NASA / TM--1998-208831 AIAA-98-3591

A Spherical Torus Nuclear Fusion Reactor

Space Propulsion Vehicle Concept

for Fast Interplanetary Travel

Craig H. Williams, Stanley K. Borowski,

Leonard A. Dudzinski, and Albert J. Juhasz

Lewis Research Center, Cleveland, Ohio

Prepared for the

34th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Cleveland, Ohio, July 13-15, 1998

National Aeronautics and

Space Administration

Lewis Research Center

December 1998

https://ntrs.nasa.gov/search.jsp?R=19990020975 2018-05-11T17:00:01+00:00Z
Acknowledgm rots

The authors wish to thank many who have been very helpful in providing guidance and expertise: to Gerald Hale

of Los Alamos National Lab (LANL) in the area of spin polariz_ d D3He fuel, to Stanley Kaye of Princeton Plasma

Physics Lab (PPPL) in the area of fusion confinement time scaling laws, to Joseph Warner of NASA Lewis Research

Center (NASA LeRC) in the area of YBCO superconductor material properties and prospects, to Ronald Moses of

LANL in the area of Dshape magnet designs and forces, to Lawrence Green of Westinghouse Science and Technology Center and Daniel Driemeyer of Boeing Co. in the area of ITER divertor design drivers, to Peter Turchi and

Hani Kamhawi of Ohio State University in the area of magnetic nozzle design and state conditions, to Larry

Grisham of PPPL and Masaaki Kuriyama of the Japan Atomic Energy Research Institute in the area of negative ion

neutral beam injectors, to Richard Kunath of NASA LeRC in the area of Ka band space communications, to Robert

Cataldo of NASA LeRC in the areas of start-up NiH batteries and Mars human payload systems; to John Miller of

JME, Inc. and Mary Ellen Roth of NASA LeRC in the area of high energy battery capacitors, to John Sankovic of

NASA LeRC and R. Joseph Cassady of Primex Technologies Corp. in the area of high power hydrogen arcjets, to

Mellissa McGuire of Analex Corp. in the area of reaction control, to Judith Watson of NASA Langley Research

Center in the area of Space Station truss network, and to Mohamed Bourham of North Carolina State University

and Stanley Milora of Oak Ridge National Lab in the area of fuel injectors.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Price Code: A03

Available from

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Price Code: A03
AIAA-98-3591

A SPHERICAL TORUS NUCLEAR FUSION REACTOR

SPACE PROPULSION VEHICLE CONCEPT FOR FAST INTERPLANETARY TRAVEL

Craig H. Williams', Stanley K. Borowski'",

Leonard A. Dudzinski"", Albert J. Juhasz ....

NASA Lewis Research Center

Cleveland, OH 44135

(216) 977-7063, -7091, -7107, 433 - 6134

ABSTRACT

A conceptual vehicle design enabling fast

outer solar system travel was produced predicated on a

small aspect ratio spherical toms nuclear fusion reactor.

Initial requirements were for a human mission to Saturn

with a > 5% payload mass fraction and a one way trip

time of less than one year. Analysis revealed that the

vehicle could deliver a 108 mt crew habitat payload to

Saturn rendezvous in 235 days, with an initial mass in

low Earth orbit of 2,941 mt. Engineering conceptual

design, analysis, and assessment was performed on all

major systems including payload, central truss, nuclear

reactor (including divertor and fuel injector), power

conversion (including turbine, compressor, alternator,

radiator, recuperator, and conditioning), magnetic

nozzle, neutral beam injector, tankage, start/re-start

reactor and battery, refrigeration, communications,

reaction control, and in-space operations. Detailed

assessment was done on reactor operations, including

plasma characteristics, power balance, power

utilization, and component design.

INTRODUCTION

The impetus for this effort was three fold.

First: in order to guide the long range NASA goal of

human expansion throughout the solar system, a

rational approach for long term advanced research and

development must be clearly articulated. Second:

currently funded nuclear fusion space propulsion

research must continually be shown to be on the critical

path for enabling order-of-magnitude improvements in

future space transportation capability. Third: a

conceptual vehicle design incorporating the proposed

design philosophies and related results of a recent series

of NASA Lewis Research Center (LeRC) papers was

the logical next step in the process 1'_.3.

