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AIAA-2004-3982REACTIONN: A Nuclear Electric Propulsion Mission Concept to the Outer Solar System
40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and ExhibitFort Lauderdale, Florida, July 11-14, 2004
Senior Futurist:Mr. A.C. Charania
Director of Advanced Concepts:Dr. Brad St. Germain
Project Engineer:Mr. Jon G. Wallace
President / CEO:Dr. John R. Olds
SpaceWorks Engineering, Inc. (SEI)
With assistance from:
Aerospace Engineer:Tara Polsgrove
NASA Marshall Space Flight Center (MSFC)
Contents
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Overview of Activity
Design Process and Assessment
Summary
Overview of Activity
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Motivation
Respond to request from NASA MSFC’s TD03 Norm Brown and Andy Gamble (In-space Systems Team Lead, Advanced Planning & Concepts Office, Transportation Directorate) to provide an assessment of future uses from NASA’s technology investment in Project Prometheus (nuclear power and propulsion). Develop a concept that would follow a Jupiter Icy Moons Orbiter (JIMO) mission.
Results detailed here include performance analysis and life cycle cost assessment of a final conceptual vehicle design for a Nuclear Electric Propulsion (NEP) mission to Pluto and the Kuiper Belt. Provide a first order design of conceptual system. Results are a collaborative product of SpaceWorks Engineering, Inc. (SEI). Assistance provided by personnel at MSFC with regards to trajectory determination.
Project Purpose
Scope
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Various Post-JIMO Mission Proposed by SEI to NASA MSFC TD03
Semi-permanent L-type stationsNear earth asteroid colonization (manned)Multiple comet sample return (multiple return canisters/pods with one mother ship)Saturn orbiter (Saturn ring sample return)Europan orbiter/lender with power beamingEuropa ocean underwater stationPluto/Neptune/Uranus/Kuiper Belt probe (all four in one mission, "Voyager"-like but with NEP)Optical interplanetary communicationLarge, high power antennas across the solar systemL-point astronomical observatoryLong duration Martian ground or aerial vehicleCometary impactorMissions to the moons of Mars (power beaming)Jupiter atmosphere sample returnVASiMR use of NEPLaser light craft from Martian surface using nuclear power for laser
TWO INITIALLY SUGGESTED CONCEPTS SELECTED FOR MORE DETAILED ANALYSISPLUTO/KUIPER CONCEPT SELECTED FOR FY03 INVESTIGATION
Baseline Concept
Science mission to orbit Pluto and Charon with additional capability to tour Kuiper Belt
Assumption of existing and slightly better JIMO-type technologies (2015+)
Use of Nuclear Electric Propulsion (NEP) consisting of fission reactor and electrostatic ion thrusters
REACTIONN (Rapid Electric Acceleration Coupling ION and Nuclear)
Mission
Timeframe
Concept
Name
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Baseline Concept Schematic 1
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Baseline Concept Schematic 2
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REACTIONN Concept Representation
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Note: Notional representationSource: SpaceWorks Engineering, Inc. (SEI)
Design Process and Assessment
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Vehicle Design Summary
Technology assumptions based upon JIMO reference documentationSource / Heritage
Development of ROSETTA model for rapid evaluation of architectureModeling
Nuclear reactor power generation system:
-Reactor, containment vessel, and cylindrical shielding at front of vehicle
-Radiators for both nuclear reactor and subsystems
Electric propulsion system:
-Electrostatic ion engines
-Thruster platform at end of vehicle
Configuration
Nuclear electric propulsion vehicle
Delta-V of 47.7km/s + 2 km/s for Kuiper Belt excursion
Baseline destination is Pluto with additional mission requirement for Kuiper Belt follow-on mission
Orbit capture at Pluto
Mission
PropertiesItem
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ROSETTA Overview
Reduced Order Simulation for Evaluation of Technologies and Transportation Architectures (ROSETTA)
- A spreadsheet-based meta-model that is a representation of the design process for a specific architecture (ETO, in-space LEO-GEO, HEDS, etc.)
