final design review 3/12/2015 - purdue engineering · tomi olokun nicole vaughn mission design...

FINAL DESIGN REVIEW

3/12/2015

AGENDA • Schedule

• Mission Design

• XM3 Human Habitats

• Entry Descent and Landing (EDL)

• Human Lander and Rovers

• Cargo Missions

• Cycler Vehicle

• Return Option

• Communications

• Vehicle Mass & IMLEO

• Risk Assessment

MISSION SCHEDULING

• DEVELOPMENT PLAN

• PRODUCTION NEEDS

STEPHEN WHITNAH

PROJECT MANAGER

03/12/2015

DEVELOPMENT SCHEDULE

2015 2020 2025 2030 2035 2040

50 years since Apollo 11

ISS Transitioned (Extension to 2028?)

SLS Block 2 Design Completed

Orion and SLS-1B development ends (EM-2)

Estimated milestones for current NASA human spaceflight programs

XM3, Rover, Crane

Whitnah 4

Human Lander

Return Option

Cycler and Boost Module, Cargo Vehicle

Mars mission hardware production begins

Human Exploration of Mars

Proposed vehicle development schedule

DEVELOPMENT SCHEDULE

2015 2020 2025 2030 2035 2040

Rigid XM Missions to

L1, L2, asteroid,

moon bases


Proposed vehicle development schedule

XM3, Rover, Crane

Whitnah 5

Cycler and Boost Module, Cargo Vehicle

Human Lander

Return Option


Inflatable BA 330

Missions



Cycler Establishment

A B Human Launches

A B

Cargo Missions

PRODUCTION SCHEDULE

2015 2020 2025 2030 2035 2040

Rigid XM Missions to

L1, L2, asteroid,

moon bases


Whitnah 6


Inflatable BA 330

Missions



Cycler Establishment

A B Human Launches

A B

L1: I L2: I Ast: I

L1: R L2: R Asteroid: R

Moon Far: RRR Near: RRR Shackleton: RRR

XC XC XC C B

XC XC XC C B

XM XM

XM XM XP R

XM XM XP L L L B

XM XM XM

L L L B

Vehicle need schedule

I – Inflatable BA 330 R – Rigid XM2 XC – XM3 (Cycler) C – Cycler core XP – XM3 (Phobos) XM – XM3 (Mars) L – Human lander B – Boost vehicle R – Return vehicle

Cargo Missions

PRODUCTION SCHEDULE

2015 2020 2025 2030 2035 2040


Vehicle need schedule

I – Inflatable BA 330 R – Rigid XM2 XC – XM3 (Cycler) C – Cycler core XP – XM3 (Phobos) XM – XM3 (Mars) L – Human lander B – Boost vehicle R – Return vehicle

Whitnah 7


L1: I L2: I Ast: I

L1: R L2: R Asteroid: R

Moon Far: RRR Near: RRR Shackleton: RRR

XC XC XC C B

XC XC XC C B

XM XM

XM XM XP R

XM XM XP L L L B

XM XM XM

L L L B

Summary of Launch Needs (Heavy Lift Vehicles)

Moon: I – 3 R – 3 (US, the 9 for moon bases assumed to be international partners) Total US launches: 6 Total International launches: 9 Mars: XC – 6 C – 2 XP – 2 XM – 9 L – 6 B – 4 R – 1 Total US launches: 30 (Not counting resupply or additional cargo missions)

QUESTIONS?

PROJECTED MANUFACTURING SCHEDULE

Whitnah 9

2028 2030 2032 2034 2036 2040

Cycler A - Humans

XM3-C

2038

XM3-C

XM3-C

XM3-C

XM3-C

XM3-C

XM3-M

XM3-M

XM3-M

XM3-M

XM3-M

XM3-M

XM3-P XM3-P

XM3-M

XM3-M

XM3-M

Human Lander

Human Lander

Human Lander Lander

Lander

Lander

Cycler B - Humans

Cycler Boost Return Cycler Boost Cycler Boost

Note: Does not include cargo missions, cranes, or rovers – estimated 3 cargo per synodic period

Cycler Boost

EMILY ZIMOVAN

ALEX DAVIS

LORENZO GARCIA

PETER GELDERMANS

PABLO MACHUCA

TOMI OLOKUN

NICOLE VAUGHN

MISSION DESIGN

FINAL DESIGN REVIEW

MARCH 12, 2015

STORY BOARD

Zimovan

S1L1 Cyclers A/B & Communications

Depart Cyclers

Cargo Supply

P-M Return to Earth

Hyperbolic Rendezvous

LEO Ops Moon

ComSats

S1L1 – A MORE ACCURATE MODEL*

Zimovan

*McConaghy, T. T., “Design and Optimizatoin of Interplanetary Spacecraft Trajectories,” Ph.D. Dissertation, Aeronautical and Astronautical Engineering Dept., Purdue Univ., West Lafayette, IN, 2004.

Date Maneuver V∞ [km/s] Altitude [km]

02/19/2031 Launch Cycler A (Low Thrust Spiral Begins) ○ ~ 0 400

04/29/2033 Hyperbolic Checkout Cargo to Cycler A (Flyby Earth) 4.47 24,950

04/19/2033 Launch Cycler B (Low Thrust Spiral Begins) ○ ~ 0 400

08/18/2033 A Cargo to Phobos and Mars (Flyby Mars) 7.58 7,070

06/28/2035 Hyperbolic Checkout Cargo to Cycler B (Flyby Earth) 4.20 2,756

11/12/2035 B Cargo to Phobos and Mars (Flyby Mars) 5.87 1,170

08/20/2037 Humans to Cycler A (Flyby Earth) 4.74 617

01/22/2038 Humans to Phobos, Cycler A (Flyby Mars) 5.66 1,454

10/26/2039 Humans to Cycler B (Flyby Earth)† 5.53 23,900

03/02/2040 Humans to Phobos and Mars, Cycler B (Flyby Mars)† 4.31 17,600

11/19/2041 Humans to Cycler A (Flyby Earth)† 7.02 37,400

04/28/2042 Humans to Phobos and Mars, Cycler A (Flyby Mars)† 5.89 9,800

11/10/2043 Humans to Cycler B (Flyby Earth)† 6.43 41,500

03/27/2044 Humans to Phobos and Mars, Cycler B (Flyby Mars)† 7.14 12,200

Crew TOF 155 days

Crew TOF 183 days

†Estimated using return to initial inertial positions of planets in Sun-frame & McConaghy, T. T., Landau D. F., Yam C. H., and Longuski, J. M., “Notable Two-Synodic Period Earth-Mars Cycler,” Journal of Spacecraft and Rockets, Vol. 43, No. 2, 2006, pp. 456-465.

○From Alex Davis

Crew TOF 138 days

Crew TOF 160 days

FUEL MASS FOR TCM FOR CYCLERS

Zimovan

TCM: ∆V = 200 m/s per complete cycle (4 2/7 years), as required in specifications.

STOUR Calculations have validated this ∆V for Trajectory Correction Maneuvers!!

Electric Propulsion Assuming: Isp = 3000 sec

Mcycler = 192.5 Mg Mlanders =115.9 Mg total

Vehicle Mass of Vehicle At TCM [Mg] TCM Fuel Mass [Mg]

Cycler (No Landers) 192.5 1.313

Cycler + 3 Landers 308.4 2.103

Rogers, B. A., Hughes, K. M., Longuski, J. M., Aldrin, B., “Establishing Cycler Trajectories Between Earth and Mars,” Unpublished. Version Dec. 15, 2014.

For ~88% of the orbit, cycler is in this orientation

Weighted Average to Approximate: TCM Fuel Mass = 1.405 Mg

OPTIMAL LAUNCH SOLUTION FROM MARS

Zimovan

Assumptions: •“Flat” Mars • Exponential atmosphere model • CD = 0.8 (drag coefficient) • Averaged mass flow rate = 85.59 kg/s

Calculus of Variations used to solve for the time-optimal launch solution

(minimum fuel required)

HYPERBOLIC RENDEZVOUS WITH CYCLER

15-Garcia

• Assume departure from Periapsis

• Need to shift orbit in order

to intersect

• TOF = 4 -6 days *

Line of Apsides

Shift in Apsides 28.2°

*done in collaboration with Sam Ferdon from Human Factors and Cory Back In Propulsion

MISSION DATES AND REQUIREMENTS

16-Garcia

Date (mm/dd/yyyy)

Vinf of Cycler (km/sec)

Rp of Cycler (km)

Description of the Mission

Total ΔV Required (km/sec)

TOF Required (days)

4/29/2033 4.47 31,328

A test run of the hyperbolic rendezvous with cargo 4.78 4

6/28/2035 4.2 9,134

A 2nd test run of the hyperbolic rendezvous with cargo 4.66 4.25

8/20/2037 4.74 6995.137 Humans to Cycler A 4.88 4

10/29/2039 5.53 30278.137 Humans to Cycler B 5.26 4

AVERAGE: 4.895 4.0625

MARS COMMUNICATION CONSTELLATION

1 - MACHUCA

Two areostationary satellites: Continuous coverage is possible with two satellites around Mars

• 17032 km altitude orbit, 63.9° < 𝛼 < 70.5° • Three antennas on Mars, three antennas on satellites, two antennas on Phobos

COMMUNICATION VIA CYCLERS

2 - MACHUCA

Communication via cyclers provide continuous communication link

• Two antennas on each cycler

LANDERS: RENDEZVOUS IN LEO

• Individual launches and rendezvous in

200 km altitude

• Rendezvous: 3 lander vehicles

• Total DV: 0.006 km/s

Rendezvous orbit statistics:

Olokun 19

Minimum DV:

190 km chasing

altitude Docking Altitude 200km

Period 1.51 hours

Chasing Altitude 190km

Mass Per Lander 24.686 Mg

DV Per Rendezvous 0.002 km/s

Total DV 6 m/s

Time Per Maneuver 0.59 hours

HUMAN LANDER TRAJECTORY DESIGN

Geldermans-20

Trajectory selected to reduce TOF with acceptable cost to total system mass.

27.418 Mg

26.149 Mg

TOF to Minimize System Mass is Unbounded

Human Health Concerns

Minimize TOF with Acceptable System Mass Cost

Fixed Vehicle Mass

Overall mission objective as defined by Project Aldrin-Purdue Mission

Specifications is to Minimize IMLEO

LANDERS: PHOBOS TO MARS

Olokun 21

4 optimal trajectories

• Low Mars Orbit requirements for 120 km altitude: • flight path angle between 0 and -20 degrees • Velocity <= 4.9km/s

True Anomaly 0

Delta V1 0.628

Delta V2 0.8805

Total Delta V 1.5085

• Lowest energy was chosen – Hohmann transfer: • Total DV = 1.51km/s

ELECTRIC PROPULSION ESTABLISHMENT OF CYCLER

Davis-22

Cycler Launch Establishment Swingby Swingby 1

Vehicle Power (kW) Cycler Launch Mass (Mg)

Establishment Swingby Mass (Mg)

Electric Propellant Mass Used(Mg)

Cycler A 250 214.4 192.5 21.90

Cycler B 240 217.5 192.5 25.04

Impulsive Delta V = 3.175 km/s To Enter Heliocentric Frame

Launch: 2/19/31 Established: 4/29/33

Launch: 5/9/33 Established: 6/28/35

POSSIBLE FIRST CARLA LAUNCH DATES

23

Several Low Cost Options in Ephemeris Models. Launch Date Selected: 7/1/35

HOHMANN TRANSFER AND AEROCAPTURE

Davis-24

Aerocapture

EDL

Cargo and Hab Earth Parking Alt [km]

Beta [kg/𝒎𝟐]

Delta V to reach Mars [km/s]

Periapsis Alt to Capture[km]

V at Entry[km/s]

Mars Case 400 20 3.667 63.4930 5.65006 (EDL)

Phobos Case 400 20 3.667 80.3351 5.65006 (Aerocapture)

*Aerocapture code adapted from Ben Tacket, Cynthia Rose, Peter Geldermans

PHOBOS RENDEZVOUS

• If Captured Orbit Apoapsis is within

1% of Phobos’ Orbit Radius, linear targeter

is a valid approximation

• Requires a 1/6km accuracy of

aerocapture entry altitude.

