Download - Centurion - OTV Presentation
AE 443S - TEAM 4APRIL 16TH, 2015
FDR PRESENTATION
2
THE TEAM
Jay Mulakala
Lead Systems Engineer
Samip ShahADCS
Systems Engineer
Bentic SebastianPower and
Thermal Systems Engineer
Yu Guan
Structures
EngineerBen Wilson
Propulsion Engineer
Derek AwtryOrbital
Systems Engineer
Kevin Lohan
Launch and Docking Engineer
PLAN OF ACTION
3
Mission Summary
Vehicle Systems Overview
Mission Architecture
Risk Analysis
Mission Costs
MISSION SUMMARYJAY MULAKALA
THE CRITERIA
5
The OTV will be stationed in 400 km AMSL circular LEO with 28° inclination.
The OTV payload capability shall be 50,000 lbs from LEO to EML1 and 15,000 lbs from EML1 to LEO.
The OTV must be capable to remain at EML1 or EML2 for at least 30 days.
Each transfer should not exceed 6 days.
The life of the OTV shall be 5 years and the OTV shall be capable of at least 10 missions to EML1 or EML2.
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CENTURION
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CENTURION
THE CENTURION
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TIMELINE
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10
TIMELINE
VEHICLE SYSTEMS OVERVIEWJAY MULAKALA
MISSION SYSTEMS
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• Orbital Systems• Spacecraft Propulsion Systems• Structural Definition• Communications• ADCS• Spacecraft Power Management
Systems• Spacecraft Thermal Systems• Launching and Docking Systems
DEREK AWTRYORBITAL SYSTEMS
DESIGN PROCESS Orbital Requirements
Maximum transfer time of 6 days to L1 and L2 Orbit around L1/L2 for at least 30 days Initial Low Earth Orbit of 400 km and 28˚ inclination Can consider Aerobraking
STK/Astrogator was used to determine trajectories to L1 and L2 Multiple trajectories were considered The trajectories with the lowest V was chosen for each Lagrange
point
We will not consider aerobraking14
CONCEPT DEVELOPMENT
15
Trajectory to L1 Different halo orbits were simulated by varying the z-amplitude
on the Earth-Moon plane. Each initial burn from LEO is 3.069 km/s, each burn back into
LEO was 3.058 km/s, with a time of flight (TOF) there and back of 4.3-4.4 days
Orbit Option Amplitude
(km)
(m/s) (m/s) Total (m/s) Halo Orbit
(days)
1 5,000 620.017 644.108 7411.944 35.999
2 7,500 622.213 644.063 7421.159 36.146
3 10,000 625.070 647.151 7422.128 36.108
4 15,000 632.660 654.051 7446.725 36.079
5 20,000 642.570 663.845 7482.638 36.140
V’s for different halo orbits
CONCEPT DEVELOPMENT
16Trajectory to L1 using STK
CONCEPT DEVELOPMENT
17
Trajectory to L2 Different halo orbits were simulated by varying the z-amplitude
on the Earth-Moon plane. Each initial burn from LEO is 3.094 km/s, each burn back into
LEO was 3. 099 km/s, with a TOF there of 5.3 days and a TOF back of 5.9 days
Orbit Option
Amplitude (km) (m/s) (m/s) Total (m/s) Halo Orbit
(days)6 5,000 1124.80
61018.23
0 8352.291 44.277
7 7,500 1122.052
1030.948 8372.584 45.08
9
8 10,000 1118.017
1003.735 8339.403 45.17
1
9 15,000 1106.847 988.601 8321.102 44.98
5
10 20,000 1097.024 974.496 8306.030 44.84
1
V’s for different halo orbits
CONCEPT DEVELOPMENT
18Trajectory to L2 using STK
CONCEPT DEVELOPMENT
19
Station-Keeping Analysis A station-keeping maneuver was added for each orbit
