis = Z;

GMAT DefaultOrbitView.EclipticPlane = Off;

GMAT DefaultOrbitView.XYPlane = Off;

GMAT DefaultOrbitView.WireFrame = Off;

GMAT DefaultOrbitView.Axes = On;

GMAT DefaultOrbitView.Grid = Off;

GMAT DefaultOrbitView.SunLine = Off;

GMAT DefaultOrbitView.UseInitialView = On;

GMAT DefaultOrbitView.StarCount = 7000;

GMAT DefaultOrbitView.EnableStars = Off;

GMAT DefaultOrbitView.EnableConstellations = On;

%----------------------------------------


%----------------------------------------

Create OrbitView DefaultOrbitView2;

GMAT DefaultOrbitView2.SolverIterations = Current;

GMAT DefaultOrbitView2.UpperLeft = [ 0.01588235294117647

0.6514575411913816 ];

GMAT DefaultOrbitView2.Size = [ 0.4947058823529412

0.4385297845373891 ];

GMAT DefaultOrbitView2.RelativeZOrder = 903;

GMAT DefaultOrbitView2.Maximized = false;

GMAT DefaultOrbitView2.Add = {DefaultSC, Earth, Luna};

GMAT DefaultOrbitView2.CoordinateSystem = EarthFixed;

GMAT DefaultOrbitView2.DrawObject = [ true true true ];

GMAT DefaultOrbitView2.DataCollectFrequency = 1;

GMAT DefaultOrbitView2.UpdatePlotFrequency = 50;

GMAT DefaultOrbitView2.NumPointsToRedraw = 0;

GMAT DefaultOrbitView2.ShowPlot = true;


GMAT DefaultOrbitView2.MaxPlotPoints = 20000;

GMAT DefaultOrbitView2.ShowLabels = true;

GMAT DefaultOrbitView2.ViewPointReference = Earth;

GMAT DefaultOrbitView2.ViewPointVector = [ 0 0 550000 ];

GMAT DefaultOrbitView2.ViewDirection = Earth;

GMAT DefaultOrbitView2.ViewScaleFactor = 1;

GMAT DefaultOrbitView2.ViewUpCoordinateSystem =

EarthMJ2000Eq;

GMAT DefaultOrbitView2.ViewUpAxis = Z;

GMAT DefaultOrbitView2.EclipticPlane = Off;

GMAT DefaultOrbitView2.XYPlane = Off;

GMAT DefaultOrbitView2.WireFrame = Off;

GMAT DefaultOrbitView2.Axes = On;

GMAT DefaultOrbitView2.Grid = Off;

GMAT DefaultOrbitView2.SunLine = Off;

GMAT DefaultOrbitView2.UseInitialView = On;

GMAT DefaultOrbitView2.StarCount = 7000;

GMAT DefaultOrbitView2.EnableStars = Off;

GMAT DefaultOrbitView2.EnableConstellations = On;

%----------------------------------------


%----------------------------------------




0.2002534854245881 ];


0.4372623574144487 ];




GMAT DefaultOrbitView3.CoordinateSystem =

EarthMJ2000Eq;









GMAT DefaultOrbitView3.ViewPointReference = Earth;


GMAT DefaultOrbitView3.ViewDirection = Earth;



EarthMJ2000Eq;












%----------------------------------------


%----------------------------------------




0.006337135614702154 ];


0.5247148288973385 ];




GMAT DefaultOrbitView4.CoordinateSystem = sun_fix;


GMAT DefaultOrbitView4.DrawObject = [ false true false ];







GMAT DefaultOrbitView4.ViewPointReference = Sun;


GMAT DefaultOrbitView4.ViewDirection = Sun;


GMAT DefaultOrbitView4.ViewUpCoordinateSystem = sun_fix;





GMAT DefaultOrbitView4.Axes = Off;







%----------------------------------------


%----------------------------------------




0.5627376425855514 ];


0.4359949302915083 ];


GMAT DefaultOrbitView5.Maximized = true;



GMAT DefaultOrbitView5.CoordinateSystem = asteroid_frame;








GMAT DefaultOrbitView5.ViewPointReference = Luna;

