Lecture Series on SPACE NUCLEAR POWER & PROPULSION SYSTEMS -3- Space Fission Power & Nuclear Electric Propulsion Systems (last updated in January 2020) Eric PROUST

(1)

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Lecture Series on SPACE NUCLEAR POWER &

PROPULSION SYSTEMS -3- Space Fission Power &

Nuclear Electric Propulsion Systems (last updated in

January 2020) Eric PROUST

Eric Proust

To cite this version:

Eric Proust. Lecture Series on SPACE NUCLEAR POWER & PROPULSION SYSTEMS -3- Space

Fission Power & Nuclear Electric Propulsion Systems (last updated in January 2020) Eric PROUST.

Engineering school. France. 2020. �hal-02979127�

(2)

Eric PROUST Lecture Series on SPACE NUCLEAR POWER & PROPULSION SYSTEMS -3- Space Fission Power & Nuclear Electric Propulsion Systems(last updated in January 2020)

FROM RESEARCH TO INDUSTRY

Atomic Energy and Alternative Energies Commission -

www.cea.fr

LECTURE SERIES ON

SPACE NUCLEAR POWER & PROPULSION SYSTEMS

-3- Space Fission Power & Electric Propulsion

Eric PROUST

Last updated in January 2020

Space Nuclear Power & Propulsion in January 2020’s news

Nuclear Thermal Propulsion

Nuclear Electric Propulsion

Space Nuclear Power Reactor

Radioisotopic Thermoelectric

Space Generators

(3)

190 MeV

Product Nuclei

(KE: 168 MeV)

Neutron

Fissile Nucleus

(U-235)

Neutrons

(2,5)

U-235

Space Nuclear Power Systems: Radioisotopic, or Fission-based?

3

Energy released by the radioactive decay

(alpha) of a radioisotope

Applications:

• Thermal management: RHU

• Power generation: RTG, DIPS

Energy released by the neutron-induced fission

of a fissile nuclide

Applications:

• Power generation

, for supplying

• Observation instruments (history: radar)

• A moon/mars base

• Electric thrusters (

Nuclear Electric Propulsion: NEP

)

• Direct propulsion (by heating a propellant gas)

• Nuclear Thermal Propulsion (NTP)

• Both combined

5.5 MeV

U-234

Pu-238

α (He-4)

3

Today’s lecture

Lecture Outline

4

The early space age years: space fission power systems developed for defense missions in earth orbit

-

reactor and energy conversion technologies:

ROMASHKA, SNAP-10A, BUK/BES-5, TOPAZ I, TOPAZ II, SNAP 50/SPUR

Safety principles and key design challenges

-

Safety principles built on the feedback of experience (including accidents)

-

Mass, shadow shield, reactor, radiator, and reliability

-

Example of later implementation: the SP-100 design

Nuclear electric propulsion for space exploration

-

An illustration: the JIMO project

-

Space propulsion basics: specific impulse and propellant mass, thrust

-

Electric thrusters: high specific impulse propulsion engines

-

Space propulsion basics: thermal vs electric propulsion

-

Solar vs nuclear electric propulsion; Manned mission to Mars

Current developments

(4)

Lecture Outline

5

The early space age years: space fission power systems developed for defense missions in earth orbit

-

reactor and energy conversion technologies:

ROMASHKA, SNAP-10A, BUK/BES-5, TOPAZ I, TOPAZ II, SNAP 50/SPUR

Safety principles and key design challenges

-

Safety principles built on the feedback of experience (including accidents)

-

Mass, shadow shield, reactor, radiator, and reliability

-

Example of later implementation: the SP-100 design

Nuclear electric propulsion for space exploration

-

An illustration: the JIMO project

-

Space propulsion basics: specific impulse and propellant mass, thrust

-

Electric thrusters: high specific impulse propulsion engines

-

Space propulsion basics: thermal vs electric propulsion

-

Solar vs nuclear electric propulsion; Manned mission to Mars

Current developments

-

NASA’s KILOPOWER, R&D at Keldysh, MEGAHIT/Democritos

ROMASHKA: first ground prototype of Nuclear Fission Space Power System (USSR)

Реактор "Ромашка" (Daisy)

(ground-tested 1964-1966, 15 000 h, 6 100 kWh)

No coolant: radial conductive heat transfer

~0.5 kWe

~28 kWth

455 kg

(reactor: 265 kg)

49 kg U5 (90%)

Active core ф: 240 mm

Reflector out ф : 480 mm

Radiating fins

(Cu, ~550°C)

Thermoelectric

elements (Si-Ge)

<815  <585°C

Radial reflector

(Graphite + Be)

Axial reflector

(Graphite + Be)

Control rod

(B

₄

C)

Active core:

UC

₂

discs

encased in

graphite

UC

₂

discs in Graphite casing

Source: Ponomarev-Stepnoi, N.N., Kukharkin, N.E. & Usov, V.A. “Romashka” reactor-converter. At Energy 88, 178–183 (2000). https://doi.org/10.1007/BF02673156 6

(5)

SNAP 10A : ~500 We Nuclear Fission Space Thermoelectric Power System (USA)

SNAP 10A: the first (and only) fission reactor launched by the US (in 1965)

Thermoelectric conversion (SiGe)

NaK coolant

UZrH fuel (95% U

₅

)

7

operated for 43 days during its flight test;

prematurely shut down by a faulty command receiver

SNAP 10A : ~500 We Nuclear Space Thermoelectric Power System (USA, 1959-1965)

UZrHx

fuel rods in

internally coated

Hastelloy N clads

ф 1.25 x L 12,45 in.