The findings of these earlier papers

emphasized that for piloted, outer solar system missions

expected within the 21 't century, adequate payload

mass fraction (5% to 15%) and multi-month trip times

would require specific impulses (l_p) and specific

powers (cx) of 20,000 to 50,000 lbf sec/lb_, and 5 to 50

kW/kg respectively 1'3. Although contestable, it is the

judgment of the authors and many in the field that only

a single space propulsion technology exists at this time

that can reasonably be expected to offer this capability:

nuclear fusion, either magnetic or inertial confinement.

Nuclear fusion reactors can be broadly

classified into at least three groups: closed magnetic

(such as tokamaks, small aspect ratio toroids,

spheromaks, field reversed, etc.), open magnetic

(mirrors), and inertial concepts. Based in part on the

results of previous studies 1'4 of the attributes and

shortcomings of these reactor groups towards space

propulsion, a closed magnetic system was chosen for

this design concept. The high power density achievable

in closed systems, improved confinement, density and

temperature profile peaking, and spin polarization of

the fuel, provided a distinct advantage in their

application towards space propulsion. Further, while

the large aspect ratio tokamaks have been the

predominant focus of the fusion research community

for many years, their great size and mass render them

unappealing for space propulsion, where light weight is

paramount. At the other end of the closed reactor

spectrum, spheromak and field reversed concepts offer

tremendous hope for compact, light weight propulsionoriented systems. However, the dearth of experimental

data on their operation and indeed their engineering

feasibility render serious engineering assessment

difficult. As a result, a compromise was struck between

existing, extrapolatable experimental databases from

tokamaks and the largely conceptual compact toroids.

The small aspect ratio spherical toms, a concept

' M.E., P.E./Aerospace Engineering, Senior Member AIAA

'" Ph.D./Nuclear Engineering, Senior Member A1AA

°" B.S./Aeronautical and Astronautical Engineering, Member AIAA

.... M.S.E./Mechanical and Aeronautical Engineering, Senior Member AIAA
midwaybetweenthese two groups, was thus chosen to

serve as the basis for the vehicle concept.

Consistent with the "top-down" requirementsdriven approach documented elsewhere t, this vehicle

design concept was initiated by first establishing a

simple set of mission requirements, then producing a

consistent engineering design that satisfied those

requirements. This meant that current state of the art

systems, along with experimental results, were used as

the basis to extrapolate to what could be

technologically available to a human presence solar

system-wide of the not too distant future --- 30 years

from now. All system engineering analysis was

performed using existing computer programs, open

literature engineering sources, and basic engineering

calculations. The preponderance of the nuclear fusion

engineering data was obtained from Department of

Energy terrestrial power and scientific research

programs, while much of the propulsion system

engineering data was derived from NASA expendable

launch vehicle and conceptual nuclear thermal rocket

(NTR) design studies. Only limited NASA fusion space

propulsion system data exists beyond what was

accomplished since the termination of the 20 year long

fusion propulsion program at NASA LeRC in 1978.

MISSION REQUIREMENTS

The top-down, mission requirements-driven

design process began with specifying the desired

mission, trip time, and payload mass fraction. These

requirements were then used to define operation

parameters: mission distance, specific power, specific

impulse, and nozzle jet efficiency. These operation

parameters are directly related to four primary system

characteristics: structure mass, power out of reactor, jet

power, and thrust. The primary system characteristics

were the focus of the engineering design effort, where

iteration was conducted until the mission requirements

were satisfied.

The vehicle concept was to be able to perform

a rendezvous (one way) mission to the Saturn, piloted

by a crew of six, with a > 5% payload mass fraction,

and a trip time of less than one year. Performance

assessment of the same vehicle was also to be made for

a similar mission to Jupiter. Also investigated was the

propellant required to arrive at the specific destinations

at Saturn and Jupiter: their major moons Titan and

Callisto respectively. The Titan mission was selected

due to its demanding performance requirement,

scientifically interesting possibility of life on its surface

requiring human presence for investigation, its

dominant size among the Saturnian satellites, its

expected abundance of hydrogen for propulsion

application, and the abundance of fusion fuels D2 and

3He in the planet's atmosphere. The Callisto mission

was selected for similar reasons, but also to represent a

less performance demanding mission for comparison

purposes (as well as the Galilean satellite farthest

outside of the Jovian gravity well). The mission

distance was predicated on optimal planetary

orientation, though performance requirements for more

demanding planetary positions were also calculated.

l'he exploratory and scientific nature of the

mission drove the requirement for human presence.