GMAT DefaultOrbitView5.ViewPointVector = DefaultSC;

GMAT DefaultOrbitView5.ViewDirection = Luna;



EarthMJ2000Eq;












%----------------------------------------

%---------- Mission Sequence

%----------------------------------------

BeginMissionSequence;

Propagate DefaultProp(DefaultSC) {DefaultSC.ElapsedDays =

100};

Valentin RichardFebruary 6, 2020

Mission Design Asteroid Tether Sling

Backup slide: References

[1] Lucia Capdevilla, David Guzzetti and Kathleen C. Howell, “Various transfer options from Earth to

Distant Retrograde Orbits in the vicinity of the Moon”, AAS 14-167

[2] David Conte, Marilena Di Carlo, Koki Ho, David B. Spencer and Massimiliano Vasile, “Earth-Mars

transfers through Moon Distant Retrograde Orbits”, Acta Astronautica, Volume 143, February 2018,

Pages 372-379

[3] Henry J. Pernicka, Steven M. Marsh, Theodore H. Sweetser and Deborah I. Scarberry, “A Search for

Low AV Earth-to-Moon Trajectories”, AIAA/AASA Astrodynamics Conference, Scottsdale (Arizona),

August 1994

Backup slide: Paths Additional info

LEO To DRO Time Of Flight (TOF) ESTIMATES[2]:

Path TOF (in days)

Lunar Near-Side Insertion (L1

Lyapunov)

LEO Departure maneuver: [ 4.15 ; 5.67 ]

Time spent in L1 Lyapunov: [ 5.84 ; 9.40 ]

DRO Insertion maneuver : Impulsive

Total: [ 9.99 ; 15.07 ]

Lunar Far-Side Insertion LEO Departure maneuver: [ 5.19 ; 10.90 ]

DRO Insertion maneuver : Impulsive

Total : [ 5.19 ; 10.90 ]

Close Lunar Flyby Insertion Only data found: Lunar DRO at C (Jacobi Constant) = 2.91

Total : 21.394

Backup slide: Asteroid Tether Dimensions estimateThe datas have been computed via my

own MATLAB Code, which can be found

on the next page.

A factor of safety = 10 is used for the

calculations, which explains why the

values are that high.

The ∆V used is 2.8 km/s and is a rough

first estimate of the ∆V needed for going

from DRO to Cycler. This value will be

updated as soon as we find a more

precise value.

However, computing these estimate

values can still be interesting to estimate

the order of magnitude of the different

parameters.

Asteroid Tether Sling dimensions

Maximum radial acceleration 3g

Material info (given by

structures team)

Dyneema

UTS = 3.325 Gpa

Density = 970 kg/m3

Tether length 319.8 km

Tether area: - Hub

- Tip

616.4 m2

6.7 * 10-3 m2

Tether mass 5.011 * 1010 kg

Maximum centripetal force 2214 kN

Backup slide: MATLAB Code

Melissa WhitcombFebruary 6, 2020

Mission Design

Orbital ∆v Calculations BACKUP SLIDES

Mass (kg)

Volumetric Mean

Radius (m)

Avg Distance

from surface of

parent body (m)

Altitude from

surface (m) at

LXO

Orbital Velocity

at LO altitude

(m/s)

Semimajor axis

a of orbit

between

Earth 5.97E+24 6371000 N/A 1000000 7,352.2

Luna 7.35E+22 1737400 382000000 110000 1,629.0

Mars 6.42E+23 3389500 N/A 400000 3,361.0

Phobos 1.06E+16 11033 5989000 0 8.0

Asteroid 1.06E+16 11033 LLO? 0 8.0

*assume m_A ~

m_P

Grav Constant, G

= 6.67E-11

LEO to LLO (the

asteroid)