95% U5

Source: SNAP Nuclear Space Reactors, Library of Congress Catalog Card Number 66-62772

Electrical Power (W)

580 We

Reactor Power (kW)

43 kWth

System overall efficiency

1.3%

Specific weight

750 kg/kWe

Design life

1 year

Overall length x diameter

3.5 x 1.3 m

2

NaK reactor out/ inlet T

833 / 761 K

Radiator area

6.8 m

2

SNAP 10A design parameters

SiGe

TE

(6)

Thermoelectric conversion

SNAP 10A

T

_H

: 777 K av.

T

_C

: 610 K av.

ZT (SiGe): 0.4 av.



ε ~ 2.0%

(

ZT = 1.0

_

ε ~ 4.0%

)

9 Eric PROUST Lecture Series on SPACE NUCLEAR POWER & PROPULSION SYSTEMS -3- Space Fission Power & Nuclear Electric Propulsion Systems(last updated in January 2020)

BUK/BES-5: 3 kWe Nuclear Space Thermoelectric Power System (USSR, 1960-1988)

35 cm

Reactor

NaK duct

Shadow shield

Radiator

Compensation

tank

Thermoelectric

generator

scale:

50 cm

31 reactors operated from 1970 to 1988

on Radar Ocean Reconnaissance Satellites

(RORSAT) on ~280 km 65° incl. orbits

average operating lifetime: 50 days

(max: 135 days)

Electric power

3 kWe

Thermal power

100 kWth

System efficiency

3.9%

System specific mass

310 kg/kWe

System mass, incl.

930 kg

Reactor

53 kg

Shadow shield

350 kg

Energy conversion

Si-Ge thermoelectric

Coolant

NaK-78

Core outlet temperature

973 K

Fuel

(SS clad)

U~7%Mo

U

5

enrichment /loading

~90% / 30 kg

form

37 (2 cm ф) rods

Fissile zone c-to-c x h

15 x 15 cm

2

Reflector

Beryllium

axial thickness

10 cm

radial, external radius

17.5 cm

Reactivity control

sliding Be drums

10 (½ SNAP-10)

(7)

Thermionic Energy Conversion

exp )

exp

)

If ΦE

> ΦE

+ V e, w/o space charge effects

For 1 amp/cm

2

ΦE

= 4.5 (W)

=> T

_E

= 2 600 K

ΦE

= 2.0 (W+Cs) => T

_E

= 1 250 K

ΦE

Electrons escaping

from the surface

Emitter Temperature

P o w e r co n v e rs io n e ff ic ie n cy

E

m

it

ti

n

g

E

le

m

en

t

Element Work Function Φ (eV)

11

TOPAZ-I ~5 kWe Space Nuclear In-Core Thermionic Power System (USRR, 1965-1988)

TOPAZ:

Thermoionic

Experiment with Conversion in Active Zone

First tests on Thermoionic conversion in 1961 in USSR (1957 in the US)

TOPAZ program launched in 1965, first TOPAZ prototype operational in 1970

Ground-tests of flight-system prototypes from 1982 to 1984

2 reactors flight-tested in 1987, operated for 143 and 342 days on 800 km orbit

5-cell Thermoionic Fuel Element

5-6 kWe

110-150 kWth

~4.3% efficiency

7 m

2

_radiator

320 kg (reactor)

1200 kg (total)

200 kg/kWe

12 (1865 K) (530 – 610 °C) 14.6 cm

Core: 28 cm D x 36.4 H

79 TFE, 12 kg

235

_U

Reactor diameter: 46 cm

(2.5 kg BOM) MONOCRYSTALLINE ZrH1.8

(8)

TOPAZ II ~5 kWe Space Nuclear In-Core Thermionic Power System (USRR)

Mo3%Nb monocrystal + W

184

_{coating (1 950 K peak T) Polycrystal Mo (~850 K)}

4.5 – 5.5 kWe

115-135 kWth

~4% efficiency

210 kg/kWe

1061 kg total

290/390 kg reactor/shield

50 kg (7.2 m

2

_{) radiator}

3.9 m height

NaK coolant BOL inlet/outlet T: 743/843 K

UO

₂

(96% U

5

_{): ~27 kg}

13

Ground tested

never flight-tested

Towards higher efficiency: Thermodynamic cycles: Rankine and Brayton

He-Xe Brayton cycle

Potassium RANKINE cycle

14 Source: Lee S. Mason, Power Technology Options for

(9)

SNAP 50/SPUR 300 kWe (US, 1959-1972) : Potassium Rankine energy conversion

Program terminated before a

complete system was ground-tested

UN/NbZr/Li reactor 2.2 MWth

Li T out/in T : 1360 / 1310 K

K/SS Rankine conversion

Turbine in: 1280 K/ 770 kPa

Turbine out: 880 K / 21 kPa

NaK/SS main radiator

NaK T in/out : 950 / 860 K

System efficiency: ~13.5%

Specific weight: 15.9 kg/kWe

15

NB: micro/zero gravity fluid mechanics

Source: SNAP-50/SPUR Program Summary, CNLM 5889, 09/24/1964

Lecture Outline

16

The early space age years: space fission power systems developed for defense missions in earth orbit

-

reactor and energy conversion technologies:

ROMASHKA, SNAP-10A, BUK/BES-5, TOPAZ I, TOPAZ II, SNAP 50/SPUR

Safety principles and key design challenges

-

Safety principles built on the feedback of experience (including accidents)

-

Mass, shadow shield, reactor, radiator, and reliability

-

Example of later implementation: the SP-100 design

Nuclear electric propulsion for space exploration

-

An illustration: the JIMO project

-

Space propulsion basics: specific impulse and propellant mass, thrust

-

Electric thrusters: high specific impulse propulsion engines

-

Space propulsion basics: thermal vs electric propulsion

-

Solar vs nuclear electric propulsion; Manned mission to Mars

Current developments

(10)

Space Nuclear Power Reactors:

Safety

Principles

Use only fresh Uranium as fuel (reactor launched free of

fission products); Use of Plutonium precluded

Reactor designed to prevent accidental criticality whatever

the emergency situation (in case of reactor compaction

and/or flooding upon impact following launch abort, …)

First criticality and operation started only once prescribed

‘’sufficiently high orbit’’ reached (“nuclear safe” orbit,

allowing for sufficient FP radioactive decay before reentry)

Minimize

fission product release

(principle ALARA)

Reactor designed either survive accidental reentry or

to disintegrate upon reentry and disperse its residual

radioactivity in the upper atmosphere (soviet strategy

adopted in the latest RORSATs, Cf. Kosmos 1402)

17

Safety objectives and regulations are currently established on

the basis of national political/legal rules: USA, Russia (Europe?)