simulated is the station-keeping burn after the first revolution is the station-keeping burn after the second revolution
Orbit Option (m/s) (m/s) Total V
(m/s)1 9.586 9.952 19.538
2 13.588 13.069 26.657
3 7.352 14.37 21.722
4 7.834 24.048 31.882
5 10.311 37.805 48.116
Station-keeping maneuvers at L1
Station-keeping maneuver at L2Orbit
Option (m/s) (m/s) Total V (m/s)
6 8.386 6.922 15.308
7 8.729 16.92 25.657
8 9.412 14.139 23.551
9 12.765
18.727 31.492
10 17.744
22.860 40.604
CONCEPT DEVELOPMENT
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Orbital Maintenance In LEO Refueling in LEO will take place at the end of the 4-5 month period
between missions to get back to required orbit is 6.797 m/s
End of Life Summary A payload will be taken on a one way mission to L1, and
dropped off Centurion will then maneuver to the stable EM-L4
from L1 to L4 is around 682 m/s Final mission = 4.914 km/s
CONCEPT DEVELOPMENT
21
Equation derived by NASA for V savings Initially used for the Martian atmosphere, can
be expanded for all celestial bodies [3]
Altitude (km) Atmospheric Density
(kg/km^3)
V savings (m/s)
50 102700.0 1,026.4
60 30960.0 295.6152
80 18449.456 157.9763
100 560.276 4.7543
120 22.234 0.2563
140 3.839 0.552
savings for different periapsis altitudes
BENJAMIN WILSONPROPULSION SYSTEMS
DESIGN PROCESS
Main Propulsion System Requirements • For a round trip to L2 provide a total V of 8.31 km/s
• Minimize fuel mass• Ensure safe operation for passengers • Reliable
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Attitude Control Propulsion System
Requirements • Provide the total V required for attitude control over the entire operational lifetime of the Centurion
• Provide a total V of 10 m/s• Minimize fuel mass• Reliable
CONCEPT DEVELOPMENT: MAIN PROPULSIONType Thruster Thrust
[N]Specific Impulse [S]
Fuel to L2[kg]
Ion Aerojet NEXT 0.235 4100 8,700
Busek BHT-20k 0.807 2320 16,600
NASA NSTAR 0.094 3195 11,600Bipropellant Aerojet CECE 111,000 465 183,00
Astrium Aestus 29,600 324 436,000CALT YF-73 44,150 420 229,000
Monopropellant Aerojet MR-80B 3780 225 1,410,000AMPAC MONARC 445
445 235 1,190,000
Nuclear Thermal
NERVA XE 1,112,000 850 62,500Escort 333,600 911 49,600CIS NTR 111,600 941 47,500
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[9] [10] [11] [12] [13] [14] [15] [16]
Escort BNTR System Characteristics
System Number of Units 3Total System Mass 6675 kg
Thrusters Thrust Per Unit 111,200 NSpecific Impulse 911 s
Propellant Propellant Liquid HydrogenPropellant Mass 49,959 kgPropellant volume 700 m3
Tank Material Composite Tank Boil Off 1% per month by mass
ReactorsFission Material
UO2-W Cermet (Uranium Dioxide in Tungsten matrix)
Exhaust Temperature ~ 2700 Kelvin
Power & Thermal Power Per Unit 25kW
Power Cycle Brayton CyclePower Fluid Helium XenonRadiator Area 65m2
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NUCLEAR REACTOR SAFETY Shielding to reduce exposure to less than 1 REM/year Reactor shutdown while docking and refueling No critical state before leaving the atmosphere Not allowed to re-enter atmosphere until all fissile material
has been used Will be decommissioned at EML4
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MAIN PROPULSION SYSTEM PROPELLANT AND TANKAGE
Liquid Hydrogen propellant
Ensures Lowest Fuel mass
700m3 composite fuel tank
Active and passive Thermal control to achieve ~1% boil off per month