Earth surface

to LEO

LLO to/from

lunar surface

LLO (asteroid)

to cycler Cycler to LMO

LMO to/from

Mars surface LMO to Phobos

m Rp = 7371000 Over the life of Over the life of 3789500

m Ra = 388481000 cycler 1, the cycler 1, the 9389533

m/s vt1 = 10300

DeltaV

expended

DeltaV

expended 4755

m/s Delta_V1 = 2948 goes from goes from 1394

m/s vt2 = 195 2.591 to 4.371, 0.09 to 1.639, 1620

m/s Vaph = 1013 with an average with an average 2136

m/s Delta_V2 = 817 of of 516

m/s Total DeltaV= 3765 4119 580 1634 1911

depends greatly

on the rocket's

acceleration

depends greatly

on the rocket's

acceleration

depends greatly

on the rocket's

acceleration

km/s Total DeltaV= 3.765 4.119 0.580 3.64041 0.868 1.634 1.911

Further calculations for ∆v

∆v and Defining the OrbitsThe Main Objectives:

1. Define the orbits for all 4 cycler vehicles (and any other vehicles we create)

2. Find the ∆v for each path shift in the orbit

3. Find the total TOF for the human passengers on each orbit

The First Challenges:

1. Multiple versions of these calculations on the loose in 450

2. Asteroid characteristics?

3. How to use/open the cycler data?

4. Defining LEO, LLO, LMO, etc.

The Plan:

1. Get rough ∆v’s with Hohmann transfers, and refine via hyperbolic orbits and

Lambert arcs to get upper limits. (MATLAB → GMAT)

2. Condense (and verify) everyone’s ∆v calculations into one spreadsheet for all

teams’ use (Excel)

Cycler 1 additions for ∆v calculation

# Date (m/d/y) Vinf Mag (km/s) Vinf x Vinf y Vinf z Altitude (km) TOF (days)DV per Segment (m/s)

DV per Cycle (m/s)

LLO to cyc delta v

LMO to

Cycl delta v

E-1 6/30/2020 4.22 3.46477 2.400842 0.20985 0 -2.591

M-2 12/15/2020 4.2 3.622656 1.933584 -0.878819 10642 168 0 -0.839

E-3 5/24/2023 5.91 4.902702 2.980457 1.397518 24853 890 0 -4.281

E-4 11/11/2024 5.92 3.567176 4.728563 0.001682 28290 537 0 0 -4.291

M-5 6/10/2025 3.37 -2.396001 2.366752 0.110712 16336 210 0 -0.009

E-6 8/15/2027 6 -2.331498 5.374726 -1.294772 26683 797 0 -4.371

E-7 1/31/2029 6 -4.122487 4.359484 -0.001202 52902 534 125 120 -4.371

M-8 9/18/2029 3.79 -3.585596 0.153237 1.231092 2706 230 0 -0.429

E-9 11/8/2032 4.44 -4.110094 -1.380301 0.979075 5027 1147 0 -2.811

E-10 5/6/2033 4.36 -0.239992 0.214641 -4.351776 5810 179 0 0 -2.731

M-11 9/15/2033 5 1.130976 -4.869649 -0.086068 1799 132 0 -1.639

E-12 3/2/2037 4.57 0.924994 -2.636173 -3.617045 28961 1264 98 -2.941

E-13 9/4/2037 4.5 -0.119045 -0.326172 4.481967 8887 186 0 100 -2.871

M-14 4/15/2038 2.53 0.854312 2.128033 -1.080243 1861 223 0 0.831

E-15 6/15/2040 6 3.516642 4.632249 -1.469633 22929 792 324 -4.371

E-16 12/3/2041 6 1.627477 5.77506 0.001464 30716 536 70 390 -4.371

M-17 6/10/2042 4.55 -4.243753 1.593782 0.416359 14512 189 0 -1.189

E-18 9/4/2044 6 -4.05995 4.228187 -1.28033 26408 817 0 -4.371

E-19 2/24/2046 6 -5.530367 2.327023 -0.001879 67547 538 8 10 -4.371

M-20 8/6/2046 5 -3.894554 -3.049263 0.731057 17025 163 0 -1.639

E-21 11/9/2049 5.2 -4.667327 -2.075194 0.979076 1556 1191 22 -3.571

E-22 5/8/2050 5.11 -0.311251 0.295232 -5.087072 24056 179 0 20 -3.481

M-23 12/31/2050 3.73 1.465533 -2.720309 2.082792 1107 237 0 -0.369

E-24 3/28/2054 4.68 4.341338 -1.754439 -0.112275 1000 1183 131 -3.051

E-25 9/30/2054 4.67 0.044935 -0.376601 4.654238 5810 186 0 130 -3.041

M-26 4/26/2055 3.3 -1.05132 3.08039 -0.563101 0 208 0

min -2.591 0.831

max -4.371 -1.639

average -3.640412 -0.868

• The last two columns on the

right detail the ∆v between the

Cycler 1 vehicle and the

planetary orbit.