List of

internationally agreed upon principles

(UN

Committee on the Peaceful Uses of Outer Space, 1992)

but no specified safety criteria or regulations so far

“Space Nuclear Safety Cuture” inspired from the

experience learned from past “nuclear launches”

Principles relevant to the use of Nuclear Power Sources in Outer Space. Report of the Committee on the Peaceful Uses of Outer Space, General Assembly Official Records Forty-seventh Session. Supplement No. 20(A/47/4/20)

A “Nuclear Safe Orbit”?

Time after shutdown (Days)

To

ta

l A

ct

iv

it

y

(C

u

ri

e

s)

Initial orbital altitude (km)

O

rb

it

a

l l

if

e

ti

m

e

(d

a

ys

)

A “Nuclear Safe Orbit”:

a (typically 1000 km or higher) orbit providing an

unattended orbital life of sufficient lifetime (typically

10 000 y or more) so that the core’s radioactive nuclide

inventory will have decayed down to ‘’acceptable’’ levels

Typical 300 kWe SNPS core activity decay

(11)

Kosmos 954 (BES-5) accidental reentry and operation Morning Light (1978)

19 Source: Operation Morning Light – An Operational History, CFD Edmonton Report, Mulroney Institute, 2018

Subsequent Kosmos 1402 (BES-5) accidental reentry and Kosmos 1900 loss of contact

Kosmos 954 (1978): failure of booster mechanism designed to bring reactor to “nuclear safe disposal orbit”

: reactor not designed for specific behavior during accidental reentry

Change made: reactor designed to ensure, in case of accidental reentry, it disintegrates and disperses its radioactive

in the upper atmosphere (mechanism to eject the reactor core)

Kosmos 1402 (1983) seemed to show that it works: after failure of the booster mechanism, the core was

ejected, reentered the stratosphere over the South Atlantic Ocean and it is believed to have completely burned up

into particles and dispersed to safe levels of atmospheric radioactivity

Change made: automatic back-up booster system (‘’rugged booster’’)

Kosmos 1900 (1988): following failure of the primary booster, the back-up booster succeeded in bringing the core

close to its prescribed disposal orbit

20 Source: Soviet Space Nuclear Reactor Incidents: Perception Versus Reality, Gary L. Bennett, Space Nuclear Power Systems 1989, Chapter 25, pp 273-278

(12)

Space power systems: minimizing mass is key for minimizing launch costs

Generator types

Running

time

Power range

Specific Mass

Fuel Cell

A few

hundreds of

hours

A few tens of kWe

~ 15 kg/kWe

Solar photovoltaic

10 years

A few tens of kWe

100 à 200 kg/kWe

(with batteries)

Solar Dynamic

7 years

20 – 100 kWe

150 à 300 kg/kWe

Radioisotope systems

(

238

_{Pu) : RTG, DISP}

A few tens

of years

Up to 0.5 kWe (RTG)

A few kWe (DIPS)

~ 200 kg/kWe (RTG)

~ 100 kg/kWe (DISP)

Nuclear Reactors

10 years

10 kWe – 1 MWe

~30 kg/kWe (200 kWe)

~100 kg/kWe (20 kWe)

The key parameter:

the Specific Mass (kg/kW

_e

)

Launch costs: access to ISS ~ US$25 000/kg (SpaceX Falcon 9 Plus Dragon, 6 t payload)

access to LEO: strongly declining costs since entry of SpaceX

21 Source: redrawn from Lee S. Mason, Power Technology Options for

Nuclear Electric Propulsion, IECEC 2002, paper n° 20159

Fission Power systems

Minimizing Specific Mass does not imply Maximizing Conversion Efficiency

22

• Minimizing System Specific Mass is a Key Objective

• Maximizing System Efficiency IS NOT an Objective (it may be one for Radioisotope Power Systems)

• System efficiency depends on Conversion Technology (and Peak Temperature)

• System Specific Mass depends much less on Conversion Technology (but significantly on Peak Temperature)

Source: Lee S. Mason, Power Technology Options for Nuclear Electric Propulsion, IECEC 2002, paper n° 20159

Illustration: ThermoElectric and Free-Piston Stirling versions of SP-100* exhibit relatively comparable specific masses

(13)

Minimizing mass (and size): the

radiator

(getting rid of waste heat)

Area for radiating 1 kW ( = 1)

!

1 #

# $

%

#

&

'()*+,)

.

1 /

$

_.

%

₎

1 000 K

0.018 m

2

400 K

0.68 m

2

550 K

0.19 m

2

700 K

0.07 m

2

minimum when

3 4

2 .

# 25%

(thermodynamic cycles)

Maximizing energy efficiency is not the goal!

High power systems

(a few 100 kWe)

radiator size constraints

high temperature reactors

fast spectrum reactors

(refractory materials

are poor moderators)

0 5 10 15 20 25 30 35 40 45 50 300 350 400 450 500 550 600

Com pre ssor Inlet Tem perature (K)

C o n v e rs io n E ff ic ie n c y ( % ) 0 20 40 60 80 100 120 140 160 R a d ia to r S u rf a c e ( m 2 )

Brayton cycle

T

_H

constant

Getting rid of waste heat: the

radiator

Reliability

: segmentation, heat pipes, protection

against micrometeorites

Deployability

for high power systems (fit inside

launcher’s fairing)

(14)

Minimizing mass: the

shadow shield

(shielding electronics/payload from radiations)

reactor

shield

protected

cone

Hardened electronic Payload Man

reactor

Li

6

_H

W

_Shield

A major weight contributor (>25% of the unmanned system mass, higher for manned missions)

Alternate dense (absorbing

γ

) and light materials

(slowing down and absorbing

n

)

Li

6

_H

Density: 0.82 g/cm

3

H concertation ~ H

₂

O

!Melting point: 680°C!