Davis-25

Captured Apoapsis Radius [km]

Delta V to Rendezvous [km/s]

Time to Rendezvous [hours]

9176 .5369 7.65

RETURN TO EARTH TRAJECTORY

Jan 22, 2038-Jul 13, 2038

TOF: 6.7 Months

Departure Delta V: 2.27 km/s

Vaughn

Aug 06, 2039-Feb 26,2040

TOF: 6.8 Months

Departure Delta V: 2.17 km/s

LUNAR OPERATIONS

Olokun 27

• Humans - Hohmann Transfer: 200 km LEO to 200 km LLO

ΔV1 [km/s] ΔV2 [km/s] Total ΔV [km/s] TOF [days]

3.1313 1.4040 4.5353 4.978

• Humans - Free return option from LLO

ΔV1 [km/s] ΔV2 [km/s] Total ΔV [km/s] TOF [days]

3.1414 0.8793 4.0207 1.437

• XM1 on stable manifold to L1 and L2

Point L2

Distance 64627.97km

DV1 to L2 3.0957 km/s

DV2 to L2 Halo 0.1010 km/s

TOF 90.9461 days

Point L1

Mass 7.3134 kg

Distance 58086.91km

DV to L1 3.2249 km/s

TOF 76.4852 days

• Cargo landers on low thrust trajectory to LLO – Propellant cost: 118.1Mg

• Refueling station orbit characteristics: Altitude 17000 km

Velocity 4.082 km/s

Period 10 hours

TOP FIVE MISSION DESIGN RISKS

Zimovan

• S1L1 Correction Maneuvers – If failed, may impact surface of Earth/Mars or miss next flyby crew is lost! Likelihood: 1 Consequence: 5 • Hyperbolic Rendezvous – Has never been attempted before, if failure occurs crew is lost! Likelihood: 3 Consequence: 5 • On-orbit Based Communications – Failure of a mission specification, could lose communications completely Likelihood: 2 Consequence: 3 • Aerocapture – Bounce off of atmosphere if too shallow, or burn up/lose control if too steep Likelihood: 2 Consequence: 5 • Cargo Trajectories – Lose significant amount of supplies if failure occurs (2 1/7 years before more cargo can be delivered) Likelihood: 1 Consequence: 4

Likelihood of Failure: 1 = Low … 5 = High

Consequence if Failure Occurs: 1 = Mild … 5 = Severe

QUESTIONS?

VERIFICATION OF SPECIFICATION ∆V

Zimovan

Event Value*

V∞ at Earth Flyby 3.98 – 7.09 km/s

V∞ at Mars Flyby 2.77 – 7.88 km/s

TOF From Earth to Mars 111 – 231 days

STOUR gives >100,000 solutions… Manually match two solutions to find two EMEE S1L1 cycles that connect: Cycle 1 depart E: 2033/04/29 Cycle 1 arrive at E: 2037/08/22, V∞ = 4.509 km/s Cycle 2 depart E: 2037/08/22, V∞ = 4.6 km/s

Burn at geocentric “∞”: ∆V = 95.4 m/s or 8.567 Mg fuel

Burn at hyperbolic periapsis: ∆V = 35.7 m/s or 3.173 Mg fuel

(for Isp = 300 s, Mcycler = 260 Mg)

*McConaghy, T. T., “Design and Optimizatoin of Interplanetary Spacecraft Trajectories,” Ph.D. Dissertation, Aeronautical and Astronautical Engineering Dept., Purdue Univ., West Lafayette, IN, 2004.

BACKUP

Geldermans-31

Trajectory selected to reduce TOF with acceptable cost to total system mass.

27.418 Mg

26.149 Mg

34.104 Mg

32.526 Mg

FINAL DESIGN REVIEW • FINALIZED DIMENSIONS

• FINALIZED MASS, POWER, AND VOLUME

• VARIANT OVERVIEWS

• FLOOR LAYOUTS

• COLONY LAYOUT

• PHOBOS LAYOUT

• POWER

XM3 VEHICLE GROUP

3/12/15

CALVIN EADS, JUSTIN GUASTAFERRO, ROBERT SKIDMORE, ANDREW

BOKHART, QIRONG LIN, AND HAONAN ZHANG

FINALIZED DIMENSIONS

Eads 33

Height (m)

Diameter (m)

Wall Thickness (m)

Floor Thickness (m)

Empty Volume (𝒎𝟑)

Structural Volume (𝒎𝟑)

Structural Mass (Mg)

9.32 7.6 0.02 0.015 404.6 7.648 22.11

Volume: 97.07 m3

Volume: 103.8 m3

Volume: 103.8 m3

Volume: 103.8 m3

FINALIZED MASS, POWER, AND VOLUME

Eads 34

Totals XM3-C XM3-P XM3-M (Core)

XM3-M (Farming)

XM3-M (Water)

XM3-M (Medical)

Mass (Mg) 42.91 45.61 43.38 36.17 43.75 43.03

Systems Volume (𝒎𝟑) 274.8 258.7 320.6 17.42 244.1 255.5

Power (kW) 63.17 55.57 120.9 35.98 106.97 112.0

XM3-M VARIANTS

Eads 35

Crew Quarters

Crew Quarters

Crew Quarters

Water Systems

Life Support Systems Power

Systems

Thermal Control Systems

XM3-M (Core)

Farming Floor

Farming Floor

Farming Floor

Water Systems


Systems


XM3-M (Farming)

XM3-M VARIANTS (CONTINUED)

Eads 36

Med Bay

Crew Quarters

Crew Quarters

Water Systems


Systems


XM3-M (Medical) XM3-M (Water)

Water Systems

Crew Quarters

Crew Quarters

Life Support Systems

Power Systems


Crew Quarters

Crew Quarters

Crew Quarters

CREW QUARTERS LAYOUT

Pottebaum-37

Stowage Food

Tunnels

Bathroom Personal Space

Maintenance/Tools

Washer/Dryer

Exercise Area

MED BAY LAYOUT

Eads 38

Med Bay

General Med Racks

Medical Tables

Storage Racks

UTILITIES FLOOR LAYOUT (STANDARD)

Skidmore - 39

FARMING IN THE XM-3 • Available Farming Space

• Number of XM-3: 2

• Number of Floors: 3

• Number of People: 18

• Area of Floors: 29.40 m2

• Farming Space Allocation (%)

• Potatoes: 50%

• Soybeans: 25%

• Pinto Beans: 12.5%

• Wheat: 12.5%

• Capabilities

• Numbers obtained for 1 year of farming

• Sustainability Goal: 40% of required calories

• Remainder from CarLa Farms

• Due to changes in the structure of the XM-3 which involved lowering the ceiling, two level farming is

no longer a viable option.

Bokhart 40

% Calories Yield (Mg)

6.91 1.02

Farming Area Walkways

SPECIALIZED WATER PROCESSING/STORAGE

Skidmore - 41

• Intended to process and store water harvested from Mars

• Primary water supply for farming floors

• Increases water storage capacity by 7.6 Mg

• Rack mounts for tanks and hardware take up 20.6 m^3 of volume

• Total mass of high-capacity processing system and empty tanks: 2.84 Mg

COLONY LAYOUT

• 2m long connectors between

modules

• Accordion style designed to

be slightly flexible. Likely to

be made of a combination of

aluminum and durable

plastics.

• Leveling legs to account for

slightly uneven ground

Eads 42

Farming Modules

Medical and Extra Core Module

Water Modules

Core Living Modules

PHOBOS SETUP

Eads 43

• Centrifuge located on ground for easier access via ports

• Ground location also allows for greater radius of rotation

• Crew uses centrifuge for 1-2 hours a day per crew member

Radius of Rotation (m)

Mass (Mg)

Volume (𝒎𝟑)

RPM to generate .38g

8 3.349 4.571 6.517

Anti Gravity Centrifuge

DOME SHIELD DESIGN

Guastaferro 44

Mass Volume (packed)

Diameter Minimum thickness

Height

13.7 Mg 8 .37 m3 60 m 0.4 m 18.4 m

Inflated Wall Structure • Fabric is 200 denier Vectran • Gas is Nitrogen

• Roof at 85 KPa • Wall at 100 KPa

• Structure inflated by detonating Sodium Azide (NaN3)

XM3-M POWER • Nuclear power production unit will

be entirely outside of XM-3 and outside of dome shielding

• Closed Brayton Cycle with Regeneration for Thermo-Electric conversion with 24.53% efficiency

45

Site Power Required

(kWe)

Thermal Power (kWt)

Brayton Converter Mass (Mg)

Reactor Mass (Mg)

Total Mass (Mg)

Phobos 67.32 274.44 1.43 0.71 2.14

Mars 196.896 802.67 3.04 2.07 15.05

Mars • 4 XM3-C (Core livable modules) • 2 XM3-F (Farming modules) • 2 XM3-W (Water modules) • 1 XM3-M (Medical module) Phobos

• 2 XM3-P (Core livable modules)

RISK ASSESSMENT

Skidmore - 46

(* Denotes Pre-Mitigation)

1) Micrometeorite Puncture

2) Buckling of cylindrical section during EDL

3) Floor failure during EDL

4) Fewer than 9 usable modules on Mars surface

5) Cabin fire

6) Water leak

7) Life support component failure

4* 4

6,7

1*, 2

1

3 5*

5

6* 7*

QUESTIONS?

SYSTEMS MASSES (MG)

Eads 48

System XM3-C XM3-P XM3-M (Core) XM3-M (Farming)

Structural 22.11 22.11 22.11 22.11

Communication 2.2 2.2 2.2 2.2

Power Systems 0.0953 0.135 0.0953 0.0953

Med Bay 2.12 2.12 0 0

Crew Quarters 3.112 4.336 7.416 0

Water Systems 3.75 6.065 6.065 6.065

Life Support 1.69 1.255 1.255 1.255

Thermal Control 4 2.9 .6 .16

Farming 0 0 0 .645

Micro Meteorite Shielding

3.837 3.641 3.641 3.641

SYSTEMS MASSES (MG) CONT.

Eads 49

System XM3-M (Water) XM3-M (Medical)

Structural 22.11 22.11

Communication 2.2 2.2

Power Systems 0.0953 0.0953

Med Bay 0 2.12

Crew Quarters 4.944 4.944

Water Systems 8.905 6.065

Life Support 1.255 1.255

Thermal Control .6 .6

Farming 0 0

Micro Meteorite Shield 3.641 3.641

SYSTEM VOLUMES (M^3)

Eads 50


Structural 7.648 7.648 7.648 7.648

Communication 0.81 .81 .81 .81

Power Systems 0.003969 0.05869 0.03969 0.03963

Med Bay 32 32 0 0

Crew Quarters 207.6 207.6 304.7 0

Water Systems 3.33 7.33 7.33 7.33

Life Support 8.6 7.1 7.1 7.1

Thermal Control 22.4 1.6 .64 .14

Farming 0 0 0 2

SYSTEM VOLUMES (M^3) CONT.