• Since the Cycler vehicle goes

at different speeds at different

encounters, our ∆v changes

as well.

• Averages were used in the

presentation to the class

Peter Salek February 4, 2020

Discipline: Power and ThermalPowering the Mass Driver

1

Code

[3] Rani, J., Thangavel, R., Oh, S.-I., Lee, Y., and Jang, J.-H., “An Ultra-High-

Energy Density Supercapacitor; Fabrication Based on Thiol-functionalized

Graphene Oxide Scrolls,” Nanomaterials, vol. 9, 2019, p. 148.

[3]

Josh SchmeidlerFebruary 6th, 2020

Power & ThermalThermal Analysis of Taxi Vehicle

Backup Slides

MATLAB Code References

[1] Benson, Tom. “Mars Atmosphere Model - Metric Units.” NASA, NASA,

12 Nov. 2014, www.grc.nasa.gov/WWW/K-12/rocket/atmosmrm.html.

[2] Girija, Athul Pradeepkumar, et al. “Feasibility and Mass-Benefit Analysis of

Aerocapture for Missions to Venus.” Journal of Spacecraft and Rockets, vol. 57,

no. 1, 24 Jan. 2020, pp. 58–73., doi:10.2514/1.a34529.

Joe TiberiFebruary 6th, 2020

Backup - Propulsion Team LeadED Tether

Calculation MethodsTether Velocity During Spinup

Power required during spinup

Matlab Scripts

Shuting YangFebruary 06, 2020

PropulsionPhobos Tether Sling

Backup Slides

Slide 1 of 6

Appendix A: References

1. Koppel, R.C., “Optimal Specific Impulse of Electric Propulsion”, Second European

Spacecraft Propulsion Conference, Edited by Michael Perry. ESA SP-398. Paris:

European Space Agency, 1997., p.137. Retrieved February 1, 2020 from

http://articles.adsabs.harvard.edu//full/1997ESASP.398..131K/0000137.000.html

2. Puig-Suari, J., Longuski, J.M., and Tragresser, S.G., “A Tether Sling for Lunar and

Interplanetary Exploration,” Acta Astronautica, Vol. 36, No. 6, 1995, pp. 291-296.

3. Redd, N. T. (2017, December 8). Phobos: Facts About the Doomed Martian Moon.

Retrieved February 3, 2020, from https://www.space.com/20346-phobos-moon.html

4. Robert, A. (2008). Rocket Propellants. Retrieved February 3, 2020, from

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

5. “Ultra High molecular Weight Polyethylene fiber from DSM Dyneema,” eurofibers,

CIS YA100, January 2010.

Slide 2 of 6

http://articles.adsabs.harvard.edu/full/1997ESASP.398..131K/0000137.000.html

https://www.space.com/20346-phobos-moon.html

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

Appendix B: MATLAB Code

Slide 3 of 6


Slide 4 of 6


Slide 5 of 6


Slide 6 of 6

Natasha Yarlagadda - Backup Slides February 6, 2020

Propulsion TeamTaxi - Mars Landing

133

Backup: Initial System Studies

Mission System System Weight (Mg) Engine Propellant

Thrust

(N) Isp (s)

Engine

Weight (Mg)