- Each traditional design discipline is represented as a contributing analysis in the Design Structure Matrix (DSM)- Based upon higher fidelity models (i.e. POST, APAS, CONSIZ, etc.) and refined through updates from such models- Executes each architecture simulation in only a few seconds
Requirement for uncertainty analysis through Monte-Carlo simulation
- Model categoriesCategory I: Produces traditional physics-based outputs such as transportation system weight, size, payload, and the NASA metric in-space trip timeCategory II: In addition to above, adds additional ops, cost, and economic analysis outputs such as turn-around-time, LCC, cost/flight, ROI, IRR, and the NASA metric price/lb. of payloadCategory III: In addition to above, adds parametric safety outputs such as catastrophic failure reliability, mission success reliability, and the NASA metric probability of loss of passengers/crew
- Outputs measure progress towards customer goals ($/lb, turn-around-time, safety, etc.)Standard deterministic outputs as well as probabilistic through Monte Carlo
ROSETTA models contain representations of the full design process. Individual developer of each ROSETTA model determines depth and breadth of appropriate contributing analyses.
More assumptions, fewer DSM links than higher fidelity models due to need for faster calculation speeds.
ROSETTA models contain representations of the full design process. Individual developer of each ROSETTA model determines depth and breadth of appropriate contributing analyses.
More assumptions, fewer DSM links than higher fidelity models due to need for faster calculation speeds.
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Vehicle Design Assumptions
Electric Propulsion- Follow-on NSTAR/NEXIS thrusters (assumed 1.2 kg/kW)
Attitude Control System (ACS)- Hydrazine (N2H4) propellant- Delta-V required: 50 m/s forward, 50 m/s aft
Nuclear Power- Particle Bed Reactor (PBR) [7 fuel element configuration]- Assumed power conversion efficiency = 30%- Assumed core density = 1,600 kg/m3
Baseline Vehicle Configuration- Payload Mass = 1 MT- Reactor Power = 1 MW- Isp = 4050 sec. (from NASA NEXT ion engine, max Isp)
Cost determined only for DDT&E and acquisition cost- Based upon historical cost estimating relationships- Includes programmatic wraps
System Test Hardware (STH)Integration, Assembly, & Checkout (IACO)System Test Operations (STO)Ground Support Equipment (GSE)System Engineering & Integration (SE&I)Program Management (PM)
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Placed in 1000 km circular nuclear safeorbit by ETO launch vehicle
NEP System Used For Journey To Pluto
NEP System Used For Propulsive Brake At Pluto
Study Of Kuiper Belt Objects After Completion Of Primary Pluto Science Mission
Initial Mass in Earth orbit = 50.0 MT
EARTH PLUTOTime of Flight = 5.2 years
KUIPER BELT
Mission Profile
Dry Mass With 1MT Payload = 10.8 MT
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Trajectory Curve Fit
0
10
20
30
40
50
60
1.E-05 1.E-04 1.E-03 1.E-02 1.E-01 1.E+00
NEP Spacecraft Initial T/Wo (referenced to Earth g)
Plu
to R
end
ezvo
us
Del
ta-V
(km
/s)
DataCurve Fit
Curve Fit of Trajectory DataInput: Vehicle Power, Initial Mass, IspOutput: Delta-V, Time of Flight (TOF)
Legend
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REACTIONN Baseline Vehicle Summary
0 m
eter
s
50 m
eter
s
100
met
erTotal Length = 115 m
Maximum Width = 101 m
Total Power Required = 1,000 kW
Isp = 4050 sec
IMLEO = 50.03 MT
Dry Mass (with 1MT payload) = 10.8 MT
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Components of NEP on REACTIONN Vehicle
REACTOR
SHIE
LD
ING
POWER CONVERSION
RADIATORS
RADIATORS
ELECTRICTHRUSTERS
POWER PROCESSING UNITS
POWER PROPULSION
SPACECRAFT
SPACECRAFT BUS
VEHICLESYSTEMS
SCIENCEPAYLOAD
POWER MANAGEMENT
AND DISTRIBUTION
XENON PROPELLANT
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Nuclear Power Source: Baseline Vehicle Efficiency Chain
Reactor
100.0%99.5% hother99.5% hcabling99.0% hshielding98.0% Total
Shielding
98.0%99.5% hother99.5% hcabling30.0% hpower-conversion29.7% Total
Power Conversion
29.1%99.5% hother99.5% hcabling95.0% hpower-conditioning94.1% Total
PMAD / Power Cond.