Eads 51


Structural 7.648 7.648

Communication .81 .81

Power Systems 0.03969 0.03963

Med Bay 0 32


Water Systems 27.93 7.33


Thermal Control .64 .64

Farming 0 0

SYSTEM POWER REQUIREMENTS (KW)

Eads 52


Communication 1 1 1 1

Power Systems 0 0 0 0

Med Bay 5 5 0 0

Crew Quarters 19.77 21.47 41.81 0

Water Systems - - - 3.075

Life Support 29.4 21.2 62.1 8.005

Thermal Control 8 1.9 16 16

Farming 0 0 0 7.9

SYSTEM POWER REQUIREMENTS (KW)

Eads 53


Communication 1 1

Power Systems 0 0

Med Bay 0 0


Water Systems - -


Thermal Control 16 16

Farming 0 0

LIGHTING INFORMATION AND DESIGN

• XM-3 Farming Variant: 1

• Floors: 3

• Light Fixture: Halo (SGL6)

Bokhart 54

Power (KW)

Volume (m3)

Mass (Mg)

Heat (W)

Number of Units

8.2 2.0 0.67 820.0 328

XM3-P THERMAL CONTROL

Zhang 55

Thermal Load of the XM3-P at different time on Phobo

Vehicle Name Mass 𝑴𝒈 Power(𝒌𝒘) V𝐨𝐥𝐮𝐦𝐞(𝒎𝟐) Rejection Rate 𝒌𝒘

XM3-P 0.95 1.9 55 126.1

Related Subsystem

Function

Louvres Adjust rejection rate

Mechanical Arm Adjust radiation area

By-Pass Valve Adjust radiation level

Phase Changing Material

Balance temperature on XM

Related subsystem

Specification of the Active Thermal Control System for XM3-P

XM3 ON PHOBOS AND MARS

56

XM3 ON PHOBOS (Parabolic Antenna) XM3 ON MARS (Adept Antenna)

Mass[Kg] 6.4

Power[W] 250

Volume[m^3] 2.4E-3

Uplink: 2.8 GHz Downlink : 3.2 GHz Beamwidth : 19 deg Data rate: 12 Mbps Antenna Diameter : 0.75 m

HD Streaming Communication through satellite 0.024 sec signal delay

Mass[Kg] 57.3

Power[W] 5

Volume[m^3] 2.285

ADEPT antenna Uplink: 2.8 GHz Downlink: 3.2 GHz Transmit Antenna Beamwidth: 10 deg Data Rate : 12 Mbps Antenna Diameter : 30 m Fixed antenna HD Streaming Communication through satellite 0.14 sec signal delay Data rate can be increased significantly

CYNTHIA ROSE

BEN LIBBEN

BEN TACKETT

ROHAN DUDANEY

ZAK SIPICH

AERODYNAMICS

3/12/15

ACRONYMS AND TERMS: GENERAL

Rose 58

• EDL: Entry Descent and Landing, the process of going to the surface of a planet when there is a considerable atmosphere

• Aerocapture: passing through a planetary atmosphere while in a hyperbolic orbit in order to decrease velocity and get in an elliptic orbit about the planet

• Ballistic Coefficient: parameter relating mass and vehicle free stream surface area in atmospheric entry

• TPS: Thermal Protection System, protective heat layer for EDL • TD: Terminal Descent, low speed propulsive deceleration • SRP*: Supersonic Retro Propulsion, propulsive deceleration in

supersonic regime

*can also mean Subsonic Retro Propulsion, we refer to it as Supersonic retro propulsion

ACRONYMS AND TERMS: SYSTEMS

Rose 59

• HuLa: Human Lander, for Mars and Phobos • CarLa: Cargo Lander, for Mars and Phobos • Mars Return: the option for getting back to Earth from Phobos

with a HuLa • ADEPT: Adaptive Deployable Entry Placement Technology, a

semi-rigid/flexible decelerator option • HIAD: Hypersonic Inflatable Aerodynamic Decelerator, an

inflatable decelerator option

MARS ENTRY PATHS

Rose 60

Mars

Mars Atmosphere

Phobos Orbit

Incoming Trajectory (Not Direct Entry)

Phobos Bound Trajectory (post Aerocapture)

Non-Phobos Bound Trajectory (post Aerocapture)

Incoming Trajectory (Direct Entry)

ENTRY OPTION FOR SELECTED VEHICLE

Rose 61

Vehicle Landing Location

Entry Mass*

Entry Type Reason Entry Architecture

Carla Mars 72.6 Mg Direct Reduces error HIAD + Ballute

Carla Phobos 72.6 Mg Aerocapture Must exit hyperbolic orbit

HIAD + Ballute

Hula** Phobos 38.6 Mg Aerocapture Must exit hyperbolic orbit

ADEPT

Hula Mars 33 Mg Aerocapture G-Load ADEPT + Ballute

Mars Return***

Earth 23.6 Mg Direct G-Load HIAD + Parachute

*Entry Mass at Aerocapture **HuLa to Phobos has a ballute for Mars Entry ***Mars Return is a HuLa with a HIAD system

HYPERSONIC INFLATABLE AERODYNAMIC DECELERATOR (HIAD)

Sipich - 62

- Inflatable aeroshell used to reduce speed and protect payload from heat loading/heat flux encountered during entry into the atmosphere

- Inflates as a series of torus rings with constant radial distance to achieve a desired ballistic coefficient.

- HIAD provides less overall mass than ADEPT while still efficiently increasing the ballistic coefficient

LEFT: Isometric view of HIAD system Note: the vehicle attached is on the other side of HIAD

BELOW: Cross sectional view of HIAD system. The grey block represents the vehicle attached to HIAD

ADEPT INTRODUCTION

Rose 63

• Adaptable Deployable Entry and Placement Technology • Hypersonic decelerator similar to an umbrella

Center of Mass Shift

Center of Pressure

Center of Gravity

Sto

wed

in T

ensi

on

N

om

inal

Sta

te

(Dep

loye

d)

• Folded for launch to reduce payload volume, then deployed before AEDL to decrease the ballistic coefficient

• Can be gimbaled to create lift vector

without need to eject ballast mass • Since the design is a 70 deg sphere-cone, landed ADEPT can be reused as a communications array

Figure based off of “Adaptive Deployable Entry and Placement Technology (ADEPT): A Feasibility Study for Human Missions to Mars,” by Venkatapathy, Ethiraj pg. 18

ARCHITECTURE OPTIONS

Libben 64

Architecture 1 Architecture 2 Architecture 3 Architecture 1 Pros: • Multiple systems reduce risk of single point failure • Creates ease for separation • Separates ADEPT much closer to the ground • Bank control for entire descent sequence Cons: • Deploying ballute may cause structural damage to

ADEPT Architecture 2 Pros: • Multiple systems reduce risk of single point failure • Creates ease for separation Cons: • Separating ADEPT at high speed / altitude • No lift control once adept is dropped Architecture 3 Pros: • Simple • Only one system needs to separate Cons: • If ADEPT fails, mission fails • In order to account for atmospheric error, ADEPT

alone would be too large

ARCHITECTURE SELECTION

Libben 65

Low

Medium

High

Low Metric (Desireable)

Medium Metric (In-Between)

High Metric (Undesierable)

Human Architecture 1 Architecture 2 Architecture 3

EDL System Mass

(Mg)Low Low Medium

ADEPT Base Radius

(m)Low Low High

Altitude at 100 m/s

(km)Low Low Medium

Descent Time (min) Low Low Low

Peak g-load

(Earth g's)Medium Medium Low

% TPS Mass Low Low High

Complexity Medium Medium Low

Control Authority Low High Low

VEHICLE EDL CHOICE: ADEPT VS HIAD

Libben 66

CarLa ADEPT Option HIAD Option XM3 ADEPT Option HIAD Option

Total Entry Mass (Mg) 76.5 76.5 Total Entry Mass (Mg) 75.5 75.5

System Mass (Mg) 19 9.2 System Mass (Mg) 18.89 9.09

System Base Radius (m)

26.25 26.23 System Base Radius

(m) 26.08 26.05

Mass Savings (Mg) 9.8 Mass Savings (Mg) 9.8

HuLa ADEPT Option HIAD Option

Total Entry Mass (Mg)

23.5 23.5

System Mass (Mg) 9.2 6.2

System Base Radius (m)

14.55 14.54

Mass Savings (Mg) 3

• Mass savings by using HIAD for CarLa and XM3 are very large.

• Due to the frequency of cargo missions versus human landings, an inflatable decelerator would most likely reduce mission cost and complexity.

• Mass saving by using HIAD on HuLa is not as large, and the need for ballast mass when using HIAD for bank angle control reduces the savings even more.

• Ability to reuse ADEPT shield for communications also saves mass and power.

Sipich - 67

FINAL HIAD SIZING FOR MARS ENTRY

Final CarLa HIAD Sizing Total Mass of CarLa = 70.39 Mg

Target Ballistic Coefficient = 20 kg/m2

Mass of HIAD = 6.89 Mg

Volume of Deployed HIAD = 3980 m3

Max Radius (rb) = 24.0 m

Torus Radius (rt) = 1.4 m

Final XM3 HIAD Delivery Sizing Total Mass of XM3 = 73.14 Mg






Maximum Radius (rb) Torus Radius (rt)

HIAD SIZING FOR EARTH RETURN MISSION

Sipich - 68

HIAD Sizing for Return Mission Total Mass of Return Vehicle = 21.6 Mg






Maximum Radius (rb) Torus Radius (rt)

ATMOSPHERIC ERROR REDUCTION

Tackett 69

1. The addition of a lifting body eliminates downrange error due to atmospheric variation.

2. To decrease terminal velocity error, a Ballute will be deployed.

3. Failure due to atmospheric error is negligible with a lifting body and the error will be based on the control system and atmospheric condition information during descent.

Atmospheric Error

(% of average density)

Downrange Distance

(km)

Peak Decleration

(Earth g's)

Terminal Velocity

(m/s)

L/D = 0.0, atme = 100% Acceptable Acceptable Acceptable

L/D = 0.0, atme = 160% Unacceptable Acceptable Acceptable

L/D = 0.0, atme = 40% Unacceptable Acceptable Unacceptable

L/D = 0.2, Added Ballute, atme = 160% Acceptable Acceptable Acceptable

L/D = -0.2, Added Ballute, atme = 40% Acceptable Acceptable Acceptable

HuLa mass ≈ 23 Mg β = 20 kg/m2

CARLA AND XM3 DIRECT ENTRY ANALYSIS

Tackett 70

1. Direct entry decreases risk by eliminating the aerocapture error.

2. Atmospheric error reduction fulfills landing requirements for CarLa and XM3 EDL

3. All other landing parameters are within acceptable ranges.

* Entry Conditions supplied by Alex D * β Provided by Zak Sipich

RETURN OPTION: EARTH ENTRY

Tackett 71

1. HIAD is used for Earth Entry

2. Direct Entry reduces Peak G load

3. Parachutes will be needed for terminal descent

4. Peak G load is slightly above 6 Earth g’s for only 44 seconds

* Entry Conditions supplied by Nicole V. & Cynthia R. * β Provided by Zak Sipich

β = 25 kg/m^2 Mass = 21.6 Mg

Ventry (km/s)

Entry Velocity

Vterminal (m/s)

Terminal Velocity

Peak q̇ (W/cm^2)

Maximum Heat Flux

Total Q (J/cm^2)

Heat Load

Peak G-Load (Earth g's)

Maximum Deceleration

11.89 20.09 60.28 5739.50 6.18

12.32 20.09 67.25 6544.10 6.60

PRE-RISK MITIGATION

Tackett 72

5 4 1 ADEPT

4 7,8 1,2 2 Mid L/D

3 3 3 Ballute

2 6 5 4 Seperation

1 5 Aerocapture

1 2 3 4 5 6 Aerobraking

7 Controls

8 Trajecctory

Like

liho

od

Consequences

Risk Analysis

Goals to decrease risk 1. Further analysis of ADEPT increased success in redesign 2. Mid L/D eliminated because error was too high and increased IMLEO 3. Further analysis of ballute and parachute deployment increased success in redesign 4. New separation architectures can decrease separation risks 5. Aerocapture risk can be lowered by using lift to improve trajectory accuracy 6. Aerobraking will not be used because the time of flight is too long 7. Estimated failure rate is too high, more research required 8. Use of a semi-lifting body and an emergency ballute may increase success rate

Trajectory

CURRENT RISK ANALYSIS

Libben 73

5

4 2,3

3 7 1,4

2 5

1 6

1 2 3 4 5

Lik

elih

oo

d

Consequence

HuLa CarLa

Overall Risk 97.8% 98.1%

RiskProb. Of

Success

1 Aerocapture 99.70%

2 ADEPT 99.00%

3 HIAD 99.00%

4 Separation 99.50%

5 Controls 99.90%

6 Trajectory 100.00%

7 Ballute 99.70%

QUESTIONS?