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NASA and Boeing’s 5.5 Meter cryogenic composite fuel tank [61]𝐼 𝑠𝑝=𝐴𝐶 𝑓 √ 𝑇𝑐
𝑀
PROPELLANT MASS CALCULATION TO L2 [ISP=911S]
Stage Description V [m/s] Time Elapsed
Mi (Includes Payload) [kg]
Mpi Propellant Spent [kg]
1 Departing LEO 3095 10 Minutes 87,745 25,6822 Transit to L2 0 5.3 Days 62,063 423 Arrive at L2 1097 3 Minutes 62,021 7,1654 Halo orbit 1 0 15 days 32,177 835 Halo
Correction 1 18 6 Seconds 32,094 646 Halo Orbit 2 0 15 days 32,030 837 Halo
Correction 2 23 6 Seconds 31,947 828 Halo Orbit 3 0 15 days 31,866 839 Depart L2 974 2 Minutes 38,588 3,986
10 Transit to LEO 0 6 days 34,602 2511 Arrive at LEO 3099 4 Minutes 34,577 10,133
Final Mass = 24,444
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𝑀𝑝=𝑀 𝑠𝑦𝑠 (𝑒∆𝑉 / 𝐼 𝑠𝑝𝑔𝑜−1)
CONCEPT DEVELOPMENT: ATTITUDE CONTROL PROPULSION
Type Engine Fuel Isp [s]
Thrust [N] Propellant Mass[ kg]
Ion Aerojet NEXT Xenon 4100 0.235 13Busek BHT-20k Xenon 2320 0.807 23
Cold Gas MOOG 58-118 Unknown 72 3.5 560AMPAC SVT01 Xenon 45 0.05 900
Monopropellant
AMPACMONARC -90 Hydrazine 235 90 215Aerojet MR-107N Hydrazine 232 109-296 220
BipropellantEADS 10N
NTO, MON-1, MON-3 and MMH
291 10 195
Aerojet R-1E MMH/NTO 280 111 190
30
[51] [52] [53]
ACS PROPELLANT AND TANKAGE
Tank Propellant
Volume (L)
Mass (kg)
Tanks Required
MOOG GEO Sat.
Hydrazine 220 27 4
ATK 80505-1 Any 134 16 4Astrium OST 31/0
MON/MMH 235 16 4
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Propellant Volume (L)
Mass (kg)
Mono methyl hydrazine (MMH)
83 73
Nitrogen tetra oxide (NTO) 81 117
ATK 80505-1[63]
YU GUANSTRUCTURAL DEFINITION
CONCEPT DEVELOPMENT
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• Systems module• ADCS sensors• Power and Thermal units
• Propellant tank• Thermal shielding
for cryogenic fuel
• Propulsion systems • Escort System• Radiation shield
SYSTEMS MODULE
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CRYOGENIC PROPELLANT TANK
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• Internal volume of 625 cubic meters• Aluminum lithium thermal shielding
PROPULSION SYSTEM
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STRUCTURAL COMPONENTS
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CYCOM 5320-1 epoxy resin systemfrom Cytek Inc. [12]
PAMG-XR1 aerospace grade aluminum honeycomb from Plascore Inc. [13]
RADIATOR DESIGN
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Hinged radiator used in systems module. [14]
Deployable radiator installed in propulsion system. [15]
MASS ESTIMATION
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STRUCTURAL TESTING
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YU GUANCOMMUNICATIONS
DESIGN PROCESSNetwork Selection
Network ideally suited for needs of the CenturionBand Selection
Allow for uninterrupted communication without incurring high pointing accuracy requirements
Radiometric TrackingEnable accurate position and velocity determination of the
CenturionAntenna Selection
Deploy methodMassGain
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CONCEPT DEVELOPMENT – NETWORK SELECTION
NEN(Near Earth Network) performance is comparable to DSN(Deep Space Network) at L1 and L2 regions(EIRP=85 dBmW)
Fewer missions in L1 and L2 using NEN Less traffic in wireless
communications.
