FINAL DESIGN REVIEW 3/12/2015 - Purdue Engineering · TOMI OLOKUN NICOLE VAUGHN MISSION DESIGN...
Transcript of FINAL DESIGN REVIEW 3/12/2015 - Purdue Engineering · TOMI OLOKUN NICOLE VAUGHN MISSION DESIGN...
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FINAL DESIGN REVIEW
3/12/2015
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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
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MISSION SCHEDULING
• DEVELOPMENT PLAN
• PRODUCTION NEEDS
STEPHEN WHITNAH
PROJECT MANAGER
03/12/2015
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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
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DEVELOPMENT SCHEDULE
2015 2020 2025 2030 2035 2040
Rigid XM Missions to
L1, L2, asteroid,
moon bases
Human Exploration of Mars
Proposed vehicle development schedule
XM3, Rover, Crane
Whitnah 5
Cycler and Boost Module, Cargo Vehicle
Human Lander
Return Option
Mars mission hardware production begins
Inflatable BA 330
Missions
Orion and SLS-1B development ends (EM-2)
50 years since Apollo 11
Cycler Establishment
A B Human Launches
A B
Cargo Missions
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PRODUCTION SCHEDULE
2015 2020 2025 2030 2035 2040
Rigid XM Missions to
L1, L2, asteroid,
moon bases
Human Exploration of Mars
Whitnah 6
Mars mission hardware production begins
Inflatable BA 330
Missions
Orion and SLS-1B development ends (EM-2)
50 years since Apollo 11
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
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PRODUCTION SCHEDULE
2015 2020 2025 2030 2035 2040
Human Exploration of Mars
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
Mars mission hardware production begins
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)
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QUESTIONS?
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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
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EMILY ZIMOVAN
ALEX DAVIS
LORENZO GARCIA
PETER GELDERMANS
PABLO MACHUCA
TOMI OLOKUN
NICOLE VAUGHN
MISSION DESIGN
FINAL DESIGN REVIEW
MARCH 12, 2015
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STORY BOARD
Zimovan
S1L1 Cyclers A/B & Communications
Depart Cyclers
Cargo Supply
P-M Return to Earth
Hyperbolic Rendezvous
LEO Ops Moon
ComSats
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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
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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
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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)
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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
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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
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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
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COMMUNICATION VIA CYCLERS
2 - MACHUCA
Communication via cyclers provide continuous communication link
• Two antennas on each cycler
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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
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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
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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
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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
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POSSIBLE FIRST CARLA LAUNCH DATES
23
Several Low Cost Options in Ephemeris Models. Launch Date Selected: 7/1/35
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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
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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
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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
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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
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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
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QUESTIONS?
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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.
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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
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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
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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
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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
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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
Life Support Systems Power
Systems
Thermal Control Systems
XM3-M (Farming)
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XM3-M VARIANTS (CONTINUED)
Eads 36
Med Bay
Crew Quarters
Crew Quarters
Water Systems
Life Support Systems Power
Systems
Thermal Control Systems
XM3-M (Medical) XM3-M (Water)
Water Systems
Crew Quarters
Crew Quarters
Life Support Systems
Power Systems
Thermal Control Systems
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Crew Quarters
Crew Quarters
Crew Quarters
CREW QUARTERS LAYOUT
Pottebaum-37
Stowage Food
Tunnels
Bathroom Personal Space
Maintenance/Tools
Washer/Dryer
Exercise Area
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MED BAY LAYOUT
Eads 38
Med Bay
General Med Racks
Medical Tables
Storage Racks
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UTILITIES FLOOR LAYOUT (STANDARD)
Skidmore - 39
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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
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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
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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
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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
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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)
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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)
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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*
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QUESTIONS?
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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
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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
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SYSTEM VOLUMES (M^3)
Eads 50
System XM3-C XM3-P XM3-M (Core) XM3-M (Farming)
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
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SYSTEM VOLUMES (M^3) CONT.
Eads 51
System XM3-M (Water) XM3-M (Medical)
Structural 7.648 7.648
Communication .81 .81
Power Systems 0.03969 0.03963
Med Bay 0 32
Crew Quarters 207.6 207.6
Water Systems 27.93 7.33
Life Support 7.1 7.1
Thermal Control .64 .64
Farming 0 0
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SYSTEM POWER REQUIREMENTS (KW)
Eads 52
System XM3-C XM3-P XM3-M (Core) XM3-M (Farming)
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
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SYSTEM POWER REQUIREMENTS (KW)
Eads 53
System XM3-M (Water) XM3-M (Medical)
Communication 1 1
Power Systems 0 0
Med Bay 0 0
Crew Quarters 27.87 27.87
Water Systems - -
Life Support 62.1 62.1
Thermal Control 16 16
Farming 0 0
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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
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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
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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
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CYNTHIA ROSE
BEN LIBBEN
BEN TACKETT
ROHAN DUDANEY
ZAK SIPICH
AERODYNAMICS
3/12/15
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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
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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
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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)
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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
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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
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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
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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
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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
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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.