TRAJECTORY ERROR REDUCTION ANALYSIS

Tackett 75

Atmospheric Error


Downrange Distance

(km)

Peak Decleration

(Earth g's)

Terminal Velocity

(m/s)

L/D = 0.0, atme = 100% 1062.60 2.92 90.818

L/D = 0.0, atme = 160% 910.28 3.52 70.24

L/D = 0.0, atme = 40% 1813.50 1.71 154.34

L/D = 0.2, Added Ballute, atme = 160% 2148.60 2.81 48.355

L/D = -0.2, Added Ballute, atme = 40% 803.58 6.05 103.06

Atmospheric Error


Downrange Distance

(km)

Peak Decleration

(Earth g's)

Terminal Velocity

(m/s)

L/D = 0.0, atme = 100% 1204.80 1.92 90.815

L/D = 0.0, atme = 160% 1110.70 1.91 70.24

L/D = 0.0, atme = 40% 1390.60 1.93 154.01

L/D = 0.2, Added Ballute, atme = 160% 1278.80 0.89 48.355

L/D = -0.2, Added Ballute, atme = 40% 1090.30 3.22 102.97

CarLa & XM3 mass ≈ 73 Mg β = 20 kg/m2

HuLa mass ≈ 23 Mg β = 20 kg/m2

DOWNRANGE ERROR MINIMIZATION

Tackett 76

Lift Optimization for Downrange Error Minimization

1. Employed a genetic algorithm optimizer from Dr. Crossley to optimize lift

2. Resulting downrange error is only 2.7 kilometers on the first landing and 0.8 meters on the 24th landing

3. Assumptions: 1. Atmospheric data is

known during flight 2. Bank angle variation can

minimize east\west error

4. Lift was varied between -0.2 and 0.2 according to Dr. Alan Cassell’s recommendation

ARCHITECTURE DESCENT EXAMPLE: HULA

Tackett 77


Sipich - 78

Background The fundamental design feature is the ballistic coefficient. From this we determine how big HIAD needs to be. HIADs are effective because they can significantly increase the surface area of a vehicle while adding minimal mass to the system as a whole. A non dimensional analysis was initially conducted for sizing purposes in order to reduce the mass of the system. From there, application of the HIAD in order to find an ideal ballistic coefficient for entry was conducted. Vehicles that will use the HIAD system: - Cargo Lander (CarLa) - XM3 Delivery - Return Option to Earth

HIAD SIZING

Sipich - 79

Torus Radius (rt)

Maximum Radius (rb)

HIAD System

Lander Vehicle HIAD sizing consists of two main dimensions, the overall maximum radius (rb) and the radius of each individual torus ring (rt). In this analysis, it is assumed all torus rings on the HIAD system will have the same radius. Each HIAD is to be composed of 8 torus rings which will mimic a 70o sphere-cone. When deployed, the system will be inflated with nitrogen gas (N2). Once entry with the HIAD system is complete, the apparatus will be discarded.

NON DIMENSIONAL ANALYSIS

Sipich - 80

The non-dimensional comparison used to assist in determining HIAD sizing. Ra is the maximum width of the payload. Rb is the desired maximum overall radius needed to achieve the target ballistic coefficient. Dt is the diameter of an individual torus ring. Do is the overall diameter of the system (2*rb). By using zetai, we can find the required zetat for a given number of torus rings. Thus giving the parameters Ra ,Rb, and Dt which define the overall HIAD system.

SRP SPECIFICATIONS

Dudaney 81

6 engines in a radial design[2][3]

• Redundancy

• Deeper throttling

• Control

• Increased mass/complexity

LOX/LCH4 [2]

• Minimal Boil-off

• More Insulation needed

Ispvacuum = 365 s [2]

• Reduces length of nozzle for landing

β = 150

Values for SRP stage:

Initiation Altitude = 11.14 km

Initiation Velocity = 670 m/s [1]

Final Altitude = 0.5 km

Final Velocity = 0 m/s

120 km

11 km / 670 m/s

0.5 km / 0 m/s

Aerocapture

Supersonic

Retropropulsion

Terminal

Descent

(30/60 s)

ADEPT HIAD

Surface

[1] Korzun, A.M. and Braun, R.D.

[2] Cianciolo et al.

[3] Edquist et al.

SRP SYSTEM MASS

Dudaney 82

mtotal = mpayload + mADEPT/HIAD + mpropellant,SRP + mengines[1] +mpropellant,terminal descent

Assumptions:

• No gimbaling/control or throttling

• Burn time = 45 s

• Leaves ~3% propellant mass

mSRP

msystem

msystem = 0.2549 * (mADEPT/HIAD + mpayload) + 462.4 (in kg)

*Values calculated simulating Equations of Motion

(See back up slides)

Human Lander Cargo

mpayload + mADEPT/HIAD

(Mg) 25 (assumed) 70 (assumed)

Msystem (Mg) 6.831 18.30

% of total mass 21.46 20.72

[1] Korzun, A.M. and Braun, R.D.

MID L/D

Dudaney 83

Pros: • Increases Entry

Corridor For Aerocapture (reducing risk)

• Reuse of Payload Fairing

Cons: • Adds Mass • Increases System

Complexity • Adds Volume

Notes: • Lower β with lower L/D may be

sufficient to lower risk • Another system would be required

for landing • Can be jettisoned after Aerocapture

Estimated Structure/Blanket Mass: 3.10 Mg HIAD Mass: 1.6 Mg Combined System: 4.7 Mg ADEPT: 2.76 Mg

MID L/D

Dudaney 84

Mid L/D Mass Components[1]:

• Structure

• Acoustic Blanket

• Separation System

• Avionics

• Flap

• TPS

Reasons for being looked at: • Mid L/D mass driven by volume • HIAD occupies small space when stowed • HIAD weighs less than ADEPT for HuLa • HIAD cannot be used for both entry and aerocapture

Supersonic

Landing

Terminal Descent

Entry/ Hypersonic

Aerocapture

Mass of Mid L/D Structure: 1.37 Mg Mass of Mid L/D Blanket : 1.74 Mg Mass of HIAD: 1.6 Mg Total: 4.70 Mg Mass of ADEPT Structure: 2.76 Mg

Would need more in-depth analysis

1 Cianciolo, A.M.D“Entry, Descent and Landing Systems Analysis Study: Phase 1 Report”, July 2010

MID L/D ARCHITECTURE

Dudaney 85

Entry Aerocapture

Aerocapture

Entry/Hypersonic

Supersonic

Terminal Descent

Landing

In the Aerocapture case, the Mid L/D would be jettisoned before entry; however in the entry case, the Mid L/D would be used for throughout the whole process, with a supersonic decelerator being used.

COMPARISON

Dudaney 86

Mid L/D Lifting Body (β = 30)

Analysis done with the same entry conditions. Mid L/D has a higher downrange range

MID L/D

Dudaney 87

Entry Conditions: • β = 87.5 [1] • L/D: -0.6 to 0.6[1] • Entry Velocity = 4.5 km/s • Mass = 25 Mg • Entry Angle = -9.1o

Notes (for a given entry trajectory): • Negative lift creates high G’s • Positive lift creates skip trajectory • Positive lift reduces max G’s

Lands Skips

ENTRY ANGLE AND LIFT

Dudaney 88

• Steeper entry results in higher G load • Shallower entry means more negative lift required (decreases rapidly)

Aerocapture Conditions (at 120km altitude): • β = 20 • Entry Velocity = 7.51 km/s (First human mission) • Exit Velocity = 4.5 km/s • No Atmospheric Error

*Vertical Line shows entry angle needed for zero lift (-8.343o)

TABULATED DATA

Dudaney 89

Entry Angle (degrees) -11.6 -11.4 -11.2 -11 -10.8 -10.6 -10.4 -10.2 -10 -9.8 -9.6

L/D 0.438 0.416 0.395 0.372 0.35 0.326 0.302 0.278 0.253 0.227 0.2

Max G load 6.2962 6.1189 5.931 5.7516 5.5604 5.3788 5.1918 4.9974 4.8055 4.6116 4.4168

Exit Angle (degrees) 9.9543 9.7276 9.5102 9.2807 9.0623 8.831 8.6038 8.3825 8.1584 7.9318 7.7036

Entry Angle (degrees) -9.4 -9.2 -9.1 -9 -8.8 -8.6 -8.4 -8.343 -8.2 -8 -7.8

L/D 0.171 0.143 0.1275 0.112 0.08 0.046 0.0105 0 -0.028 -0.069 -0.113

Max G load 4.2272 4.0215 3.9244 3.8246 3.6224 3.4208 3.2138 3.1544 3.0096 2.8025 2.5931

Exit Angle (degrees) 7.4603 7.2489 7.127 7.0107 6.7809 6.5453 6.3239 6.2626 6.0867 5.862 5.6621

Entry Angle (degrees) -7.6 -7.4 -7.2 -7 -6.8

L/D -0.162 -0.215 -0.275 -0.3418 -0.417

Max G load 2.3909 2.1862 1.9926 1.8042 1.6249

Exit Angle (degrees) 5.4209 5.2525 5.0018 4.7799 4.5772

REFERENCES

1 Hollis B.R. and Hollingsworth K.E, “Experimental Aeroheating Study of Mid-L/D

Entry Vehicle Geometries: NASA LaRC 20-Inch Mach 6 Air Tunnel Test 6966 ” NASA,

November 2014. 2 Mars Architecture Steering Group, “Human Exploration of Mars Design Reference

Architecture 5.0 ” NASA, July 2009.

Dudaney 90

HUMAN LANDER AND ROVER

3/12/15

Team Members: Kevin Lapp, Kyle Schwinn, Ted Danielson, Eiji Shibata, Jake Johnson, Cory Back, Rohan Dudaney, Zach Jochum, Daniel Ingegno, Kudzo Ahegbebu, Charlie Hartman, Cynthia Rose, Ben Libben, and Tony Sepkovich

OVERALL STRUCTURE

• Mission Overview and Human Factors

• Booster Vehicle

• Propulsion

• LEO to S1L1

• Cycler to Phobos

• Phobos to Mars

• Landers

• Rovers

• Cranes

• Summation

• Failure Analysis

• Conclusion

Lapp 92

MISSION OVERVIEW AND HUMAN FACTORS

• Landers group together in Leo to meet

Cycler

• Landers leave Cycler and head to Mars

• 2 Landers to Phobos

• 1 Lander to Mars

• 1.3 day trip max

• 6 people total (All to Phobos)

• Two years later astronauts travel to Mars

• Same Lander from Phobos must travel to Mars.

• Landers on Mars transform into rovers for reusability

• Rovers must reach 20 km/h and be able to climb a 30º degree incline.

• Rover will transport astronauts to the XM3s.

• Rovers will also be used for exploratory missions.

Lapp 93

Item Mass (Mg)

Volume (m3)

Water 0.421 0.004

Nitrogen 0.005 5.77e-6

Food 0.146 0.018

Miscellaneous Items

0.972 12.22

Total 0.991 12.25

BOOSTER VEHICLE

Lapp 94

• Will be used to get the three Landers to

the Cycler

• Three Landers attach to top of

vehicle around the

circumference

• Vehicle will contain propellant tanks and

food for 180 days. (30 day buffer)

• Vehicle will attach to an XM3 for transit

to Phobos/Mars

• Volume

• 856.8 m3 for tanks

• 8.911 m3 for food

• Total = 865.71 m3

• Dimensions

• Diameter = 8.2 m

• Height = 16.4 m

LANDER RENDEZVOUS PROPULSION

Back - 95

Total Lander Mass LEO 368.6 Mg

Lander LEO Mass

LEO Payload Mass 115.9 Mg

Chemical Impulsive Burn - Boost Vehicle

Parameter LH2/LOX

Propellant Mass (Mg) 224.7

Inert Mass (Mg) 28.0

Fuel Tank Volume (m3) 703.6

Oxidizer Tank Volume (m3) 153.2

Thrust: 1739 kN Burn Time: 9.5 min

Δ𝑉 = 4.21𝑘𝑚

𝑠

S1L1 Boost Vehicle

Specific Impulse (s) 450

Throat Diameter (m) 0.26

Exit Diameter (m) 1.64

Propellant Flow Rate (kg/s) 393.7

Lander LEO Mass Breakdown

Human Factors Inert (Mg) 7.1

Consumables (Mg) (3) 1.0

Structure (Mg) (3) 5.22

Propulsion Phobos-Mars (Mg) (2) 23.03

Propulsion Mars Direct (Mg) (1) 6.89

Controls (Mg) (3) 1.33

Power and Thermal (Mg) (3) 2.58

Aerodynamics (Mg) (3) 8.39

Communications (Mg) (3) 0.08

Total Mass 115.9 Mg

LEAVING CYCLER PROPULSION

Ingegno 96

Illustration based on figure 6-9 from Rocket Propulsion Elements by Sutton and Biblarz.