Network of choice: Near Earth Network
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CONCEPT DEVELOPMENT – FREQUENCY BAND SELECTION
Several bands S-band
Low frequency Low data rates Lower pointing requirements Low atmospheric attenuation
X-band Frequencies, data rates, pointing
requirements, and attenuation in between S and Ka-band
Ka-band High frequency High data rates Higher pointing accuracies Higher atmospheric attenuation
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BandFrequency
(GHz)
S-band 2-3
X-band 7-11
Ka- band 18-30NEN Frequency Band Characteristics [2]
Atmospheric attenuation as a function of frequency [1]
CONCEPT DEVELOPMENT – RADIOMETRIC TRACKINGDoppler Provides velocity estimates Orbital maneuvers
ΔV calculations
Ranging Provides position estimates Orbital maneuvers
Orbital transfer points Docking maneuvers
Get within range of fuel depot and payloads to use proximity sensors
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Characteristics ValueRanging Accuracy 10 Meters (1 sigma)Doppler Accuracy 1 millimeter per second (1 sigma), 5 second
integration timeAngle Accuracy 0.1 Degrees
Maximum Velocity 2.0 Degrees/second (az and el)Near Earth Network Tracking Characteristics [4]
CONCEPT DEVELOPMENT
Parabolic reflectorX & S bandLow gain as backupFoldable design 46
CRITICAL DESIGN ISSUES
Structural Critical Issues
- Thermal Cycling
- Material Failure Communication Critical Issues
- Costs of components
- Lack of details on ground station
47
SAMIP SHAH
ATTITUDE DETERMINATION & CONTROL SYSTEMS
DESIGN PROCESS
Sensor SelectionHigh attitude sensing requirements for docking and refueling
Actuator SelectionHigh pointing requirements for docking and refuelingAbility to actively mitigate disturbance torquesLarge volume and mass require high performance actuators
Control SystemsSensor processing and actuator controlRedundancy and flight proven hardware for reliability
49
CONCEPT DEVELOPMENT – SENSOR SELECTION
Star Trackers Absolute attitude sensor Highest accuracy Low update rate
Selection: Surrey Rigel-L 2 Units
50Characteristics of common star trackers [22] [23] [24]
Surrey Rigel-L Star Tracker
Manufacturer/Model Lifetime (yrs)
Max Resolution(arc sec)
Update Rate (Hz)
Tracking Rate
(deg/s)
Surrey/Rigel-L 7+ X/Y < 3Z < 25 1-16 6
Terma/HE-5AS N/A RMS pitch, yaw <1RMS roll <5 4 0.5-2.0
Jena Optronik/ASTRO APS >18 X/Y < 1
Z < 8 10 0.3-3
CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.)Inertial Measurement Units Relative attitude sensor Prone to drift High update rates
Selection: Honeywell HG9848 2 Units
1 Backup
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Manufacturer/Model
Gyro Bias Repeatability/
Stability (deg/h)
Gyro ARW (deg/√hr)
Gyro Scale Factor (ppm)
Accel Bias Repeatability
(μg)
Scale Factor (ppm)
Northrop Grumman/LN-200S
1/<0.1 <0.07 100 300 300
Honeywell/HG9848 <0.005 <0.005 <10 <50 150
Kearfott/KI-4901 0.005/0.003 .003 50 400 500
Characteristics of common IMUs [26] [27] [28]
CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.)Sun Sensors Useful for solar array pointing Relatively inexpensiveSelection: Adcole Course Sun Sensor Pyramid 2 Units
52Characteristics of common sun sensors [5]
Adcole Course Sun Sensor Pyramid
Manufacturer/Model Field of View Accuracy (deg) # required for full
coverage
Adcole/Digital Sun Sensor +/- 64 deg 0.25 5
Adcole/Course Sun Sensor Pyramid 2π steradians 1 2
CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.)Proximity Sensor Vital for autonomous docking maneuvers Enables accurate knowledge of relative position and attitude
53
Distance to Target X (m) Y (m) Z (m) Pitch
(deg)Yaw
(deg)Roll
(deg)Range
(m)
2m 0.0003 0.0007 0.0029 0.25 0.09 0.06 0.003
30m 0.0098 0.0039 0.0061 1.04 0.81 0.33 0.012
Demonstrated Accuracy of AOS Proximity Sensors [15]
OperationApproach Velocity
(m/s)
Lateral Alignment
(m)
Lateral Velocity
(m/s)
Angular Misalignmen
t (deg)Angular Rate
(deg/s)
Docking 0.3 0.2 0.05 5 0.25Berthing 0.01 0.5 < 0.01 < 10 < 0.1
Capture Tolerances for Docking and Berthing [14]
CONCEPT DEVELOPMENT – ACTUATOR SELECTIONAttitude Control Thrusters Provide both large and
fine attitude adjustments
16 Units Clusters of 4
Roll slew time - ~4 min Pitch/Yaw slew time -
~6 min
54
Manufacturer/Model Thrust (N)Specific Impulse
(sec)Propellant
Vacco/2 LBF Cold Gas 8.9 - GN2
Moog/DST-11H 22 310 Hydrazine/MON
Aerojet/R-1E 111 280 MMH/NTO
Aerojet/MR-107V 67 229 HydrazineCharacteristics of common thrusters [31] [32] [33] [34]
Configuration of attitude control thrusters
CONCEPT DEVELOPMENT – ACTUATOR SELECTION (CONT.)