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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
Target Ballistic Coefficient = 20 kg/m2
Mass of HIAD = 7.14 Mg
Volume of Deployed HIAD = 4230 m3
Max Radius (rb) = 24.4 m
Torus Radius (rt) = 1.4 m
Maximum Radius (rb) Torus Radius (rt)
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HIAD SIZING FOR EARTH RETURN MISSION
Sipich - 68
HIAD Sizing for Return Mission Total Mass of Return Vehicle = 21.6 Mg
Target Ballistic Coefficient = 25 kg/m2
Mass of HIAD = 2.1 Mg
Volume of Deployed HIAD = 543 m3
Max Radius (rb) = 12.4 m
Torus Radius (rt) = 0.58 m
Maximum Radius (rb) Torus Radius (rt)
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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
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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
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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
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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
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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%
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QUESTIONS?
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TRAJECTORY ERROR REDUCTION ANALYSIS
Tackett 75
Atmospheric Error
(% of average density)
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
(% of average density)
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
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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
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ARCHITECTURE DESCENT EXAMPLE: HULA
Tackett 77
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HYPERSONIC INFLATABLE AERODYNAMIC DECELERATOR (HIAD)
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
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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.
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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.
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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.
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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.
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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
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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
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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.
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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
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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
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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)
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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QUESTIONS?
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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
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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
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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
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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
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HYPERSONIC INFLATABLE AERODYNAMIC DECELERATOR (HIAD)
Sipich - 111
Maximum Radius (rb)
Torus Radius (rt)
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 m
Torus Radius (rt) = 1.36 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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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QUESTIONS?
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HIAD FOR XM3 DELIVERY
Sipich - 124
Final XM3 HIAD Delivery Sizing Total Mass of XM3 = 73.14 Mg
Target Ballistic Coefficient = 20 kg/m2
Mass of HIAD = 7.14 Mg
Volume of Deployed HIAD = 4230 m3
Max Radius (rb) = 24.4 m
Torus Radius (rt) = 1.4 m
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
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Backup
Lee - 125
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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
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CYCLER FINAL DESIGN
127
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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
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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
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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
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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
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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%
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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
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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
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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
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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
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QUESTIONS?
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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
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CYCLER’S SOLAR PANELS
Maxime PINCHAUD - 139
(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
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BACK UP SLIDE
Maxime PINCHAUD - 140
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
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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
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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)
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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.
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DANIEL GOLDBERG, MAX DZIS, TRISTAN LOUDEN, LEE WESTROPP, ALIBEK YERTAY, YUE GUO, MORGAN LUCAS, NICOLE VAUGHN, JOCELINO RODRIGUES
MARS RETURN OPTION
12 MAR 2015
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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
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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
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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
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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
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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 Mass (Mg) 0.040
Total Volume (m3) 0.014
BA-330: 2 Circular ATK Panels
Total Power (kW) 13.96
Total Mass (Mg) 0.182
Stowed Volume (m3) 0.349
Deployed Area (m2) 173
Diameter (m) 10.5
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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
Total Mass [Mg] 0.042
Power [W] 1.7
Total Volume [m^3] 0.25
Sensors
Star Sensors
Number 2
Total Mass [Mg] 0.01
Power [W] 20
Sun Sensors
Number 2
Total Mass [Mg] 0.004
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
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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
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MAM
Dzis 152
Total Volume 24.52 m3
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
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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
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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
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● 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
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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)
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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
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MAM DESGIN
Louden 158
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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
Total Volume 0.682 m3
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
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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
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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
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QUESTIONS?
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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
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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
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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
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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
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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
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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)
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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
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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
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RETURN VEHICLE COMMUNICATIONS
Houtz 171
Antenna Properties Units Values
Maximum Power [kW] 4.159
Mass [kg] 261.0*
Stowed Volume [m3] 4.859*
Diameter [m] 10
*includes two backup Low-Gain Antennas (LGA’s)
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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
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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
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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
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MARS COMMUNICATIONS SATELLITES
Houtz 175
Antenna Properties Units Values
Constant Power Required [kW] 14.31
Mass [kg] 532.1*
Stowed Volume [m3] 7.283*
Diameter [m] 10 (HGA)
*Totals for entire satellite (HGA (and components), LGA’s (and components), batteries, solar panels, structure)
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MARS ORBIT COMMUNICATIONS SATELLITE
Sepkovich 176
Solar Array Properties Values
Diameter [m] 15.5
Mass [Mg] 0.464
Stowed Volume [m3] 0.6
Solar array sizing performed by Maxime Pinchaud (Power/Thermal)
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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)
Antenna Properties Values
Power [W] 8.000
Mass [Mg] 0.0573
Stowed Volume [m3] 2.285
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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
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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
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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
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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
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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
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QUESTIONS?
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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
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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
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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
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CALCULATIONS OF DEPLOYABLE SATELLITES
Mangueira 187
Earth to Mars/Cycler
Mars to Mars Orbit
Antenna Properties Values
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
Stowed Volume [m3] 4.8490
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
Stowed Volume [m3] 2.2850
Power [W] 5
*Calulations derived from example of antenna diameter = 9 [m], weight = 4.55 [lbs] given by Dr. Jeffery Marks of Northrup Grumman
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ALTERNATIVE RETURN VEHICLE COMMS
Houtz, 188
20m
20m 20m
20m
8m
8m
Main Power
Main Power Charge Discharge
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• 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
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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.
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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
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RISK ON MARS
Pre-Mitigation Post-Mitigation
Dominguez, 192
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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
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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
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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
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FINAL QUESTIONS?
Thank you for attending