Based primarily on existing humans to Mars mission

studies, a reasonable crew size of six was chosen. The

crew habitat payload design and its requirements, being

outside the scope of this study, were largely adopted

and scaled-up from current Mars mission spacecraft

designs. Although an ample (>5%) payload mass

fraction was initially set as a requirement, even a

generous_:y scaled-up Mars mission payload mass was

found to result in a lower payload mass fraction due to

the excessive total mass of the entire vehicle concept.

Therefore, an explicit payload mass fraction was

replaced with a specific payload mass value.

The piloted nature of the mission also drove

the requirement for relatively fast trip times. The one

year maximum was somewhat arbitrary, but was

represent ttive of long duration human experience in

low Earth orbit and consistent with some current Mars

mission ;tudies. The intent was to force the vehicle

concept to perform multi-month (vs. multi-year)

missions, despite interplanetary distances that dwarf the

more commonly thought of Earth to Mars transfers.

VEHICLE OVERVIEW

:igures 1 through 4 illustrate the overall

layout o, the vehicle concept. The cylindrical crew

habitat payload was forward of the propulsion system.

It was liaked to the central truss through an adapter

which also attached to the avionics suite and truss

booms supporting the communication antennas. The

forward central truss supported the four square panel

heat reje :ting radiators. Along the outside of the mid

central _rass were six slush hydrogen propellant tanks.

Within t_ e mid central truss was the D3He fuel tank and

refrigeration system for all propellant/fuel tankage.

Throughtmt the central truss were also various data,

power, coolant, and propellant lines. Outside of the aft

central truss were two Brayton power conversion

systems, including recuperators. Within the aft central

truss were the power management and distribution

system, the neutral beam injector, the refrigeration

system for all the superconducting magnetic coils, and

the starvre-start reactor and battery bank. Aft of the
Figure 1: Vehicle Concept Overview

(Aft View)

central truss was the spherical torus nuclear fusion

reactor and the magnetic nozzle. The overall vehicle

length was 260 m. The greatest deployed system

dimension was 150 m of the square heat rejection

(radiator) system and its connecting central truss.

However the maximum stowed diameter for any

individual system was limited to 10 m so as to fit within

a reasonable payload fairing, facilitating launch and onorbit assembly. The fully tanked initial mass in low

Earth orbit (IMLEO) was 2,941 mr.

The design philosophy followed was to locate

as much of the vehicle mass (power conversion,

reactor, and propellant) as close to and forward of the

thrust vector as possible to facilitate steering control

authority. Although shielded and emitting limited

neutron radiation, the reactor was kept as far aft of the

crew payload as possible. The modular vehicle layout,

system packaging, launch operations, and in-space

assembly sequence was expected to be conducted in a

way that maximized docking maneuvers rather than

labor intensive Extra Vehicular Activity (EVA). Since

the vehicle was designed for interplanetary cruise, only

two docking ports, located forward and side, were

provided to accommodate surface-based landing craft.

Table 1: Mass Property Summary

Payload 108

D3He fuel 45

Hydrogen propellant 1,292

Main impulse 1,220

Reaction control 21

Flight performance reserve 12

Residuals/losses 39

Structure 1,496

Central truss 15

Reactor 436

Power conversion 381

Refrigeration 54

Magnetic nozzle, divertor 8

Neutral beam injector 65

Start/restart reactor, battery 11

Reaction control 48

Avionics, communication 2

Cryo-tankage 13 I

Weight growth contingency 345

IMLEO(mt) 2,941
Figure 2: Vehicle Concept Overview

(Axial View)

I $I

Figure 3: Vehicle Concept Overview

(Top View)
Table 1 illustrates the mass property summary

for the fully loaded stack. The "payload mass" was 108

mt and consisted of useful payload only (crew, habitat,

consumables, scientific instruments, etc.) The "fuel"

was 45 mt and was solely D3He used to fuel the reactor.