Control Moment GyroscopesCommon on large spacecraftEasily mitigates disturbance
torques of roughly 4 x 10-3 NReduces thruster firings
Actuate spacecraft in ADCS thruster failure emergency
6 units total2 units per axis1 backup per axis
Roll slew time - ~25 minPitch/Yaw slew time - ~40 min
55
Manufacturer/ModelAngular
Momentum
(N-m-s)
Output Torque
(N-m)
Weight
(kg)
Power Requirements
(W)
Honeywell/M50 25-75 0.075-75 28 95 @ peak torquing
L-3/DGCMG 4800/250 4760 258 272 -
Airbus/CMG 15-45s 15 45 18.4 25Characteristics of commonly used control moment gyroscopes [37] [38] [39]
56
CONCEPT DEVELOPMENT – CONTROL METHODS
Onboard Processing Comprehensive processing is required to control attitude and
position RAD750
Powerful, dependable, flight proven 3 used together for redundancy
57
Characteristics of radiation hardened flight processors [42] [43] [44]
Manufacturer/Model Speed (MIPS) Power (W) Flight ProvenIBM/Rad6000 35 10 YesIBM/Rad750 400 10 YesProton300k 8000 12 Yes
CONCEPT DEVELOPMENT – CONTROL METHODS (CONT.)
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SOURCE LINES OF CODEComputer Software Component SLOCExecutive 1,000Communications 2,000Attitude/Orbit Sensor ProcessingSun Sensor 500IMU 1,000Star Tracker 2,000Attitude Determination and ControlKinematic Integration 2,000Kalman Filter 8,000Error Determination 1,000Orbit Propagation 10,000Attitude Actuator ProcessingThruster Control 1,000CMG Control 1,500Fault Detection 10,000UtilitiesBasic Mathematics 1,000Transcendental Mathematics 1,500Matrix Mathematics 2,300Time Management & Conversion 700Coordinate Conversion 2,500Other FunctionsMomentum Management 3,000Power Management 2,000Thermal Control 1,500Total 54,500
BENTIC SEBASTIANSPACECRAFT POWER MANAGEMENT
DESIGN PROCESSFour performance requirementsSupply power for all instruments onboard.Provide suitable radiation shielding.Provide suitable temperature for onboard
electronics.Provide backup power when there is no power
generation.
61
PROPOSED SOLUTIONESCORT Bimodal Nuclear Rocket Thermal Propulsion System.Use radiation hardened instruments in Systems Module, and 3cms of
shielding on reactors.Hinge radiators used to control temperature of Systems Module.Solar panels during emergency mode
CONCEPT DEVELOPMENT- POWER GENERATION AND DISTRIBUTION
62
Main source of power is BNTRS Three Nuclear Reactors, producing 25 kW each. Will be run at 2/3rds of maximum output, 50kW
CONCEPT DEVELOPMENT- POWER GENERATION AND DISTRIBUTION
Peak voltage = 50V for Surrey/Rigel-L star tracker Peak power input = 113 W for Honeywell/M50 moment gyroscope
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System component Voltage Required(V) Power Required(W)
Surrey/Rigel-L star tracker 16-50V 0.5-6.5W
Honeywell/HG9848 IMU 5V 10W
Adcole/Coarse Sun Sensor Pyramid 0V 0W
Aerojet/R-1E attitude control thruster 28V 36W
Honeywell/M50 moment gyroscope 28V 11-113 W
IBM/Rad750 2.5-3.3V 5W
POWER GENERATION AND DISTRIBUTION
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CONCEPT DEVELOPMENT – POWER STORAGE
Requirements for batteries Light Rechargeable High energy density Long discharging time
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CONCEPT DEVELOPMENT – POWER STORAGE
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Technology Specific
Density(Wh/kg)
Energy
Density(Wh/l
)
Operating
temp.