Cycler to Mars Propulsion Engine Parameters

Fuel RP-1

Oxidizer LOX

Thrust (KN) 865.6 KN

ΔV One Stop

1.695 km/s

ΔV No Stop

0.257 km/s

One stop: Representing stop at Phobos before Mars No Stop: Representing Direct to Mars

PROPULSION REQUIREMENTS

Ingegno 97

Figure 1: Engine Chamber, Throat, Nozzle and Tank Proportional Model. The length from nozzle exit to the top tank is about 7.8 meters

One Stop No Stop

Total Vehicle Mass

40.92 (Mg) 22.30 (Mg)

RP-1 Tank

Volume

7.42 m3 0.88m3

LOX Tank

Volume

11.06 m3 1.31 m3

Ahegbebu 98

Propellant Mass Breakdown

Total Propellant Mass (Hover + Delta V kill)

5.12 (Mg)

Propellant for Entry Delta V kill

1.33 (Mg)

Hover Maneuver Alone 3.74 (Mg)

Structural Mass : 17.9 (Mg) Total Mass (with Prop): 22.9 (Mg) Propellant Volume : 4.57 (m3)

Fuel Mixture: N204/MMH Incoming Vehicle Velocity : 100 m/s Required Hover: 60 s

HOVER MANEUVER

HUMAN LANDER

Vehicle Mass [Mg] Volume [m^3]

Lander 5.224 29.42

Specification Measure

Bottom diameter 5.842 meters

Top diameter 2.641

Height 1.895 meters

Wall materials Al7075-T6

Wall Thickness 1.59 cm

Inclination angle 32.5 degrees

Stringers are 0.00425 meters in width and spaced 0.176 meters apart.

Jochum - 99

Shibata - 100

Insulator

Structure

Ablator

radq

genq

Strain Isolator Pad (SIP)

Adhesive

aeroq

Layer Thickness [cm] Mass [Mg]

SLA-561 1.0 0.318

Aerogel 2.54 0.538

RTV-560 0.0203 0.0406

Nomex 0.1372 0.0168

THERMAL PROTECTION SYSTEM FOR LANDER

MARS ROVER

Schwinn - 101

Rover Mass

Chassis 1.05 Mg

Capsule 9.85 Mg

Total 10.84 Mg

Capsule Mass

Human Factors 0.99 Mg

Structures 4.79 Mg

Controls 0.02Mg

Power 3.26 Mg

Insulation 0.71 Mg

Communications 0.08 Mg

Total 9.85 Mg

ROVER SYSTEMS • Rover Instruments

• ChemCam, REMS, APXS, MB, CheMin, SAM, MAHLI, Mini-Tes, MEDA,

SHERLOC, RAD, DAN

• Individual use does not exceed 700 W

• Total mass of instruments will not exceed 0.5 Mg

• Rover Controls

• There are navigational cameras and instruments

• Fore and Aft hazard cameras for hazard avoidance during remote operation

• Two 5 DoF arms with graspers and instruments

Danielson 102

Component Power Usage Mass Volume

Robotic Arm 70 W 5 kg 59500 cm^3

Hazard Camera 2.14 W 220 g 188.5 cm^3

Navigation Camera

2.14 W 220 g 188.5 cm^3

Star Tracker 8 W 5.48 kg 499.2 cm^3

IMU 12 W 748 g 529 cm^3

Sun Sensor 28 W 1.3 kg 108.3 cm^3

CRANE • Mass: 9 Mg

• Power: 100 kW

• Volume: 226 m3

Danielson 103

Component Mass

Chassis[1] 4 Mg

Power Supply 1.5 Mg

Crane System 3.5 Mg

Condition Power Usage

Driving at 10 km/hr 78.3 kW

Driving with XM-3, .18 km/hr 8.55 kW

Lifting XM-3 onto flatbed 2 kW

Component Volume

Chassis 226 m3

Crane System Height 15 m Crane Chassis based on the JPL ATHLETE Rover [1]

References:

1. “ATHLETE (All-Terrain, Hex-Limbed, Extra-Terrestrial Explorer),” NASA, Washington, DC

PICTURE HERE!

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TOTAL SUMMATION

Lapp 104

Stages Mass (Mg) Volume (m3)

Hover – Touchdown 11.11 0.746

Entry – Hover 23.15 5.106

Phobos – Mars Entry 29.81 19.06

Aerocapture - Phobos 38.24 19.21

Cycler – Aerocapture 42.05 31.30

Leo – Cycler 277.8 865.7

Note: Leo – Cycler is for three landers

Lapp 105

FAILURE ANALYSIS Component Risk Type L C Mitigation

Propulsion Delta V Error Technical 4 1 Added fuel for additional burns

Crane Leg issues Technical 3 1 Replace leg

Crane Tether issues Technical 3 1 Replace tether

Crane Wheel issues Technical 4 1 Replace wheel

Rover Chassis Wheel Issues Technical 4 1 Replace wheel

Rover Chassis Overall destruction

Technical 2 2 Can be down a rover chassis.

Rover/Lander Power

Explosion, depressurization, loss of life support systems

Technical 1 4 Replace with a minor or not used power source, redundancy

QUESTIONS?

DESIGN DECISIONS • CAPSULE DIMENSIONS

• CARGO RESUPPLY MISSION PARAMETERS

• HIAD ENTRY TECHNOLOGY

• ENTRY CONTROL AND GUIDANCE

• FARMING IN THE CARGO SHELL

• PROPULSION COMPARISON AND CHOICE

• RISK ASSESSMENT

CARGO LANDER

MARCH 12, 2015

DRAKE WISSER, ANTHONY MILLER, BEN TACKETT, DJ LEE, ZAK SIPICH, KEVIN LAPP,

ALEX MANGUIERA, KEVIN LAPP, CHARLIE HARTMAN, EIJI SHIBATA

CARLA CAPSULE

Wisser, Drake - 108

Height: 6 meters

Outer Radius: 3.8 meters

Wall Thickness: 0.02 meters

Leg Length: 3 hydraulic legs at 2 meters

Structure Volume: 6.449 m3

Empty Volume: 265.7 m3

Once Landed and emptied of contents, top two floors

repurposed for farming and bottom for support

systems.

Top Floors: 2 meters in height

44.87 m2

Bottom Floor: 1.92 meters in height

44.87 m2

2 m

2 m

1.92 m

PAYLOAD WALL TPS (CARLA)

Shibata - 109

Insulator

Structure

Ablator

radq

genq

Strain Isolator Pad (SIP)

Adhesive

aeroq

Layer Thickness [cm] Mass [Mg]

SLA-561 1.5 0.7805

Aerogel 1.09 0.3783

RTV-560 0.0203 0.0665

Nomex 0.1372 0.0276

Total 2.764 1.253

CARGO RESUPPLY MISSIONS

Wisser, Drake - 110

Assuming an entry mass similar to that of the XM3 delivery mission, which is 25 Mg.

Since the CarLa is using Hohmann Transfers, it is limited to sending resupply missions

every 2 1/7 years ( 782 days).

Therefore the best viable option for CarLa resupply missions is to send 3 vehicles during

the Hohmann transfer window. This allows to even cut back on Human Factor supplies for

other mission needs.

Cargo Mass (Mg)

Number of crew to support

Frequency of missions (days)

Number of CarLas to send

40% of farming present

21.35 18 365 1 Yes

24.83 6 1275 3 No


Sipich - 111

Maximum Radius (rb)

Torus Radius (rt)

Final CarLa HIAD Sizing Total Mass of CarLa = 70.39 Mg




Max Radius (rb) = 24 m


CarLa Lander

HIAD

Mission Requirement : Safely decelerate CarLa into the Martian atmosphere for direct entry

Critical Assumptions: Total mass of HIAD includes surface material, inflation gas (N2), and inflation system Flexible TPS materials will properly withstand any heat loading/heat flux it may encounter Before being deployed, the HIAD will be packed tightly around CarLa vehicle

ENTRY CONTROL AND GUIDANCE Prebank

• Initial bank angle at entry, eject

first ballast

Hartman-112

Range Control

• Control bank angle to minimize

downrange error

Heading Alignment

• Control bank angle to minimize

crossrange error

Terminal Descent • Prior to engine ignition,

eject second ballast to re-center C.M

Bank

Prebank

Bank

Propellant Mass [Mg]

4.30

ENTRY CONTROL AND GUIDANCE

Hartman-113

1. Eject Ballast Mass

1. Eject Ballast Mass

-α A.C.

2. C.M. re-centered

Pre-Terminal Descent Re-center C.M.

2nd Ballast Mass [Mg]

5.5

Prebank Phase Generating Lift

1st Ballast Mass [Mg]

3.25

α A.C.

2. C.M. Moves

HIAD Still Packaged

FARMING IN THE CARGO SHELL

K. Lapp 114

• Dimensions

• 3.78 m inner diameter

• 6.0 m height

• 2 Floors

• 29.40 m2 farming per level

• 58.80 m2 total farming area per cargo

• Using aeroponics

• Power Requirements Per Cargo

• Thermal Control = 1.212 kW

• General Air = 2.000 kW

• Lighting = 6.65 kW

• General Water = 3.34 kW

• Total Power = 13.20 kW

• Goal of 40% Sustainable for a crew of 18

• Need 10 CarLas

• Rest will be sent

• 7.396 Mg

• 9.299 m3

Cargo numbers obtained from Drake Wisser, Farming results obtained from Andrew Bokhart

Farming Results per CarLa

% Calories

Yield (Mg)

3.41

0.38

Potatoes

Soy Beans

Pinto Beans

Wheat

CAPSULE FARMING

- The ten CarLas would surround the central trigonal base with enough room in

between the ring and center for the rovers to be able to get through.