The "propellant mass" was 1,292 mt and was slush

hydrogen for momentum transfer, reaction control,

reserves, and losses. It does not include system or

tankage mass. (For the purposes of calculating the

velocity ratio to be defined in the following section, the

propellant mass (Mp) was restricted to only that for the

main impulse. The balance of slush hydrogen (reserves,

etc.) was book-kept with the structure mass.) The total

structure mass (Ms) was 1,496 mt and referred to all

mass required to operate the propulsion system.

MISSION ANALYSIS

Fusion propulsion systems are expected to

operate at high enough Isp and a to produce

accelerations greater than the local acceleration due to

solar gravity at Earth's orbit (0.6 milli-g; where 1 millig = 32.1739 10 -3 ft/sec :) i. The normally thought of

conics of minimum energy trajectories followed by

today's chemical systems degenerate into nearly

straight line, radial transfers at these high acceleration

levels with continuous thrust. A "field-free space"

approximation can be invoked to greatly simplify the

usually complex orbital mechanics. Gravity losses and

optimum steering concerns can be neglected without

introducing too much error, obviating the need for

computationally intensive, numerically integrated

solutions to support preliminary analysis. In addition, a

"launch at anytime" approach to mission design is a

luxury that can usually be assumed for fusion systems

so long as the thrust to weight is great enough

compared to the local acceleration due to solar gravity.

As will be shown, despite an initial thrust to weight of

only 0.887 milli-g, the vehicle concept's trajectory was

reasonably close to that of a radial transfer.

At the initiation of and throughout the design

effort, a simple high I,v/high thrust algorithm 3 for radial

transfers was used to initiate and guide the design

process and monitor convergence towards study

requirements. This analytic, closed-form solution

previously published by the co-author was used for the

initial trip time and performance analysis. The

governing relations are based on the classic rocket

equation and auxiliary relations for high (constant)

thrust, constant mass flow rate (variable acceleration)

travel through field-free space.

At the conclusion of the design effort, the high

Figure 4: Vehicle Concept Overview

(Forward View)
fidelity, variational calculus-based trajectory

optimization program VARITOP was used to verify

vehicle performance and overall mission design.

VARITOP is a two body, heliocentric transfer

computer program for modeling low thrust space

propulsion systems 5. It is well known throughout the

preliminary design community for its good accuracy in

solving the two point boundary value problem and

integrating state variables. Good agreement was found

between the two trajectory design computer programs.

Figure 5 illustrates a heliocentric view of the integrated

VARITOP trajectory from Earth to Saturn.

A few definitions of operational characteristics

and system parameters are as follows. Specific power

(ix) is calculated differently throughout the space

propulsion community. Here we will use the common,

though perhaps not universal, def'mition of the ratio of

power out from the reactor system (P_0 (and sent into

the thrust generating device) divided by M s (including

reaction control propellant, reserves, and residuals)

(equation (1)). at for the concept was 3.92 kW/kg.

Paul

a = (1)

/_s

The nozzle jet efficiency (rlj) is the effectiveness of

converting transport power out of the reactor into

directed jet power in the thrust exhaust as defined in

equation (2). The _j for this class of systems remains

largely conjecture, consequently a value of 0.8 (i.e.

80%) was assumed based on known low power electric

propulsion systems and analytically derived minimum

fusion system efficiencies 6'7.

!

!

!

I

I

I

!

!

!