Mars Insight

Thrusters during atm entry,

parachute descent, 12 descent

engines 0.61

Aerojet Rocketdyne

MR-107K Hydrazine 222.4 226 0.00091

Mars PathfinderAeroshell entry capsule, supersonic

parachute, 3 solid rockets 0.89 Solid tractor rockets N/A N/A N/A N/A

Mars CuriosityGuided entry, parachute descent, 8

rocket engine descent, sky crane 0.9

Aerojet Rocketdyne

MR-80B Hydrazine 3113 213 0.00851

SpaceX Falcon 9Grid fins to guide into atm, cold gas

thrusters to flip, 3 back boost engines 30.39 Merlin 1D

LOX/Kerose

ne 854000 311 0.47

Blue Origin New

Shepard

8 large drag brakes, hydraulically

actuated fins to steer, engines to

slow 34 BE-3 LOX/LH2 90000 N/A N/A

Space Shuttle 77 RS-25 LOX/LH2 1860000 366 3.17

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Backup: References

[1] “EXTERNAL TANK,” NASA Available: https://science.ksc.nasa.gov/shuttle/technology/sts-newsref/et.html.

[2] “Shuttle technical facts,” ESA Available:

https://www.esa.int/Science_Exploration/Human_and_Robotic_Exploration/Space_Shuttle/Shuttle_technical_facts.

[3] “RS-25 Engine,” RS-25 Engine | Aerojet Rocketdyne Available: https://www.rocket.com/space/liquid-engines/rs-25-engine.

[4] “Propellant Tanks,” Purdue Engineering Available: engineering.purdue.edu › reportfinaluploads › appendix › structures

[5] Pietrobon, S. S.,“Analysis of Propellant Tank Masses,” NASA Available: https://www.nasa.gov/pdf/382034main_018%20-

%2020090706.05.Analysis_of_Propellant_Tank_Masses.pdf

[6] Greene W. D., “Space Shuttle Main Engine – Liquid Rocket Engines (J-2X, RS-25, general),” NASA Available:

https://blogs.nasa.gov/J2X/tag/space-shuttle-main-engine/.

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https://science.ksc.nasa.gov/shuttle/technology/sts-newsref/et.html

https://www.esa.int/Science_Exploration/Human_and_Robotic_Exploration/Space_Shuttle/Shuttle_technical_facts

https://www.rocket.com/space/liquid-engines/rs-25-engine

https://engineering.purdue.edu/AAECourses/aae450/2008/spring/report_archive/reportfinaluploads/appendix/structures/A.5.2.1.3%20Preliminary%20Design%20-%20A.5.2.1.5%20Effects%20of%20PropellantTypes.doc

https://www.nasa.gov/pdf/382034main_018%20-%2020090706.05.Analysis_of_Propellant_Tank_Masses.pdf

https://blogs.nasa.gov/J2X/tag/space-shuttle-main-engine/

Rachel RothFebruary 6, 2020

Backup Slides

Rachel Roth Backup Slides – Material PropertiesAl 7075-T7351 [5]

Rachel Roth Backup Slides – Calculations [4]

Rachel Roth Backup Slides – References[5] Aluminum Association, Inc. (2001). Aluminum 7075-T73; 7075-T735x. Retrieved from

http://asm.matweb.com/search/GetReference.asp?bassnum=MA7075T73.

[6] Federal Aviation Administration. (2003). Metallic Materials Properties Development and

Standardization (DOT/FAA/AR-MMPDS-01). Retrieved from http://everyspec.com/FAA/FAA-

General/DOT-FAA-AR-MMPDS-01-JAN2003_24102/.

[7] National Aeronautics and Space Administration. (2014). Structural Design Requirements and Factors

of Safety for Spaceflight Hardware (JSC 65828). Retrieved from

https://standards.nasa.gov/standard/jsc/jsc-65828.

[8] National Aeronautics and Space Administration. (2011). Standard Materials and Processes

Requirements for Spacecraft (NASA-STD-6016). Retrieved from

https://standards.nasa.gov/standard/nasa/nasa-std-6016.

http://asm.matweb.com/search/GetReference.asp?bassnum=MA7075T73

http://everyspec.com/FAA/FAA-General/DOT-FAA-AR-MMPDS-01-JAN2003_24102/

https://standards.nasa.gov/standard/jsc/jsc-65828

https://standards.nasa.gov/standard/nasa/nasa-std-6016

Documents

Section III Presentations - engineering.purdue.edu · The Problem: Determine the Configuration of the Cycler Vehicle Requirements • The cycler must accommodate three docked taxi