Li

6

_{: 945 barns}

th. n absorption XS

with no secondary

γ

Sliding radial reflector

segments / drums

SP-100

BES-5

Minimizing mass: minimizing the

reactor

height

Reactor to be optimized to minimize the shadow shield diameter (=mass)  minimize reactor height

Owing to reactivity control through control of radial neutron leakage

(radial reflector of variable efficiency)

Radial reflector ‘’shutters’’

Rotating control drums

equipped

with neutron absorbing sectors

SNAP 50

TOPAZ-II

(Narciss exp, KI)

OPUS

Source: Reactivity control options of space nuclear reactors, Timothy M. Schriener, Mohamed S. El-Genk, doi:10.1016/j.pnucene.2008.11.003

26

(and diameter)

(15)

Minimizing Mass: minimizing the

Reactor Mass

by using the highest Uranium Enrichment

27

Core Parameters

93% U5

20% U5

Power (kWth)

600

500 Fuel Pin OD (mm)

15.9

25.4 No. of Fuel Pins

211

200 Core Height (cm)

40.6

74.9 Core Diameter (cm)

22.9

37.6 Fuel Form

UZrH

1.8

UZrH

1.8

Uranium (wt. %)

10.35 10.35-18.0

Relative Core Volume

1.0

5.0 Moderated Reactor Design: SNAP-8 (600 kWth)

The lowest reactor/system mass is always obtained by using HEU

Possible issues: availability/constraints of HEU or non-proliferation policy

The reactor (and radiation shield) mass penalty of using Uranium enriched

to significantly less than HEU depends:

• on whether the reactor operates with a fast or thermal neutron spectrum

• on the reactor design concept

• on the reactor thermal power

x 4.2

x 12.6

x 1.6

x 2.3

Source: Impact of the Use of Low or Medium Enriched Uranium on the Masses of Space Nuclear Reactor Power Systems, DOE/NE-0112, December 1994

System Parameters 93% U5 35% U5 20% U5

Reactor Thermal Power (kWth) 522 604 668

Reactor Height (cm) 111.5 111.5 111.5

Reactor Diameter (cm) 55.4 60.7 80.2

No. of TFE - Emitter D (cm) 48 - 1.8 36 - 1.8 90 - 2.8

No. of ZrH1.8 pins 156 72 180

UO2 Mass / U235 Mass (kg) 37.6/30.7 69.3/21.4 173/30.5

MASS SUMMARY (kg) Reactor Subsystem 610 738 1677 Reactor Controls 50 110 154 Reentry Shield 35 45 96 Radiation Shield 452 526 869 Heat Rejection 282 293 389

Natura Threat Protection 80 92 133

Power CC&D 392 392 392

Boom/Structure 121 140 237

TOTAL 2020 2336 3947

Mass Increase Baseline 16% 95%

Moderated Reactor Design: 20 kWe S-PRIME in-core Thermionics

Minimizing mass: minimizing the

reactor

core diameter

Reactor to be optimized to minimize its mass and the shadow shield diameter  minimize reactor diameter

Example: NaK-cooled UO

₂

/ZrH core

Smallest cores: with well thermalized spectra  good moderators (ZrH, Be, …)

Good moderators not compatible with high temperature

 High power reactors (requiring high temperature for system efficiency) = fast spectrum cores

C

ri

ti

ca

l D

ia

m

e

te

r

(c

m

)

Vol. ZrH / Vol. UO2

fresh fuel

critical after 10 y

Be reflector

Energy (MeV)

N

o

rm

a

lis

e

d

F

lu

x

(n

/c

m

2

.s

)

ZrH/UO2 = 0

ZrH/UO2 = 1 (F1/F2 = 11.5)

ZrH/UO2 = 3

ZrH/UO2 = 8

ZrH/UO2 = 18 (F1/F2 = 2.2)

28

(and height)

(16)

What about

Reliability

?

High-cost long-duration missions without maintenance

Avoid single points of failure when possible

 Provide redundancy trough

• Segmentation of cooling loops

or use of heat pipes instead of loops

• Segmentation of energy conversion

Prefer lowest failure probability equipment

Avoid equipment with moving mechanical parts (wear, fatigue) when possible:

 electromagnetic pumps

 static energy conversion: thermoelectics, thermionics

If not possible (dynamic conversion), adopt quasi frictionless devices

 Brayton turbomachinery with gas foil bearing

 Free piston Stirling engines

See the example of the SP-100 design

(on next slides)

29

The SP-100 (100 kWe) Program (US, 1983-1994)

Joint DOD/NASA/DOE program

100 kWe system

scalable from 10s to 100s of kWe

7 years at full power

10 years overall operation

with 95% reliability

In earth orbit (DOD) or exploration

Key technology choices

99.9%

7

_{Li cooled reactor}

_{(1 350 K outlet)}

Nb(/Re) clad

UN

(97%

5

_{U) fuel pins}

Li containing structures: Nb alloy PWC-11

Thermoelectric

conversion (

SiGE/GaP

)

C/C matrix structure + Ti/K heat-pipe radiator

Test scheduled for early 90’s. Program restructured due to funding

restrictions to demonstrate a complete technology and lifetime

test by 1998. Program terminated in FY95

US$ 520 millions (as spent over 83-94)

(17)

The SP-100 (100 kWe) Program (US, 1983-1992)