27.4%
PPU
99.5% hother99.5% hcabling95.0% hppu94.1% Total
25.8%
Electric Thrusters
99.5% hother99.5% hcabling79.7% helectric-thrusters78.9% Total
20.3%
Propellant Feed System
99.5% hother99.5% hcabling95.0% hppu94.1% Total
24.2%
Hotel Loads
99.5% hother99.5% hcabling99.0% Total
27.1%
Science Loads
99.5% hother99.5% hcabling99.0% Total
27.1%
Communication Loads
99.5% hother99.5% hcabling99.0% Total
27.1%
The efficiency of converting electric power to thrust power (thruster efficiency), based upon xenon propellant
79.7%η-electric-thrusters
For both nuclear and solar power systems99.5%η-cabling
99.0%η-shielding
The efficiency of power conversion for the reactor30.0%η-power-conversion
The efficiency of power conditioning for the reactor95.0%η-power-conditioning
95.0%
95.0%
99.5%
Value
η-propellant-feed-system
η-ppu
Including radiation and thermal, for both nuclear and solar power systems
η-other
DescriptionEfficiency
EFFICIENCIES FOR NUCLEAR POWER SOURCE AND ELECTRIC PROPULSION
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Baseline Vehicle Mass Breakdown Statement (MBS) and Power Budget
1.0 Nuclear reactor power system
Nuclear core
Containment vessel
Radiation shield
Power conversion
Power conditioning
2.0 PropulsionElectric propulsion system
Attitude control system
3.0 Thermal ControlPrimary radiators
Secondary radiators
Misc blankets, heaters, thermostats
4.0 Primary Central Structure
5.0 Data ProcessingAttitude/Orbit determination
Attitude/Orbit control
Device pointing
Integrated function
6.0 Navigation Sensing/ControlCelestial
IMU
7.0 Telecom TCM Module
Command and data handling
Communications payload
8.0 Growth Margin (15%)
Dry Mass (w/o payload)
9.0 Payload
Dry Mass (w payload)
10.0 Propellants NEP propellant
Forward attitude control
Aft attitude control
Near Earth Departure Mass
Mass [kg]Mass Item
Two-Level MBS
4,680300
2,370
1,175
200
635
2,7402,700
40
10560
20
25
685
7020
20
20
10
4020
20
20025
10
165
1,280
9,800
1,000
10,800
39,23036,700
1,265
1,265
50,030
Communication Loads
Science Loads
Hotel Loads
Propellant feed systems
Power required for electric thrusters
PPU
PMAD / Power Conditioning
Power conversion losses
Shielding losses
Total cabling losses
Total other losses
Total power required from reactor
Power [kW]
5.0
25.0
5.0
2.1
207.0
13.6
14.4
679.2
10.0
19.3
19.4
1,000.0
Power Item
Two-Level Power Budget
50.03 MT50.03 MT
1.0 MW1.0 MW
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Trade Study A: Isp Versus Near Earth Departure Mass (1 MT Payload)
Payload = 1.0 MT
20
30
40
50
60
70
80
90
100
110
120
130
140
150
160
3,500 3,750 4,000 4,250 4,500 4,750 5,000
Isp, seconds
Nea
r E
arth
Dep
artu
re M
ass,
MT
TOF = 3 yrs
TOF = 5 yrs
TOF = 7 yrs
TOF = 9 yrs
TOF = 11 yrs
TOF = 13 yrs
TOF = 15 yrs
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Trade Study B: Isp Versus Reactor Power (1 MT Payload)
Payload = 1.0 MT
0.2
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
2.0
2.2
2.4
2.6
2.8
3,500 3,750 4,000 4,250 4,500 4,750 5,000
Isp, seconds
Rea
ctor
Pow
er, M
W
TOF = 3 yrs
TOF = 5 yrs
TOF = 7 yrs
TOF = 9 yrs
TOF = 11 yrs
TOF = 13 yrs
TOF = 15 yrs
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Trade Study C: Payload Versus Near Earth Departure Mass (Isp = 4,050 seconds)
Isp = 4,050 seconds
40
45
50
55
60
65
70
75
80
85
90
95
0.25 0.50 0.75 1.00 1.25 1.50 1.75 2.00
Payload, MT
Nea
r E
arth
Dep
artu
re M
ass,
MT
Reactor Power = 0.50 MW
Reactor Power = 0.55 MW
Reactor Power = 0.65 MW
Reactor Power = 0.75 MW
Reactor Power = 1.00 MW
Reactor Power = 1.25 MW
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Baseline Vehicle Cost Breakdown:Non-Recurring (DDT&E) Cost (without wraps and margin)
† - FY2003 US$
Nuclear reactor power system59.0%
Propulsion27.1%
Main structure5.7%