Range(C)
Design
life(years)
Cycle life
Ag-Zn 100 191 -20 to 25 2 100
Ni-Cd 34 53 -10 to 25 3 25,000-40,000
Super Ni-Cd 28-33 70 -10 to 30 5 58000
IPV Ni-H2 8-24 10 -10 to 30 6.5 At least 60000
CPV Ni-H2 30-35 20-40 -5 to 10 10-14 50,000
SPV Ni-H2 53-54 70-78 -10 to 30 10 At most 30,000
Lithium Ion 90 250 -20 to 30 1 At least 500
To store at least 1.4 kW of power: Weight of IPV Nickel-Hydrogen batteries: 83 kg Weight of Lithium batteries(Quallion QL075KA):20 kg
CONCEPT DEVELOPMENT – POWER STORAGE
67
Design choice: Eight Quallion QL075KA batteries, 72Ah, 3.6V Eight additional batteries for redundancy. Total weight of 16 Quallion batteries: 20kg
Advantages of battery. Quallion QL075KA has a cycle life of 100,000 cycles. No need to replace batteries.
CONCEPT DEVELOPMENT – RADIATION SHIELDING
Material for Radiation Shielding: Aluminum
Material Thickness 3cms
Shielding placed in front of Nuclear Reactors
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CONCEPT DEVELOPMENT – EMERGENCY MODE Two solar panels will generate additional energy during emergency
conditions Minimum energy required for operations:
1.4kW Material for Solar Panels:
Amorphous Silicon Solar panels will have total surface area of 7.7 meters square Accounting for efficiency of 15%, solar panels provide 23.3 W of power, as
emergency power for computers
69
BENTIC SEBASTIANSPACECRAFT THERMAL SYSTEMS
DESIGN PROCESS
Control temperature withinNuclear ReactorsFuel TankSystems ModuleSolar Panels and Radiators
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PROPOSED SOLUTIONTwo radiators for Nuclear ReactorsTwo radiators for Systems ModuleCryogenic fuel tank with MLI and SOFI
CONCEPT DEVELOPMENT – THERMAL CONTROL OF NUCLEAR REACTORS
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Brayton Cycle of ESCORT System [29]
Peak temperature of HeXe working fluid =929K
CONCEPT DEVELOPMENT – THERMAL CONTROL OF NUCLEAR REACTORS
Deployable radiators designed by Lockheed Martin.
Total area of 65 meters squared.
Will use ammonia coolant loops to control temperatures on radiator.
Radiators will use beta gimbals to keep them perpendicular to the Sun
73
CONCEPT DEVELOPMENT – THERMAL CONTROL OF FUEL TANK
74
Detailed wireframe of the OTV
Cryogenic fuel tank made of Aluminum-Lithium Alloy 60-90 layers of MLI SOFI with thickness of 30.48 cms
CONCEPT DEVELOPMENT – THERMAL CONTROL OF SYSTEMS MODULE AND SOLAR PANELS
75
Two hinged Radiators of size 2m by 0.5m will be used. Temperature on solar panels without thermal control can reach upto 846K Ammonia coolant loop will be used to control temperatures closely.