Reactor

700 Shield

1000

Primary heat transport

500 Reactor I&C

290 Power conversion

370 Heat rejection

850 Power cond. Control

390 Mechanical/structure

480 Total

4580

(45.8 kg/kWe)

SP-100 Mass breakdown

Rated Power

100 kWe

Thermal Power

2 400 kWth

Reactor outlet T

1 350 K

inlet T

1 300 K

Heat rejection in. T

860 K

out. T

780 K

Radiator area

104 m2

Key SP-100 parameters

Key design drivers:

Safety

Reliability

Mass

(4% system efficiency)

31 Source: SP-100 Nuclear Space Power Systems With Application to Space Commercialization, NASA TM 101403, 1989

The SP-100 Program (US, 1983-1992)

Reactor

Vessel

Radial Sliding

Reflectors (12)

Fuel Pins (858)

Safety Rods (3)

Honeycomb

Coolant Inlet

Flow Passage

Auxiliary (NaK)

Coolant & Thawing

Loop U-tubes (52)

Reactor

Vessel

Protects against LOCA

Prevents fuel pins

spreading upon impact

Assures intact reentry

Assure shutdown in

accident situations

TEM pump

Primary & secondary HTS (12 loops)

Lithium/Helium gas separator/accumulator

SP-100 core cross-section

32 Source: Scott F. Demuth, SP100 Space Reactor Design,

Progress in Nuclear Energy, No.3, pp. 323-359, 2003 Safety rod latches

Retains rods in core for impact and explosion accidents

Reflector launch locks

Safety features

TEM pumps: decay heat removal

w/o external power

(18)

The SP-100 Program (US, 1983-1992)

Envisaged SP-100 ground testing facility

The challenge (cost) of ground testing in a simulated space (vacuum) environment

33 Prevention of Significant Deterioration Application for Approval to Construct SP-100 Ground Engineering System Test Site DOE/RL-90-14, 1990

The ERATO 20/200 kWe Program (France, 1982-1989)

Partnership CNES – CEA with industrial

participation (SAGEM, TURBOMECA,

NOVATOME…)

Design features and technology bases of

20/200 kWe Nuclear Power System for

electric

propulsion

& comparison with conventional

space propulsion means

Conceptual design studies completed

Reactor definition, system integration,

development time and cost estimation

Studies on 3 systems (all with

Brayton cycle

)

UO

2

/ sodium / S-Steel T < 700 °C

UC

2

/ He-Xe / super alloys T < 850 °C

UN / Lithium / Mo-R T < 1150 °C

Range of power considered: 20-200 kWe

Radiator: 32 m² to 140 m²

Mass balance (reactor + shield + conversion

system + electrical network):

~ 2000/7000 kg



100 kg/kWe at 20kWe



35 kg/kWe at 200kWe

(19)

Space Nuclear Fission Power & Nuclear Electric Propulsion

35

The early space age years: nuclear fission power systems in earth orbit for defense missions

-

reactor and energy conversion technologies:

ROMASHKA, SNAP-10A, BUK/BES-5, TOPAZ I, TOPAZ II, SNAP 50/SPUR

Safety principles and key design challenges

-

Safety principles built on the feedback of experience (including accidents)

-

Mass, shadow shield, reactor, radiator, and reliability

-

Example of later implementation: the SP-100 design

Nuclear electric propulsion for space exploration

-

An illustration: the JIMO project

-

Space propulsion basics: specific impulse and propellant mass, thrust

-

Electric thrusters: high specific impulse propulsion engines

-

Space propulsion basics: thermal vs electric propulsion

-

Solar vs nuclear electric propulsion; Manned mission to Mars

Current developments

-

NASA’s KILOPOWER, R&D at Keldysh, MEGAHIT/Democritos

Nuclear Electric Propulsion for Space Exploration: example of the JIMO project

JIMO, the

Jupiter

Icy Moons Orbiter (NASA project, 2003-2005)

• Main target: Europa, where an ocean of liquid water may harbor alien life.

• Other targets of interest: Ganymede and Callisto

, which are thought to have

liquid, salty oceans beneath their icy surfaces

Using electric propulsion (8 ion engines, plus Hall thrusters of varying sizes) to

go into and leave orbits around the moons of Jupiter, creating more thorough

observation and mapping windows than for current spacecraft, which must

make short fly-by maneuvers because of limited fuel for maneuvering

Once on Jupiter moon orbit, the nuclear power system would provide

electricity to feed the high power radar needed to penetrate beyond the thick

icy surface

Reactor design conducted by the naval reactors branch of DOE

(20)

Eric PROUST

Lecture Series on SPACE NUCLEAR POWER & PROPULSION SYSTEMS -3- Space Fission Power & Nuclear Electric Propulsion Systems(last updated in January 2020) 37

Nuclear Electric Propulsion for Space Exploration: example of the JIMO project

• Gross mass in low Earth orbit: 36 500 kg

• Mass of xenon propellant: 12 000 kg

• Reactor module mass: 6 700 (+ 3 300) kg (200 kWe output)

• Spacecraft module dry mass: 16 200 kg

• Science payload mass: 1 500 kg

• Electric turboalternators: multiple 104 kW

• Deployable radiator: 422 m² surface area

• Electric Herakles ion thrusters:

multiple 30 kW high efficiency,

specific impulse 7 000 s (69 kN·s/kg)

• Hall thrusters: high power, higher thrust

• Deployed size: 58.4 m long × 15.7 m wide

• Stowed size: 19.7 m long × 4.57 m wide

• Mission design life: 20 years

• Launch date: 2017

• Launch Vehicle: Delta IV Heavy (3 launches)

• Development cost: 16 billion $ excluding launch : the show stopper!!