! IntegratedI

Rendezvous

Trajectory

Earth Orbit

Figure 5: Integrated Earth - Saturn Trajectory

(Heliocentric View)

Tal:le 2: Performance Analysis Results

Destinati.m Saturn Jupiter

Travel distance (AU) 9.1 4.9

a (kW/kg) 3.92 same

I_p(lbf sec/lbm) 40,485 32,590

Payload mass (mt) 108 same

IMLEO _mt) 2,941 2,925

Rendezvous trip time (days) 235 150

Jet power (MW) 4,916 same

r h 0.8 same

Thrust (lbr) 5,567 6,916

Mpdot (kg/sec) 0.062 0.096

cN_ 1.113 1.120

Thrust/mass (initial) (milli-g) 0.887 0.928

e.jet

_s - (2)

Pa_,

Using _ese definitions and constant total mass flow

rate (rh tm_, the total flow rate of propellant and reactor

fuel and g_ = 32.1739 lbmft/(lbf sec2)), the jet thrust

power (Pj_, thrust (F), Isp, and exhaust velocity (c) can

all be solved for using the familiar equation (3):

• 2

Fgj,p F2g_ m,o,_j c

- --- (31

Pj,, = "IsPo_,= 2 2rh ,o,., 2g .

Table 2 contains the overall performance

analysis results for the design concept. All vehicle mass

properties, including propellant loading, P_t, rb (and

thus Pj_), ct, and payload mass were fixed for both

missions• The I,p, F, and /h tot,t were solved to satisfy

the mission. For the Saturn rendezvous mission, the

thrust v,as 5,567 Ibf, tsp of 40,485 lbf sec/lb=, and a

rhto,, I {,f 0.062 kg/sec. For the Jupiter rendezvous

mission the thrust was 6,916 Ibf, l_p of 32,590 Ibf

sec/lb=, and a rhtml of 0.096 kg/sec. Rendezvous

missions were integrated for the optimal departure

dates _om Earth to Saturn/Jupiter using the same

payload module of 108 mt. The 235 day (8 month) trip

time to Saturn and 150 day (5 month) trip time to

Jupiter are rapid compared to trip times of

represe:ltative alternate concepts. A similar rendezvous

missior trip time to Saturn using chemical or even

nuclear thermal propulsion would be measured in

years. An analytic approximation was made of the trip

time for opposition orientation of the planets. The

Earth-Saturn trip time was lengthened by only 15%,

illustrating the relative insensitivity to launch date due

to the high thrust to weight capability. These results

6
demonstrate that the vehicle concept could accomplish

fast interplanetary trip times with significant payloads

over broad launch opportunities. Further improvements

in payload capability and trip time could be achieved if

fusion technology advances to permit the use of

optimal propellant and payload mass fractions.

The vehicle concept was designed to carry a

sub-optimal payload mass with a non-optimal

propellant loading, contrary to earlier study

recommendations _. Nuclear fusion technology offers

perhaps the greatest useful power out for high

thrust/high lsp propulsion technology appropriate for

interplanetary mission requirements. But its a for

credible propulsion designs, coupled with M, in excess

of 1,000 mt, is still too low for optimum payload and

propellant mass allocations below several 1,000's mt.

Simply put, projected fusion propulsion systems are

still too massive despite the large jet power they

produce. The ratio of exhaust to characteristic velocities

(V_) (c/V c, a convenient measure of optimal propellant

loading with respect to Ms) vs. payload mass fractions

for representative mission difficulty factors (AV/V_)

(where velocity increment (AV)) are illustrated in

figure 6, with the design concept also plotted. It shows

that the design concept was not optimal (AV/V c =

0.626) and considerably removed from the maximum

potential payload fraction (or similarly, propellant

loading). An optimal propellant loading for the same

propulsion system mass and AV/Vc would have been

4,100 mt, representing 20 propellant tanks (and

equivalently propellant tank launches) as opposed to

the 6 baselined. Less than 6 propellant tanks would

have driven the c/Vc to even greater values, lengthening

0.8

0.6

-_ 0.4

0.2

0

0.1

DV/Vc = 0. I

_h¢ / : Concep_ cle

c/V I 10

Figure 6: Payload Mass Fractions vs. c/Vo

trip times and driving the concept further from the

optimal value of c/Vc = 0.65. Thus a 6 propellant tank

configuration was chosen, yielding a non-optimal,

though otherwise reasonable c/V_ = 1.! 1.

SPACE OPERATIONS

Space operations issues pertain to IMLEOdriven Earth to orbit (ETO) launch requirements,

assembly and departure/arrival park orbit basing, and

rendezvous vs. round trip propellant loading modus

operandi. These issues are interrelated and have a

profound influence on total system viability.

