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.

6

CENTURION

7

CENTURION

THE CENTURION

8

TIMELINE

9

10

TIMELINE

VEHICLE SYSTEMS OVERVIEWJAY MULAKALA

MISSION SYSTEMS

12

• 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

20

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

23

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

24

[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

25

26

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

27

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

28

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

29

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

31

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

33

• Systems module• ADCS sensors• Power and Thermal units

• Propellant tank• Thermal shielding

for cryogenic fuel

• Propulsion systems • Escort System• Radiation shield

SYSTEMS MODULE

34

CRYOGENIC PROPELLANT TANK

35

• Internal volume of 625 cubic meters• Aluminum lithium thermal shielding

PROPULSION SYSTEM

36

STRUCTURAL COMPONENTS

37

CYCOM 5320-1 epoxy resin systemfrom Cytek Inc. [12]

PAMG-XR1 aerospace grade aluminum honeycomb from Plascore Inc. [13]

RADIATOR DESIGN

38

Hinged radiator used in systems module. [14]

Deployable radiator installed in propulsion system. [15]

MASS ESTIMATION

39

STRUCTURAL TESTING

40

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

42

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

43

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

44

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

45

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

51

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.)

58

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

63

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

64

CONCEPT DEVELOPMENT – POWER STORAGE

Requirements for batteries Light Rechargeable High energy density Long discharging time

65


66

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


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

68

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

71

PROPOSED SOLUTIONTwo radiators for Nuclear ReactorsTwo radiators for Systems ModuleCryogenic fuel tank with MLI and SOFI

CONCEPT DEVELOPMENT – THERMAL CONTROL OF NUCLEAR REACTORS

72

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