_{Source: Prometheus Project Final Report, NASA 982-R120461, 10/01/2005}

Nuclear Electric Propulsion for Space Exploration: example of the JIMO project

38

Selected: minimized

development

challenges & most

readily tested

Source: Documentation of Naval Reactors Papers and Presentations for the Space Technology and Applications International Forum (STAIF) 2006

The various nuclear power system concepts envisaged (200 kWe)

(21)

Nuclear Electric Propulsion for Space Exploration: example of the JIMO project

39 Documentation of Naval Reactors Papers and Presentations for the Space Technology and Applications International Forum (STAIF) 2006

Subsystem

mass (kg)

Reactor core

2 800

Radiation Shield

1 650

I&C

380 Power conversion

1 960

Heat rejection

3 300

Power System mass

10 090

(~

50 kg/kWe

)

Reactor outlet / inlet T 1150 / 911 K

HeXe pressure 2.0 / 1.0 MPa

Turbine outlet T / P 943 K/ 1.0 Mpa

Compressor outlet T / P 538 K / 2.0 Mpa

Radiator inlet /outlet T 505 / 379 K

Radiator area: 542 m2

Vessel/Reflector diameter 62 / 85 cm

Vessel length 160 cm

Core fuel height 61 cm

Safety rod diameter 13 cm

U235 fuel load 375 kg

av. fuel power density 26 w/cm3

Base Case Design Point

_{~1 MWth / 200 kWe}

Nuclear Space Propulsion: Electric (NEP) or Thermal (NTP)?

40

Nuclear Thermal Propulsion

Propellant

(H

₂

)

Nozzle

Nuclear Reactor

Heat addition

Nuclear Electric Propulsion

Spacecraft subsystems Experiments & Spacecraft Subsystems Heat Electric power Waste Heat (low T)

Nuclear Power Subsystem

(22)

Space Propulsion: some basics

41

For

‘’thermal’’ rocket engines

(chemical, nuclear thermal)

6 78 9 7

:

;<

_;,

=

%>% 7

;:

_;,

?

;<

=

%>% 7

?

_{: ;:}

1 @

ABCDE

@

BCBF

GH

IJ

G

_K

LMLNO

∝

_{R S)}

QR

TU

_V

W

Launch mass (cost) exponentially decreases

with

K

_LMLNO

X

YZ

Y)

K

LMLNO

_[

Specific Impulse:

\

]^]7

\

_]^ `

a

∆J

/

K

LMLNO

Chemical

(LH

₂

/LO

₂

_{) : M~13.8 g/mol, T ~3420 K }

X

_YZ

~480 s

Thrust ~ 2 000 kN; burn time ~500 s; thrust/weight ~150

and ‘’energy limited’’ performances (energy stored in chemical bounds)

Nuclear Thermal

(LH

₂

_{propellant): M~2 g/mol, T ~2700 K }

X

_YZ

~900 s

Thrust ~50 - 1 000 kN; burn time ~1 000 s; thrust/weight ~10 - 30

and performances limited by fuel resistance to high temperature H

₂

∆J

=

%>% 7

ln

_\

\

]^]7

_]^ `

K

LMLNO

U

: chamber temperature (K)

V

: molecular weight

f '

g

'

G

2 \

]^]7

\

_]^ `

a

∆J

/

K

LMLNO

(Tsiolkowsky rocket equation)

Space Propulsion: some basics

42

Round trip low Earth orbit  low Mars orbit:

minimum ΔV ~6 km/s

Chemical (Thermal) propulsion:

X

YZ

~480 s  ΔV/

X

YZ

~13 

V

hihO

/

V

jhikl

~3.7

Nuclear Thermal Propulsion:

X

YZ

~900 s  ΔV/

X

YZ

~6.7 

V

hihO

/

V

jhikl

~1.9

Example of Earth-Mars round trip: for thermal propulsion,

a twice higher

X

_YZ

reduces mass in LEO (cost) by a factor ~2

or enables to shorten manned round trip time (space radiations!)

X

YZ

= 480 s

X

YZ

= 900 s

Mars

Earth

(23)

Much higher I

_sp

with Electric Space Propulsion Engines (but much lower thrust)

43

Electrothermal

I

_sp

: 500 - 1 000 s

gas heated via resistance

or arc and expanded

through nozzle

3 classes of electric space propulsion concepts

Electrostatic

I

_sp

: 2 000 - 20 000 s

ions electrostatically

accelerated

Electromagnetic

I

_sp

: 1 000 – 7 000 s

plasma accelerated interaction of

current and magnetic field

Resistojets

Arcjets

Hall thruster

Ion thruster

(those are ‘’thermal’’ engines)

Magnetoplasmadynamic

thruster

Variable Specific-Impulse

Magnetoplasma thruster

VASIMR®

High-Power Electric Space Propulsion Engines: Estimate Performances

44 Source: Air Force Research Laboratory High Power Electric Propulsion Technology Development, Daniel L. Brown, Brian E. Beal, James M. Haas, 2010 IEEE Aerospace Conference (2010)

HET: Hall Effect Thruster, NASA-457M: Hall Effect Thruster

ELF: Electrodeless Lorentz Force (ELF) thruster

VASIMR: Variable Specific-Impulse Magnetoplasma thruster

Estimated performances of high-power propulsion scaled to 200 kWe

Concentric channel HET

ELF device

(24)

Electric Space Propulsion basics: Chemical vs Electric Propulsion

45

6 78 9 7

m

:n =

%>% 7

o $

%

)

1 :n =

₂

%>% 7

1

6 78 9 7

_{2 [ p}

6 78 9 7

o

$

m

%

2 [ p

Compared to chemical or nuclear thermal:

Electric thrusters: much higher

X

_YZ

much lower thrust

much lower thrust/weight ratio

low thrust/power ratio

Ex: advanced annular ion thruster (AAIT):

X

YZ

= 5 000 s, = 80%:

q

OrstYO

= 1 N 

u

v

~ 30 kWe

Mars manned mission with NEP:

requires 2.5 MWe feeding 10 AAITs (5 000 s

p )

Required electric power

o :

o $

%

) 6

78 9 7

[ p

₂

CP limited in total available energy at liftoff

EP is limited by electrical power available at any moment

CP: propellant velocity independent on thrust

EP: propellant velocity increased with thrust

CP: total mission impulse delivered at mission start-up

EP: total mission impulse accumulated during mission

Electric Space Propulsion basics: the highest I

_SP

is not the best choice

o $

%

) 6

78 9 7

[ p

₂

\

7w7

\

gx

y \

w

\

gx

_{| }}

z

{_~•

6 78 9 7

p

\

w

?