Wisser, Drake - 115

XM3 and connector design by Jani Dominguez

PROPULSION CHOICES

Miller 116

Considerations • Fuel Type

• Xenon vs LOX/LH2 • Possibility of In-Situ Propellant

Production for LOX/LH2

• Power Requirements • Required power for Electric

Propulsion to move this large of mass in time frame required is unrealistically large

• Cost of One-Way Mission • Costs for Power Supply is high

for multiple one way missions

Best Case Electric

Propulsion VASIMR

Payload Mass (Mg)

31 50

Power Required (kW)

100 200

Propellant Mass (Mg)

10.49 3.9

Propellant Volume (m^3)

3.56 1.3576

Inert Mass (Mg)

2.9 2.9

Total Mass (Mg)

44.39 56.8

Time of Flight (Days)

1532 1598

PROPULSION FUEL COMPARISON

Miller 117

Lunar Fuel Production

LH2

• Pros: Efficient Fuel

• Cons: Boil Off Issues

CH4

• Pros: No Boil Off/Current

Tech

• Cons: Moon is Carbon

light

SiH4

• Pros: Great for Lunar

production

• Cons: TRL too low

LOX/LH2 LOX/CH4 LOX/SiH4

Payload Mass (Mg)

76.97 76.97 76.97

Delta V Required

(Km/s) 3.6 3.6 3.6

Oxidizer Mass (Mg)

91.8 130.2 162.9

Oxidizer Volume (m^3)

80.45 114.1 142.9

Fuel Mass (Mg)

24.6 34.9 43.7

Fuel Volume (m^3)

347.4 82.6 32.6

Total Vehicle Mass (Mg)

218.4 267.1 308.7

REQUIREMENTS & SPECIFICATIONS

Lee - 118

Optimized Nozzle Properties

Resulted Values

Exit Area (for each) 0.817 𝑚2

Throat Area (for each) 0.014 𝑚2

O/F mixture ratio 2.25

• N2O4/MMH Used

• Thrust Required = 247 kN

Mission Requirements Given Values

Duration of Hover 30 sec

Horizontal Landing Error 6 km (with 99% probability)

Entering Velocity to kill (Δ𝑉)

120 m/s

6 Nozzles instead of a single nozzle - For redundancy - One nozzle fails We have extra

MASS & VOLUME UPDATES

Lee -119

Mass Breakdown Values (Mg)

XM3 54.39

Propellant for terminal descent 4.66

Propellant for hover 5.63

Total Propellant 10.27

Total Mass 64.66

Volume Values (𝒎𝟑)

Propellant Tank 10.2

ENGINE CYCLE CHOICE

120

Gas Generator Pros/Cons

• Pros

• Simple

• Wide F range

• Cons

• Possible loss in performance

• Gives low Isp

Mass flow rates Values (kg/s)

N2O4 129.8

MMH 57.69

Volume flow rates Values (m3/s)

N2O4 0.0900

MMH 0.0656

head P rise Values (kPa)

N2O4 16.77

MMH 12.21

ABLATIVE COOLING CHOICE

121

Ablative Cooling:

• Silica fiber in phenolic

resin

• Single-use application

• Low-cost

Risk:

• Performance Decrease

Temperature Values (K)

Chamber Temperature 3374.67

Throat Temperature 3199.67

Exit Temperature 1358.02

RISK ASSESSMENT

Wisser, Drake - 122

1. EDL propellant chemicals

degrade the inner wall of the

fuel chamber.

2. An EDL nozzle stops

functioning.

3. Hull is ruptured by space

debris.

4. Main engine failure.

5. Control thruster failure.

1

2

3

Risk Mitigation

1 No need for horizontal propulsion, EDL hover shouldn’t be effected

2 6 nozzles creates redundancy

3 Micrometeoroid shielding prevents this

4 Component redundancy

5 Redundant control thrusters

4,5

QUESTIONS?

HIAD FOR XM3 DELIVERY

Sipich - 124

Final XM3 HIAD Delivery Sizing Total Mass of XM3 = 73.14 Mg






The HIAD system for the delivery of the XM3 was designed under similar conditions as the CarLa. The main driving difference was the total mass of the system, which was

larger in the XM3 resulting in a larger HIAD for the same ballistic coefficient

Backup

Lee - 125

CYCLER VEHICLE

3/12/15

ARJUN JAYARAJ, ALEC MUDEK, CORY BACK, SAM FERDON, SAPHAL ADHIKARI, ALEX DAVIS, MAXIME PINCHAUD, NATHAN HOUTZ, JOCELINO RODRIGUEZ, AND JULIAN WANG

CYCLER FINAL DESIGN

127

DIMENSIONS

128

Cycler General Dimensions (Optimized

using Stress Analyses)

• Hull Thickness = 25 mm

• Net Volume = 59.72 m3

• Length = 30 m

• Radius = 4 m

Net Mass of Cycler IMLEO = 521.7 Mg

Mass at S1L1 = 225.1 Mg

CYCLER HUMAN FACTORS

• Prepared for 180 days in cycler

• Using Environmental Control and Life Support System (ECLSS) based on the ISS

• Assume system provides 86% water recovery

Crew of 18

Total Mass (Mg) 32.746

Total Volume (m^3) 788.911

Power Required (Kw) 147.3

Human Factors Totals:

Note: these numbers are not total mass and power requirements for the cycler vehicle.

Mass (Mg) Volume (m^3) Power (kW)

Food1 7.09 8.91 NA

Crew Quarters2 (x3) 3.11 248.4 19.77

Water Systems2 (x3) 3.75 3 NA

Life Support2 (x3) 1.69 8.6 29.4

1. Food mass and volume calculated by Matlab script: Master_HumanFactors.m 2. Crew quarters, water systems, and life support values taken from mass , power,

volume spreadsheet for XM3 –cycler variant

CYCLER’S SOLAR PANELS

Maxime PINCHAUD - 130

(1) – ATK communication, « MegaFlex — Leverages UltraFlex Flight and Production Heritage », ATK-Goleta Paper, June 2014

Cycler without crew

Cycler A B

Power requirement 221.4 kW 271.4 kW

Configuration 2 x 24.5 m 2 x 27 m

Area 833.7 m2 1022 m2

Weight (cells) 1.03 Mg 1.27 Mg

Stowage volume 5.41 m3 6.66 m3

24.5 m diameter

Cycler A

Height (m) 8.17 9

Thickness (m) 0.28 0.31

Stowed dimensions of one

solar panel

(1)

Height =

R*2/3=8.17 m

Width = 3.2 m

CYCLER ESTABLISHMENT MASS REDUCTION • LH2/LOX Boost Vehicle

• Chemical Burn out of LEO

• Electric Propulsion into S1L1

Back - 131

Cycler LEO Mass Breakdown

Cycler Mass 192.5 Mg

EP Inert Mass 10 Mg

EP Propellant 21.90 Mg

Payload Mass 224.4 Mg

Impulsive Chemical Burn – Depart LEO Δ𝑉 = 3.175 𝑘𝑚/𝑠

Chemical Impulsive Burn - Boost Vehicle

Parameter LH2/LOX

Propellant Mass (Mg) 267.6

Inert Mass (Mg) 29.7

Fuel Tank Volume (m3) 837.7

Oxidizer Tank Volume (m3) 182.4

Electric Propulsion – Into S1L1 Spiral Out: TOF = 800 Days Power: 250 kW

*Worked closely with Alex Davis – Mission Design

Total Cycler Mass LEO 521.7 Mg

S1L1 Boost Vehicle

Specific Impulse (s) 450

Throat Diameter (m) 0.26

Exit Diameter (m) 1.64

Propellant Flow Rate (kg/s) 393.7

Chemical Thrust: 1739 kN Burn Time: 11.3 min

Ma

TRAJECTORY CORRECTION MANEUVER (TCM)

Adhikari-132

Electric Propulsion - Hall Thruster Isp Total

Efficiency Electric

Efficiency Propellant Propellant

Mass Operation Time (OP)

3000 s 0.70 0.85 Xenon 0.655 Mg 1.067 yr

Thrust Required • Half Period (OP) = 0.5728 N

Power Required • Half Period (OP) = 12.04 kW

Optimized Current and Voltage • Current = 13 A • Voltage = 790 V

Fig 2: Current and Voltage Optimization

Advantages of Electric over Chemical Propulsion • Propellant mass reduced by around 82% • Propellant volume reduced by around 94% • Inert mass reduced by around 90%

Vehicle Name Mass 𝑴𝒈 Power(𝒌𝒘) V𝐨𝐥𝐮𝐦𝐞(𝒎𝟐) Rejection Rate 𝒌𝒘

Cycler 6.8 13.67 382.2 126.1

CYCLER THERMAL CONTROL

Related Subsystem Function

Louvres Adjust rejection rate

Mechanical Arm Adjust radiation area

By-Pass Valve Adjust radiation level

Reaction Wheel Adjust Orientation of Cycler

Thermal Load of Cycler with Different Pitch Angle 𝛼 Radiator Schematic

Specification of the Active Thermal Control System for Cycler

Rodrigues, Jocelino 134

LOUVER ANALYSIS

Conclusion: Louver system helps reduce temperature range and does not add

much mass, but it is not enough.

Configuration A (Closed) Configuration B (Open)

Louver mass (Mg)

Total cycler mass (Mg)

% of total mass

5.6 220 3.8

Total surface area = 1428.6 m2 (Structures team)

Design: Weight/area = 3.2 kg/m2

CONTROL SYSTEM SELECTION

Mudek - 135

Astrionics Mass (Mg) Power (W) Volume (m^3)

Inertial

Measurement Unit 1.88e-2 64 3.94e-2

Sun Sensor 2.5e-5 3.6e-2 1.08e-5

Star Tracker 5.49e-3 9 6.42e-5

Total 2.43e-2 73.0 3.95e-2

Control Scheme Total Mass (Mg) Total Power (W) Total Volume (m^3)

Reaction Control Wheels 7.8e-3 52 9.40e-3

Electric Prop Thrusters 5.14e-6+Mdry .96 8.94e-3

Control Moment Gyros 2.80e-2 11/113 2.92e-2

Conclusions • Standard thrusters are not desirable because they provide too much thrust and are not reliable • Control wheels are also not very reliable, but with proper redundancy they have a lower resource cost and are less massive solution • Electric propulsion would require very little fuel and uses the least resources, but does not allow for as accurate control and still requires fuel • Control moment gyroscopes provide a more robust control system than we need, but are more reliable than reaction control wheels • Recommend the use of both a control moment gyroscope and control wheels for redundancy, accuracy, and lifespan

= designed for 3σ performance

RISK ASSESSMENT

Jayaraj - 136

1) Micrometeorite Puncture

2) Failure of Structural components

3) Life support component failure

4) Cycler Power Failure

1 2

3

4

QUESTIONS?

138

Risk Potential Outcomes Mitigation Plan

Run out of food/water Crew cannot operate at full capacity. Worst case: cannot survive entire flight

Carry extra food/water, planned diet has more calories than are required for basic survival.

Oxygen production by electrolysis fails

Without oxygen, crew will not survive the flight

-Carry three life support systems(1 for each hab), if one fails the other two can maintain environment. -Carry reserve of liquid oxygen (30 days)

Cabin fire Destruction of equipment, possible loss of life.

-Fire detection and suppression systems included on each module. -Triple redundancy on equipment

Radiation Exposure -Numerous negative health effects. Including increased cancer risk, and tissue degeneration. -Large solar flares could emit deadly levels of radiation.

-There is currently no plan to shield the cycler crew from radiation, as per the mission requirements. -A lot of research is currently being done on this subject as little is truly know about the long-term effects.

Zero-G Environment Lack of gravity causes bones and muscles to weaken over time. This may impede astronauts ability to carry out the mission.

Astronauts will attempt to mitigate the effects by exercising and spending time in a small centrifuge that creates artificial gravity.

HUMAN FACTORS RISK ASSESSMENT

CYCLER’S SOLAR PANELS


(1) – ATK communication, « MegaFlex — Leverages UltraFlex Flight and Production Heritage », ATK-Goleta Paper, June 2014

(1)

Height =

R*2/3=6.7 m

R=10 m

Width = 3.2 m

• Lightweight

• Reliable

• Low stowage volume

• TRL: 5

BACK UP SLIDE


Cycler’s power requirement: Human factors: XM3-C: 14.5 kW (with crew) XM3-C: 2.185 kW (without crew) Communication XM3-C: 7.96 kW (With crew onboard) 24.8 kW (without crew) Propulsion: Cycler B: 190 kW (800 first days) Cycler B: 240 kW (800 first days) Cycler A and B: 6 kW (during cycles) Then Cycler’s total power requirement is: With crew onboard: 3*14.5+4.4+3.18= 51 kW Without crew onboard: 3*2.185+20.5+3.18= 30 kW

Cycler without crew (30 kW)

Configuration 1.6 x 20 m

Weight 0.334 Mg

Stowage volume 0.75 m3

Rodrigues, Jocelino 141

ADDITIONAL LOUVRE DESIGNS

TABULATED POWER INPUT RESULTS BASED ON SMALL SELECTION OF MATERIALS

Summary of louver designs [reference 8, pp 333]

OSC chosen for presentation as an example due to it’s consistency in flight history (see

table below).