Thermal Control2.4%
Data Processing0.7%
Navigation Sensing/Control0.2%
Telecom and Data4.9%
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Baseline Vehicle Cost Breakdown
Item Non-Recurring (DDT&E) Cost,
$M-FY2003
Acquisition Cost,
$M-FY2003 Hardware Cost
Nuclear reactor power system $725.00 M $115.00 M Propulsion $333.52 M $58.56 M
Thermal Control $30.00 M $0.52 M Main structure $70.00 M $0.10 M
Data Processing $8.40 M $4.20 M Navigation Sensing/Control $2.00 M $0.60 M
Telecom and Data $60.49 M $21.90 M Cost Summary
Sub-total $1,229.41 M $200.87 M Total Programmatic Costs (30%, and 10%) $368.82 M $20.09 M
Total Cost (without Margin) $1,598.24 M $220.96 M Margin (+55%) $879.03 M $121.53 M
Total Cost (with margin) $2,477.27 M $342.49 M Total Cost-Development and Acquisition (with margin) $2,819.76 M
Item Non-Recurring (DDT&E) Cost,
$M-FY2003
Acquisition Cost,
$M-FY2003 Hardware Cost
All sub-systems $1,213.61 M $339.63 M Cost Summary
Sub-total $1,213.61 M $339.63 M Systems Integration $376.15 M $33.83 M
Fee (+5%) $79.49 M $12.79 M Program Support (+10%) $166.92 M $26.85 M
Contingency (+15%) $275.45 M $44.30 M Total Cost (with margin) $2,111.59 M $339.63 M
Total Cost-Development and Acquisition (with margin) $2,451.22 M
REACTIONN Baseline Spacecraft Cost Assessment: NAFCOM 2004 Cost Model
REACTIONN Baseline Spacecraft Cost Assessment: ROSETTA Cost Model
$2.82 B$2.82 B
$2.45 B$2.45 B
Total Cost EstimateDoes not include
technology maturation cost, science instrument
cost, or launch vehicle/in-space assembly costs
REACTIONN Concept Visualization
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Note: Notional representationSource: SpaceWorks Engineering, Inc. (SEI)
Summary
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Summary
Developed a first level conceptual design of a NEP architecture to Pluto/Charon and Kuiper Belt- Near Earth Departure Mass = 50.03 MT, 10.8 MT (dry)- Total ∆V = 49.7 km/s (Pluto/Charon/Kuiper Belt)- 1 MT of payload at a reactor power level of 1 MW- Cost for development and acquisition = $2.45-$2.82 B- Spacecraft Operations Cost Model (SOCM): $106.4 M (FY2003) which consists of $77.7 M for flight operations, $13.6 M for
navigation and tracking, and $15.1 M for science operations
Development of ROSETTA model to encompass most important engineering and economic disciplines- Integrates trajectory, performance, weights, power, sizing, and cost disciplines
Generally vehicle is large and will require in-space assembly of constituent parts- Subsystems are generally small enough to be launched individually or in combination with other subsystems
Trade studies indicate that for lower payload classes (under 1 MT), larger reactor power does not necessarily relate to smaller IMLEO, at this point of lower payloads the power reactor seems to be oversized for the payload required
- Effect most noticeable for power levels approaching 1 MW and beyond for the payload range (0.25-2MT) in question
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SpaceWorks Engineering, Inc. (SEI)
Contact Information Business Address:SpaceWorks Engineering, Inc. (SEI)1200 Ashwood ParkwaySuite 506Atlanta, GA 30338 U.S.A.
Phone: 770-379-8000Fax: 770-379-8001
Internet:WWW: www.sei.aeroE-mail: [email protected]
President / CEO: Dr. John R. OldsPhone: 770-379-8002E-mail: [email protected]
Director of Hypersonics: Dr. John E. BradfordPhone: 770-379-8007E-mail: [email protected]
Director of Advanced Concepts: Dr. Brad St. GermainPhone: 770-379-8010E-mail: [email protected]
Project Engineer: Mr. Matthew GrahamPhone: 770-379-8009E-mail: [email protected]
Project Engineer: Mr. Jon WallacePhone: 770-379-8008E-mail: [email protected]
Senior Futurist: Mr. A.C. CharaniaPhone: 770-379-8006E-mail: [email protected]