;:

_;,

7] %

₆

_{789 7}

_,€:a

[ p

6 78 9 7

=

%>% 7

;:

_{;, [ p}

;:

_;,

M

a

ss

Specific Impulse (Isp)

Optimum Isp

Maximum payload

Propellant Mass

\

_w

Electric Power

Supply Mass

\

_gx

Total Mass

\

_7w7

Constant

6 _{78 9 7}

over mission

,€:a:

Choice of particular thruster type for matching optimum

p

so as to maximize payload

NB: for chemical propulsion, highest payload mass for highest

p

• gx

$

_‚[

%

_\

o

gx

(25)

Electric Space Propulsion basics: the longer the mission time, the more efficient

47

\

`w

\

]^]7

\

_]^ `

\

]^]7

\

gx

\

]^]7

\

`w

\

]^]7

aƒ„

Δ<

[ p

2 •

[ p

xg

,€:a 1 aƒ„

Δ<

[ p

\

`w

\

]^]7

\

_]^ `

\

]^]7

\

`w

\

]^]7

aƒ„

Δ<

[ p

Assume DeltaV = 13 km/s,

p = 3 000 s, •

_xg

= 20 We/kg*

and constant continuous thrust

= 80%

= 60%

(*

•

_xg

of SP-100)

Chemical

p = 480 s

NTP

p = 900 s

The longer the mission time, the more efficient

(NB: for manned missions, the shortest, the better!)

Solar or Nuclear Energy to Power Electric Propulsion Engines

48

ESA’s Rosetta spacecraft required 64m

2

_{of solar array}

(26)

Eric PROUST

Lecture Series on SPACE NUCLEAR POWER & PROPULSION SYSTEMS -3- Space Fission Power & Nuclear Electric Propulsion Systems(last updated in January 2020) 49

Solar Power: System Integration Constraints

Solar or Nuclear Energy to Power Electric Propulsion Engines

Nuclear Electric Propulsion for a Manned Mission to Mars?

50

The Manned Transfer Vehicle

(NB: you also need a cargo TV…)

: 0.41 AU from the sun!

Subsystem

Mass (MT)

Nuclear Power System (2 x 5 MWe)

100 Propulsion System

30 Propellant tank (dry)

65 Structure

40 LH2 propellant

280 Payload

43 Total MTV mass

558

Source: Use of High-Power Brayton Nuclear Electric Propulsion (NEP) for a 2033 Mars Round-Trip Mission, M. L. McGuire, NASA/TM-2006-214106

60 days Mars

stay time

(27)

Nuclear Thermal Propulsion beats them all for manned missions to Mars!

51 Source: B. G. Drake “Human Mars Mission Definition: Requirements & Issues”, Human 2 Mars Summit, 2013

Electric Propulsion:

SEP: Solar

NEP: Nuclear

Chemical

(433 d)

NTP

(316 d)

Van Allen Belts

20 mSv

Mars Surface

10 mSv

Galactic Radiation

310 mSv

220 mSv

Solar Flares

260 mSv

150 mSv

Nuclear Reactor

0 mSv

50 mSv

Total

600 mSv

450 mSv

Comparison of radiation exposures

(same liftoff and payload mass)

Source: Sager, 1993; quoted by T. J. Laurence, ‘’Nuclear-Thermal-Rocket Propulsion Systems’’, in Nuclear Space Power and Propulsion systems, ed. By C. Bruno (Am. Inst. Aeronautics & Astronautics, 2008)

Space Nuclear Fission Power & Nuclear Electric Propulsion

52

The early space age years: nuclear fission power systems in earth orbit for defense missions

-

reactor and energy conversion technologies:

ROMASHKA, SNAP-10A, BUK/BES-5, TOPAZ I, TOPAZ II, SNAP 50/SPUR

Safety principles and key design challenges

-

Safety principles built on the feedback of experience (including accidents)

-

Mass, shadow shield, reactor, radiator, and reliability

-

Example of later implementation: the SP-100 design

Nuclear electric propulsion for space exploration

-

An illustration: the JIMO project

-

Space propulsion basics: specific impulse and propellant mass, thrust

-

Electric thrusters: high specific impulse propulsion engines

-

Space propulsion basics: thermal vs electric propulsion

-

Solar vs nuclear electric propulsion; Manned mission to Mars

Current developments

(28)

Towards Affordable Fission Surface Power Systems for Moon/Mars base applications

53

‘’The affordability goal

led to a decision to limit reactor fuel-clad temperature to 900 K

to minimize fuel and material development costs and maximize the use of existing technologies’’

40 kWe

* UZrH not selected ‘’because of unproven long life at high temperature

and the recapturing/development of a hydrogen retention barrier

*

Towards Small (and affordable) Fission Power Systems: one driver: the

238 _{Pu supply issue}

54 54 Source: P. McClure, D. Poston, Design and Testing of Small Nuclear Reactors for Defense and Space Applications, LANL/NASA, LA-UR-13-27054, 2013

E le c tr ic P o w e r (k W e ) Operational Time

Unlimited

238 _{Pu supply}

_Limited

238 _{Pu supply}

This chart includes estimates of

mass, practicality and utility of

each power source

The utility of solar power is

obviously dependent on distance

from sun and/or possibly of

day-night cycle

Yellow curve is estimate of utility

at 10 AU, dotted line is estimate

at 1 AU (no eclipse application)