Hinge Radiator by Swales Aerospace[30] Position of radiator and solar panels on inner truss
KEVIN LOHANLAUNCHING AND DOCKING
DESIGN PROCESS
Launch vehicleOTV launch vehicle must be reliableSingle launch
Universal docking mechanismCarry a variety of payloads
Refueling process development
77
CONCEPT DEVELOPMENT : LAUNCH VEHICLE
78
Launch Vehicle
Payload Capacity
(kg)
Launch Cost
(millions of $)
Falcon XX 140,000 300Falcon X Heavy 125,000 280
SLS 70,000 500Falcon 9 Heavy 53,000 85
Delta IV Heavy 22,560 300
US developers OTV Launch
Vehicle Delta IV
Heavy Custom
fairing Payload and
Fuel Launch Vehicle Falcon 9
Heavy
Heavy Class Launch Vehicles [71],[72],[73],[74]
CONCEPT DEVELOPMENT : DOCKING SYSTEM
79
International docking system standards (IDSS) Regulates where connections are placed Compatibility for future systems
NASA docking system Limited active cycles IDSS compatible
NASA Docking System [75]
Docking System IDSS
Compatibility
Probe and Drogue NO
APAS NO
NASA Docking
System
YES
IDSS compatibility [72] [73]
CONCEPT DEVELOPMENT : REFUELING
80
Robotic Refueling Mission (RRM) Modified Dextre
Arm Successful test in
2013
Robotic Refueling Mission Arm [78]
Refueling System
Flow rate (L/min)
Time to Refuel (hours)
Aerial Refueling 1112 10.5
Gas Pump 37.9 307.8
RRM 1 13725.5
Fluid Flow Rates [60],[61],[62]
Primary Critical Design Issue Fuel Flow Rate
JAY MULAKALARISK ANALYSIS
TECHNOLOGY RISK ANALYSIS
82
TECHNOLOGY CONSEQUENCE POF SHORT CODE
Star Trackers 3 3 [1]
IMU's 3 3 [2]
Magnetometer 1 3 [3]
Sun Sensor 1 4 [4]
Reaction Wheels 2 3 [5]
Magnetic Torque Rods 1 1 [6]
Control Moment Gyroscopes 4 2 [7]
Thrusters 3 4 [8]
Nickel-Hydrogen 4 2 [9]
Lithium-Hydride 3 2 [10]
Solar Panels 3 1 [11]
Chemical Thrusters 3 2 [12]
Nuclear Thermal Thrusters 5 3 [13]
Delta IV 5 2 [14]
NASA Docking System 4 3 [15]
OPERATIONAL RISK ANALYSIS
83
OPERATIONS CONSEQUENCE PO SHORT CODE
Political setbacks due to technologies 4 4 [1]
Use of nuclear power in space 4 4 [2]
Issues with nuclear power across nations 4 3 [3]
Improper disposal of nuclear fuel 5 1 [4]
Accidental failure of nuclear engines 5 2 [5]
Lack of funding to complete production 5 1 [6]
Improper decommissioning of Centurion 4 2 [7]
Launch vehicle failure 2 2 [8]
Inclement weather delaying launch 1 5 [9]
Launch failure 3 2 [10]
Components failing before expected date 1 3 [11]
Sabotage 4 2 [12]
Human error/negligence 3 3 [13]
84
RISK MITIGATION
Technology Risk Analysis
Operational Risk Analysis
Main Factors:• Nuclear Thermal
Propulsion System• Nickel Hydrogen
Batteries
Main Factors:• Political setbacks• Nuclear material
JAY MULAKALAMISSION COSTS
COST ESTIMATION
86
Fixed
Cos
ts • Development
• Production• Assembly
• $2.518 billion Cl
ient
Cos
ts • Fuel Transport
• Fuel Costs
• $94 million
87
COST ESTIMATION
COST ESTIMATION
88
Centurion’s Liquid Hydrogen Fuel• 10 Missions• 10 Falcon 9
Heavys• $850 million
Conventional Bipropellant Fuel• 10 Missions• 50 Falcon 9
Heavys• $4.25 billion
About 5 times more over the course of 10 missions
COST ESTIMATION
89
Our Solution• Delta IV• Nuclear bi-
modal propulsion system
• Modified NASA Docking System
Current Solution• Falcon 9 Heavy• Bi-Propellant
• 3x Fuel• 2x Cost
• NASA Docking System – 2 Cycles
THE CENTURION
90
REFERENCES
REFERENCES
REFERENCES
COST ESTIMATION – FIXED COSTS
94