OSC Swales Starsys

Blades 3 to 42 - 1 to 16

Area (m2) 0.07 to 0.6 0.08 to 0.5 0.02 to 0.2

Weight/area (kg/m3) 3.2 to 5.4 ~ 4.5 5.2 to 11.6

Flight history

Niambus, Landsat, Viking, SolarMax,

SPARTAN, EUVE, MGS, MSP

XTE, Stardust Rosetta, Mariner,

Voyager, MLS, Cassini

COMMUNICATIONS RISK ASSESSMENT

Houtz, 142

5

4

3 Before

2

1 After

1 2 3 4 5

Like

liho

od

Consequences

*Before and after the inclusion of triple redundancy. Redundant units are two low-gain antennas (incapable of high data rates, but good enough for telemetry)

COMMUNICATIONS RISK ASSESSMENT

(NOTES ONLY)

143

•Justification for before: deployable mesh antennas have not been on too many missions, hence higher risk. Failure of the HGA would pose no immediate threat to the crew but would make it difficult (though not impossible) to have the vehicle accurately perform maneuvers. If the crew is aboard the vehicle when the failure occurs, it may be possible to repair.

•Justification for after: Likelihood of failure is virtually zero – LGA’s are well tested, easy, and reliable. If the HGA fails, leaving only LGA’s, the data rate drops significantly, which could have small consequences for the crew. The rate of transmitted medical information, as an example, may need to be cut, but neither the vehicle nor the crew will be lost in this case.

DANIEL GOLDBERG, MAX DZIS, TRISTAN LOUDEN, LEE WESTROPP, ALIBEK YERTAY, YUE GUO, MORGAN LUCAS, NICOLE VAUGHN, JOCELINO RODRIGUES

MARS RETURN OPTION

12 MAR 2015

OVERVIEW/REQUIREMENTS

Goldberg 145

Immediate Return Date

22 January 2038

Later Return Windows

Aug 06 – Sep 15 2039

Sep 22 – Sep 30 2041

MAM- Mars Ascent Module

BA 330- Bigelow Aerospace Inflatable Hab

HuLa- Human Lander

Later Returns

Immediate and Later Returns

HUMAN LANDER

Goldberg 146

Human Lander

Dry Mass [Mg] 19.5

EDL Mass [Mg] 2.1

Propellant Mass [Mg] 5.5

Available ΔV [km/s] 0.3067

Image courtesy of Ted Danielson

BA 330

Dzis 147

Total Volume 483.0 m3

Pressurized Volume 330.0 m3

Avg. Wall Thickness 0.46 m

Consumables 7.03 Mg

BA 330 Mass 20.0 Mg

Total Mass: 27.03 Mg

CONSUMABLES

Mass [Mg] Volume [m^3] Tank Number

H2O 1.6438 1.6438 11

Food 2.6325 3.3098 N/A

O2 0.4476 0.1294 4

N2 0.0774 0.0031 1

Item 2.5990 324.3528 N/A

Total 7.0253 329.4389 N/A

• H20: Water used in oxygen generation, general consumption, etc.

• Food: Food required to keep crew sated

• O2: Reserve of liquid oxygen to be used in case of failure of the oxygen generation system (includes tanks values)

• N2: Reserve of liquid nitrogen to replace that which is lost due to leaks, airlock, etc. (includes tanks values)

• Item: Miscellaneous consumables and fixtures required

Lucas 148

RETURN VEHICLE POWER & THERMAL CONTROL

Radiator

EDL

Westropp 149

Radiator

Mass (Mg) 1.16

Power Required (kW) 2.33

Area (m2) 65.2

EDL

TPS Mass 1.542 Mg

HIAD Material Carbon Composite

MAM: Lithium-Ion Batteries

Total Energy (MJ) 70.57


Total Volume (m3) 0.014

BA-330: 2 Circular ATK Panels

Total Power (kW) 13.96


Stowed Volume (m3) 0.349

Deployed Area (m2) 173

Diameter (m) 10.5

ATTITUDE CONTROL SYSTEM

Yertay 150

Actuators

Reaction Control System

Number 32

Total Mass [Mg] 0.545

Total Thrust [N] 431

Isp [s] 280

Reaction Wheel

Number 5


Power [W] 1.7

Total Volume [m^3] 0.25

Sensors

Star Sensors

Number 2


Power [W] 20

Sun Sensors

Number 2


Power [W] 3

• Fuel cost included in RCS mass • Additional sensors and actuators for

redundancy • 0.001 deg accuracy on sensors

Total Mass = 0.601 Mg

PHOBOS LAUNCH VEHICLE

Goldberg 151

Performance Parameters

Isp 330s

Thrust 1.44MN

Chamber Pressure 4.29Mpa

Chamber Temperature 3284K

Optimized Values1,2,3

Payload 48.6 Mg

ΔV 1.7242 km/s

Flight Duration 204 days

GLOW 100.9 Mg

Model by Tristan Louden

Rocket Sizing

Fuel Volume 17.118m3

Oxidizer Volume 20.895m3

Rocket Length 8.5m

Rocket Diameter 4m

MAM

Dzis 152


Pressurized Volume ~23.54 m3

Avg. wall thickness 2.54 cm

Crew Mass ~0.302 Mg

MAM Mass 2.768 Mg

Propulsion Mass 34.76 Mg

Total Mass: 37.83 Mg

Vehicle Requirements

Crew: 3

Internal Volume: ~20 m3

BA 330 docking: 4 reaction wheels and RCS thrusters

Ascent to Phobos: Two stage ascent

153

MAM ATTITUDE CONTROL SYSTEM The MAM control system consists of 16 RCS thrusters and 4 reaction wheels.

Guo 153

RCS Thrusters

Number Thrust each [N] Isp [s] Total mass [Mg]

16 360 280 0.519

Reaction Wheel

Number Total mass [Mg] Total power[W] Total volume [m^3]

4 0.042 96.098 0.0512

• Fuel cost included in RCS thrusters’ total mass

• The propellant mass of RCS thrusters is 6.876 kg (0.006876 Mg)

• Total control system mass is 0.561 Mg

• The RCS thrusters are in charge of adjusting the MAM to the desired orientation during orbit insertion into LMO and landing

• The reaction wheels will counteract the gravitational and environmental torques in orbit

MARS ASCENT CONCEPT

Louden 154

●Total ΔV - 6.7486 km/sec

○Includes transfer, losses and ability to

orbit at 200 km

○Ascend to Low Mars Orbit (LMO) at

200 km above surface

○Initiate a Hohmann Transfer to head

towards Phobos orbit

○Burn at apoapsis to provide an

inclination change for final burn at

to maintain a Phobos orbit.

**Orbital Path courtesy of Nicole Vaughn

Transfer Point ΔV

LMO 3.451

Hohmann 1.244

Inclination Change 1.567

● SaTA

FINAL MARS ASCENT OVERVIEW

Louden 155

Propellant Mass [Mg] Total Mass [Mg] Total Fuel Vol. [m3]

Stage 1 23.8585 27.5585 23.8579

Stage 2 6.3039 10.2719 7.4789

Total Vehicle 30.1624 37.830 31.3368

• Propellant type: 1st & 2nd CH4/O2

• Vehicle Characteristics

Pc [Mpa] Tc[ K] Burn [sec] Thrust [kN]

Stage 1 5.171 3496.19 171.38 459.5

Stage 2 4.1368 3523.74 197.11 110.8

• Drop 1st stage after 171 seconds - in LMO

• 2nd burn not continuous; burn at apoapsis, Hohmann transfer burn

Total Est. Height [m] Vehicle Diameter [m]

Stage 1 19.43 3.50

Stage 2 4.84 4.70

Total Vehicle (St. 1,2 shell, inter-stage) 24.27 N/A

• Overall Dimensions

MARS ISRU: CH4& LOX PRODUCTION SYSTEM

MARS – MASS AND POWER VALUES

• Sent as cargo inside first CarLa

• Goal: have propellant needed for Return Vehicle ready by the time the first manned

lander arrives on Mars (~ 2 years)

• Inputs CO2 (atmosphere) and H2 (electrolysis of water obtained from regolith*)

IMISPPS (Integrated Mars In Situ Propellant Production System) – Zubrin [2]

Power / unit (kW) Mass / unit

(Mg) # of

Units Total power

(kW) Total mass

(Mg)

0.9 0.12 63 56.7 7.12

1x IMISPPS unit produces 1 kg/day of propellant combination

Hence, 41 units needed in total x 1.5 (safety factor)

* See Julian’s work on water extraction from regolith (Power & Thermal Controls team)

MARS ISRU: CH4& LOX PRODUCTION SYSTEM

MARS – MASS SAVINGS

Conclusion: CH4/LOx yields significant mass savings for IMLEO.

MARS RETURN VEHICLE*

No ISRU UMDH/N2O4 + solid

ISRU CH4/Lox

∆

Propellant Mass (Mg)

51.4 37.8 - 47.7 %

Tank Volume (m3) 33.2 35.8 + 7.5 %

IMMARS (Mg) 59.1 45.7 - 22.7 %

CarLa IMLEO (Mg) 267.1*** 222.97 - 16.5 %

* Values obtained from Propulsion team ** Integrated Mars In Situ Propellant Production System

*** Would require 2 CarLa trips

0

10

20

30

40

50

60

70

No ISRU ISRUM

ass

(Mg)

MAM Vehicle as CarLa payload

∆ = 44.1 Mg

MAM DESGIN

Louden 158

MAM SHELL DESGIN

Force Magnitude

(MPa)

Maximum Stress -Von

Mises (MPa)

Maximum Stress – Tresca

(MPa)

Max Deflection

(mm)

Drag 0.0118 1.661 1.936 .0532

*Al 7075-T6 Ultimate Tensile Strength = 503 MPa → Well within safety limits


Min. Wall Thickness 1.50 cm

Stringer Thickness 2.50 cm

Outer Area 66.31 m2

Density 2810 kg/m3

Mass 1.916 Mg

Dzis 159

MAM STRESS ANALYSIS

Dzis 160

Force Magnitude

(MPa) Maximum Stress -Von Mises (MPa)

Maximum Stress – Tresca (MPa)

Max deflection (mm)

Gravity+Thrust 10272+50529 122.8 137.5 3.308

*Al 7075-T6 Ultimate Tensile Strength = 503 MPa → Factor of safety is 4.12 at point of max. stress

*Next highest stress range is 51.73-8.377 MPa through majority of structure Factor of safety between 10-70 at min. stress

RISK ANALYSIS

Goldberg 161

5

4

3 2,3,4 5,6

2 1

1

1 2 3 4 5

5

4

3

2 4 6 5

1 2,3 1

1 2 3 4 5

Lik

elih

oo

d

Consequences

Lik

elih

oo

d

Consequences

Risk Mitigation

1 MAM Drag Buckling Add fillets, increase floor thickness

2 Communications failure Add two low-gain antennas

3 Sensor/Actuator Failure Redundant sensors/actuators

4 Insufficient Propellant Generation Additional propellant generation units

5 Engine Failure Extensive testing and analysis

6 Insufficient ΔV burn Assuming additional engine inefficiencies Pre-mitigation Post-mitigation

QUESTIONS?

CONTINUOUS COMMUNICATION SYSTEM • DEPLOYABLE MESH ANTENNA

• EARTH TO MARS

• CYCLER TO EARTH AND MARS

• SATELLITES IN MARS ORBIT

• ADEPT ANTENNA

• SHORT RANGE COMMUNICATIONS

.

