spaceworks engineering, inc. (sei)science mission to orbit pluto and charon with additional...

Copyright 2004 SpaceWorks Engineering, Inc. (SEI) All rights reserved.Engineering Today, Enabling Tomorrow

www.sei.aeroSpaceWorks Engineering, Inc. (SEI)

SpaceWorks Engineering, Inc. (SEI)

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


With assistance from:

Aerospace Engineer:Tara Polsgrove

NASA Marshall Space Flight Center (MSFC)

Contents



Overview of Activity

Design Process and Assessment

Summary

Overview of Activity



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





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



Baseline Concept Schematic 1



Baseline Concept Schematic 2





REACTIONN Concept Representation

Copyright 2004 SpaceWorks Engineering, Inc. (SEI) All rights reserved.Engineering Today, Enabling Tomorrow Page 9


Note: Notional representationSource: SpaceWorks Engineering, Inc. (SEI)

Design Process and Assessment





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



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.



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)



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



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



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



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



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


27.1%

Communication Loads


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



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



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



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



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








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%



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



Note: Notional representationSource: SpaceWorks Engineering, Inc. (SEI)

Summary





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




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]

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