Limited

238 _{Pu supply has}

lowered the threshold for

entry-level fission systems

(29)

Towards Small and Affordable Space Reactors: KILOPOWER (US, 2015-present)

55 Source: Design and Testing of Small Nuclear Reactors for Defense and Space Applications, P. McClure, D. Poston, LA-UR-13-27054, 2013

Surface power

Spacecraft power

Kilopower

Towards Small and Affordable Space Reactors: KILOPOWER (US, 2015-present)

A compact,

low cost

,

scalable (1 to 10 kWe)

fission

power system integrating

available component

technologies

:

• UMo fuel,

• **passive sodium heat pipes* for reactor heat**

transport,

• flight-ready Stirling convertors (ASRG program)**

• Titanium/water heat pipe radiator

to bridge the gap between RTGs and 40 kWe class

fission power technologies

for space science and exploration

56

* Tested with NASA Stirling Convertor development program

(30)

Towards small and affordable space reactors: KILOPOWER (US, 2015-present)

4.5 kWth core (1 kWe): KRUSTY nuclear Test

57

U8%Mo Cast metal core

(93%

235

_U)

Free Piston Stirling (FPS) Engines

FPS: a long time 100’s of million US$ NASA technology development program(s) (solar dynamic, SP-100, ASRG, ...)

Contact-free moving parts / planar springs

(31)

Towards small and affordable space reactors: KILOPOWER (US, 2015-present)

KILOPOWER scalability

59 Source: Nuclear Systems Kilopower Overview, Don Palac et al. NASA, 02/22/2016; KiloPower Project - KRUSTY Experiment Nuclear Design, D .I Poston, LA-UR-15-25540, 07/20/2015

1 kWe Thermoelectric

~4 m long

600 kg or 1.7 We/kg

Fuel core

Ф

_out

11.0 x H 24.0 mm

2

2.2 dm

3

_{; 2.3 W/cm}

3

32.9 kg

235

_U

800 We Stirling

~2.5 m long

400 kg or 2 We/kg

Fuel core

Ф

_out

12.0 x H 26.0 mm

2

1.9 dm

3

_{; 6.0 W/cm}

3

28.4 kg

235

_U

3 kWe Stirling

~5 m long

750 kg or 4 We/kg

Fuel core

Ф

_out

15.0 x H 28.0 mm

2

2.9 dm

3

_{; 15.1 W/cm}

3

43.7 kg

235

_U

10 kWe Stirling

~4 m tall

1800 kg or We/kg

Towards small and affordable space reactors: KILOPOWER (US, 2015-present)

60 Source: Mission design for the exploration of ice giants, Kuiper belt objects and their moons using KILOPOWER electric propulsion, S. L. McCarty et al., AAS 18-241

= ~2,500 kg dry mass spacecraft (w/o propellant tank, w/o science payload)

Preliminary analysis of exploration of Ice Giants, Kuiper Belt Objects and their moons

using

KILOPOWER electric propulsion

Assumption: 10 kWe KILOPOWER power system feeding two NEXT-C (gridded ion)

thrusters (+ 1 spare) operated at a constant 4,000 s I

sp

and 0.28 N thrust

Uranus / best performing 11 year mission:

~400 kg science payload,

~1,650 kg Xe, launch date 03/2044

total wet mass: ~4,500 kg

Neptune / best performing 15 year mission: ~350 kg science payload, 1,925 kg Xe, launch date 10/2037

NB: For radioisotope electric propulsion, 16y mission, 130 kg science payload, 3,200 kg wet mass: need for ~4 kWe RTG (or SRG) power!

An illustration of a potential application to electric propulsion for space exploration

(32)

Keldysh MWe Power Transport Module (Russia)

61

Transport Power Module (space tug):

3.5 MWth gas cooled fast reactor

4 x 250 kWe Brayton conversion units

Droplet or panel type radiator

High-Power 7,000 s Xe ion thrusters

2010 – Launch of project

2012 – TPM and NPSS draft design

Development

2013 – 2018: ground testing and

TPM flight test

preparation

Source: Concept of Electric Propulsion Realization for High Power Space Tug, L E. Zakharenkov et al., Progress in Propulsion Physics 8 (2016) 165-180

R&D at Keldysh: joint operation of Brayton power conversion loop and ion thruster

62

Nuclear power and propulsion system based on closed Brayton

cycle power conversion unit and electric propulsion

High-rpm test bench

Turboalternator compressor

Electric gas heater

35 kWe Electric propulsion test bench

Ion thruster cluster

Source: Study of Operation of Power and Propulsion System based on Closed Brayton Cycle Power Conversion Unit

(33)

And what about Europe?

63

(2012-2013)

roadmaps

Start Megahit roadmaps implementation

by preparing demonstrators for a

MWe class nuclear electric space propulsion

still waiting for a real start…

(2015-2017)

Nuclear Fission Reactors for Space Power & Space Electric Propulsion

64

Nuclear Electric Propulsion

versus

Solar Electric Propulsion

Nuclear reactors for

• Moon outposts (lunar nights last 14 days)

• Mars outposts (dust storms, 24 h nights)

Nuclear fission power vs Solar power

E

le

c

tr

ic

p

o

w

e

r

(k

W

e

)

Mission duration

*

(34)

Radioisotope and Fission Space Power Systems

10 W

100 W

1 kW

10 kW

100 kW

MW

Surface Power

In Space

Surface Power

Radioisotope

Systems

Fission

Systems

Mobile science, samples

Automatic mission

Human exploration

Samples, deep drilling, …life support

65

Atomic Energy and Alternative Energies Commission -

www.cea.fr

Thank you for your attention

Any question?

Last update: January 2020