TONY SEPKOVICH, NATHAN HOUTZ,

QIRONG LIN, ALEX MANGUEIRA

COMMUNICATIONS

3/12/15

OVERALL COMMUNICATION REQUIREMENTS

• 12 Mbps signal for HD video streaming

• Provide necessary communications link for sending control signals:

• Cycler

• Crane

• Rovers

• Constant communication between:

• Mars

• Earth

• Phobos

• Cycler

Mangueira 164

DEPLOYABLE MESH ANTENNA

• Based on Northrop Grumman AstroMesh

• Stows roughly in a cylinder

• 10m deployed diameter for Mars to Earth/Cycler

• Assume 100 kg of pointing and signal modulation

hardware per antenna

Mangueira/Sepkovich 165

Antenna Properties Values

Power [kW] 24.01

Mass [kg] 106.4

Stowed Volume [m3] 4.849

Mass and volume based on conversation between Alex Mangueria and Dr. Jeffery Marks of Northrup Grumman

EARTH TO MARS ORBIT

Sepkovich 166

• Direct line of sight from

Earth to Mars

• Signal sent directly from

Earth to Mars

• Constant contact with

cycler from both Earth and

Mars

CYCLER TO EARTH AND MARS

Houtz 167

• Sun interferes with direct

Mars to Earth

communication

• Use cycler to relay signal

around sun

Earth-Cycler = 0.4 AU

Mars-Cycler = 2.6 AU

CYCLER COMMUNICATIONS

Houtz 168

Antenna Properties Units Values

Maximum Power (crew on board)

[kW] 7.964

Maximum Power (communications Relay)

[kW] 24.84

Mass [kg] 365.3*

Stowed Volume [m3] 9.708*

Diameter [m] 10 (x2)

*includes two backup Low-Gain Antennas (LGA’s)

CYCLER RISKS

• Failure would be harmless to crew and possibly

reparable.

• Can use redundancy, small backup antennas, or both to

mitigate risk

• Example: Galileo’s mesh HGA failed to deploy,

mission relied on low gain, S-band antenna.

(1/10,000th the received power on Earth)

• Sending telemetry using lower data rates is very easy,

but crew may not get streaming HD video.

Houtz 169

LOW GAIN ANTENNA

• Calculations for cycler, can be used on any vehicle

• High gain antenna may fail to deploy (no HD video)

• Use low gain antennas for essential communications

• Antenna diameter = 15.91 cm

• Data rate = 40 bps

Sepkovich 170

Receiver Antenna Diameter [m]

Power [W]

10 19,000

34 1655

70 400

Component Mass [kg]

Antenna 0.5368

TWTA 77.82

Total 78.36

RETURN VEHICLE COMMUNICATIONS

Houtz 171


Maximum Power [kW] 4.159

Mass [kg] 261.0*


Diameter [m] 10

*includes two backup Low-Gain Antennas (LGA’s)

MARS ORBIT TO MARS SURFACE

Sepkovich 172

• Areostationary orbit

• 2 satellites offset by 67.2°

• 4 antennas per satellite

• 1 to Earth

• 1 to Mars

• 1 to Phobos

• 1 to other satellite

A B

P

ANTENNAS IN MARS ORBIT

Sepkovich 173

Destination Diameter [m] Power [W] Antenna Mass [kg] TWTA Mass [kg]

Cycler 10 24,010 6.4 97.94

Mars surface 0.6646 8 0.4 1.530

Mars orbit 0.6646 24 0.4 1.595

Phobos Surface 0.6646 35 0.4 1.639

Total - 24,077 7.6 102.7

MARS COMMUNICATIONS SATELLITES

Houtz, 174

Main Power Main

Power Charge Discharge

Satellite #1 Satellite #2

Cycler/Earth Antenna

Satellites switch off communicating from Mars to cycler/Earth every 2 hours

MARS COMMUNICATIONS SATELLITES

Houtz 175


Constant Power Required [kW] 14.31

Mass [kg] 532.1*


Diameter [m] 10 (HGA)

*Totals for entire satellite (HGA (and components), LGA’s (and components), batteries, solar panels, structure)

MARS ORBIT COMMUNICATIONS SATELLITE

Sepkovich 176

Solar Array Properties Values

Diameter [m] 15.5

Mass [Mg] 0.464


Solar array sizing performed by Maxime Pinchaud (Power/Thermal)

ADEPT ANTENNA

• ADEPT 14.55m diameter

• Can add parabolic mesh after landing to make antenna

• Used to communicate from Mars surface to Mars Orbit

Mangueira 177

Based on designs by Ben Libben (Aerodynamics)


Power [W] 8.000

Mass [Mg] 0.0573


ADEPT ANTENNA RISKS

• Difficult to predict how ADEPT will move after it is ejected

• May not land within range of crane

• May impact ground moving very fast and damage structure

• Could add parachute to ADEPT

• Would add extra mass to system

• May not be feasible depending on final ADEPT architecture

Sepkovich 178

BACK UP PLANS FOR ADEPT

Lin 179

Parabolic Antenna

Mass[Kg] 6.4

Power[W] 480

Volume[m^3] 2.4E-3

Uplink: 2.8 GHz Downlink : 3.2 GHz Beamwidth : 10 deg Data rate: 12 Mbps Antenna Diameter : 0.75 m

HD Streaming Communication through satellite

SHORT RANGE COMMUNICATIONS ON SURFACE

• Crane

• Max distance from colony is 6km

• Line of sight communication may be possible

• Rovers

• Max distance from colony is 100km

• Communicate up to satellite in orbit and signal

is sent down to Mars colony

Lin 180

Rover ( Helical Antenna)

Mass [Kg] 1

Power [W] 160

Volume[m^3] 0.008

Uplink: 3.2 GHz Downlink : 2.8 GHz Beamwidth : 19.4 deg Data rate: 1 mbps Receiver Diameter : 0.7 m

LASER COMMUNICATION

PROS

1. Reduced Power

• Due to very high frequency

• Results in small diameter ,and therefore, mass

2. X10 - X100 Better Data Return

• Unregulated and immune to interception so minimal to no interference

3. Better Data Rates

• High Speed and High Capacity

• Ideal for HD video

4. Up and Coming

• Technology is advancing quickly

• Expected to deploy 2034 on LISA

Mangueira 181

CONS 1. Unavailable

• Technology not expected for deep space use ~2030

• Currently too high frequency for AstroMesh conversion

2. Narrow Beamwidth • Makes antenna controls

system required • New. More complex satellite

configuration

PROJECTED PARAMETERS WITH OPTICAL COMMUNICATION

Mangueira 182

Comm. Properties Optical Laser Ka Band

Frequency [GHz] 193,000 32.00

Data Volume [Gb/day] 10.00 10.00

Final Power [kW] 0.0446 20.00

Total Mass [kg][2] 12.10 104.3

Parameter Comparison of Earth to Cyclers

Assuming: 1. 267 [Kpbs] data rate (enough for HD streaming) 2. Ka Band has an additional 6.370 [kg] for AM-Lite

QUESTIONS?

REFERENCES

[1] K. Wilson, and M. Enoch, "Optical Communications for Deep Space Mission," IEEE Commun. Magazine 38,134-139 (2000).

[2] H. Hemmati et al., "Comparative Study of Optical and Radio Frequency Communication Systems for a deep space Mission," JPL TDA Prog. Rep. 42-128, Feb., 1997.

[3] M. Toyoshima, “Trends in Satellite Communications and the Role of Optical Free-Space Communications,” Journal Of Optical Networking, 2005 Jun, Vol.4(6), pp.300-311

[4] Northrop Grumman, AstroMesh Reflector Parametrics, Aero Astrospace Headquarters,

Los Angeles, California, January 2013, file:///Users/mac/Desktop/AMLite.pdf

[5] Dr. Jeffery Marks, Chief Enginner of AstroMesh, Northrop Grumman Aero Astrospace

Headquarters, Phone Call

[6] Northrop Grumman, AstroMeshLite (AM-Lite), Aero Astrospace Headquarters, Los

Angeles, California, January 2013, http://www.northropgrumman.com/BusinessVentures/

AstroAerospace/Products/Documents/pageDocs/Parametrics.pdf

Mangueira 184

EARTH ACADEMIC ANTENNAS Locations in:

• Moorehead, Kentucky (20m)

• Bochum, Germany (20m)

• Bangkok, Thailand (12m)

• Most common size is 8m

• Less common 12m

• Rare 20m+

• Assumed 10m receiving antennas for communications to Earth

• Larger antennas reduce power requirements

Mangueira 185

ANTENNA LATITUDE/LONGITUDE

Sepkovich 186

Antenna Latitude Longitude Country

Morehead 38°11’31’’N 83°26’20’’W USA (KY)

Bochum 51°25′40″N 07°11′39″E Germany

CALCULATIONS OF DEPLOYABLE SATELLITES

Mangueira 187

Earth to Mars/Cycler

Mars to Mars Orbit


AstroMesh Mass [kg] 2.5480

Aluminum Structure Mass [kg]

3.8220

Total Mass [kg] 6.3700

Stowed Height [m] 1.4000

Stowed Diameter [m] 2.1000


Antenna Parameters Values

AstroMesh Mass [kg] 22.932

Aluminum Structure Mass [kg]

34.397

Total Antenna Mass [kg] 57.329

Stowed Height [m] 2.2000

Stowed Diameter [m] 1.1500


Power [W] 5

*Calulations derived from example of antenna diameter = 9 [m], weight = 4.55 [lbs] given by Dr. Jeffery Marks of Northrup Grumman

ALTERNATIVE RETURN VEHICLE COMMS

Houtz, 188

20m

20m 20m

20m

8m

8m

Main Power

Main Power Charge Discharge

• MASS & IMLEO

• POST-MITIGATION RISK ON MARS

• PRE-MITIGATION RISK ON MARIO

• STORYBOARD: XM1, XM2, AND MARS SURFACE

JANI DOMINGUEZ

ASST. PROJECT MANAGER

03/12/2015

MASSES BY DISCIPLINES

Wisser - 190

- A breakdown of each vehicle being used in the mission and what each team has

contributed to every vehicle.

- Each color signifies the spreadsheet/Vehicle Group it came from. Grey signifies that

that discipline wasn’t utilized in the mass analysis for the vehicle.

IMLEO MASSES

Wisser - 191

- Definitive numbers given by Project Manager for multiplicity of vehicles needed for full

mission success.

- Number of cargo missions is dependent on farming success and the immediate needs

of the crew on Phobos and Mars.

**The propulsion system transporting the Mars Return vehicle has not been determined

RISK ON MARS

Pre-Mitigation Post-Mitigation

Dominguez, 192

RISK ON MARS CONT. Rank Risk Prob.

1 Aerocapture[1] 99.41%

2 ADEPT[2] 98.50%

3 Hyperbolic Rendezvous

98.40%

4 Tethered Ballute 99.00%

5 XM3 Thermal Management

99.80%

6 Cycler Engine Failure 99.85%

7 HuLa: Fligth Path

Angle 99.90%

8 Phobos to Mars

Lander 99.80%

9 Reaction Wheels 99.40%

10 Cycler Structural

Failure 99.87%

Probability of Success: 93.84%

Rank Risk Prob.

1 Aerocapture 99.86%

2 ADEPT 99.00%

3 Hyperbolic Rendezvous

98.70%

4 Aerobraking 99.86%

5 M2P Solid Prop 97.66%

6 Mid L/D Failure 98.60%

7 Delta V/ incorrect

trajectory 98.50%

8 Ballute Tethered 98.60%

9 Cycler Engine Failure 95.00%

10 HuLa: Flight Path

Angle 94.60%

Probability of Success%: 81.89

Dominguez, 193

PRE-MITIGATION: RISK ON MAREO Rank Risk Prob.

1 In-Situ Resource

Utilization 97.50%

2 MAM 98.20%

3 Uncertainty with

Martian Atmosphere 97.20%

4 Engine Failure 99.90%

5 Phobos Launch 99.80%

6 Delta V - Trajectory 99.85%

7 Solid Prop System 99.90%

8 Re-entry Earth 99.85%

9 BA 330 Deployment 98.10%

10 Guidance Systems 99.95%

Probability of Success: 90.70%

Dominguez , 194

REFERENCES [1] Kremic, T., Munk, M.M., “Aerocapture Summary and Risk Discussion,” NASA Presentation, March 26, 2008. [2] Venkatapathy, E., and Glaze, L., “ADEPT-VITaL Mission Feasibility Report,” NASA Ames Research Center, Aug. 2013

Dominguez , 195

FINAL QUESTIONS?

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