The large M_-driven IMLEO represented a

fundamental obstacle to viable space operations. The

vehicle concept's IMLEO was -30 times what could be

delivered to LEO by a launch system in the Space

Shuttle-class (that is: 80 mt (orbiter) + 22 mt (payload)

= 102 mt to a 140 nmi circular orbit inclined 28.5

degrees) 8. This could be accomplished with the socalled "Shuttle-C" booster, a long proposed derivative

of the existing system that would operate without the

Shuttle orbiter. More likely, a new heavy lift launch

vehicle (HLLV) would be required. Sized for the

greatest single payload masses and volumes, the HLLV

payload capability could be as great as 250 mt to LEO.

This throw-weight capability of almost 2 "equivalent

three stage Saturn V's", was at the upper range of past

design studies 9 and would be required to deliver a

single, fully loaded propellant tank to LEO. Even this

class of launch vehicle would require a dozen HLLV's

to launch the initial configuration (and a half dozen for

subsequent propellant re-loading, on a mass basis only).

Even with a new HLLV, a serious viability

issue would still exist regarding launch availability. The

ability to launch up to a dozen HLLV's within a

reasonable time period will remain a dubious

proposition for the foreseeable future. And launch costs

associated with that many HLLV's would represent a

significant percentage of total mission costs. Thus,

dramatic reductions in launch processing and increases

in robustness in launch availability would be mandatory

for such a launch campaign to be viable.

On-orbit assembly would be a necessity for

this concept. Individual systems would be configured to

maximize simpler rendezvous and docking techniques

as opposed to telerobotics or labor intensive EVA. A

significant amount of on-orbit operations will still

nonetheless require human presence despite the

significant complexity, cost, and human factors issues

surrounding EVA. Major systems (payload, central

truss, heat rejection, etc.) would be assembled/deployed

in orbit sequentially, facilitating these operations rather

than attempting to minimize gross vehicle size by

7
maximizing final vehicle configuration packing

density. Assembly orbit altitude will most likely be no

higher than 260 nmi due to human radiation exposure

limits and HLLV performance limits, while minimum

orbits much below 140 nmi are unlikely due to long

term atmospheric drag and monatomic oxygen effects.

Expensive launch and operations, coupled with the

already high cost fusion system technology, will

mandate the design requirement for long life and

extended re-use. If fusion concepts actually prove to be

as massive as the current design suggests, limits on onorbit assembly, launch availability, and maximum

practical HLLV performance and volume capability

could prove development-lethal for fusion propulsion.

The vehicle concept was designed for

interplanetary cruise, thus a high altitude, sub-parabolic

orbit space basing would obviate the need for multiweek spiral escapes and captures at its origin and

destination. This also lessened the operation limits that

would otherwise be imposed on a vehicle that is a

source of high energy radiation and neutrons,

particularly near populated areas such as a space

station. These departure/arrival orbits could be low

lunar, lunar-altitude, or Lagrange orbits at the Earth,

and at very high minor moon or sub-parabolic orbit at

the major planet destinations. This would require

autonomous propulsion systems, possibly solar electric,

to transport and change-out propellant tanks at

departure and arrival points following their launch on

HLLV's [Earth) and yet to be determined launch

vehicles at the outer planets. Crew transport between

Earth/Titan/Callisto and the vehicle concept could be

performed by small, high thrust propulsion vehicles

designed for fast orbit transfer.

Rendezvous missions were selected as the

modus operandi due to their enabling of dramatic

reductions in propellant requirements, vehicle size, and

improved performance 3compared to carrying sufficient

propellant for round trip missions. The implied

requirement of a planetary refueling capability is of

great concern, but is consistent with a solar systemclass transportation system regularly journeying to and

between large outer planets with atmospheres and

moons rich in H2, D2, and 3He. Sources of available

propellant near high departure/arrival orbits, such as

water ice at the lunar poles, minor moons, outermost

major moons, and even asteroids, would greatly

facilitate refueling without entering into deep gravity

wells, provided the facilities could be established and

maintained at a sufficiently low cost. It is reasonable to

assume that in the time frame of fusion propulsion

systems, other technologies and infrastructure (such as

semi-robotic mining encampments) would be available.

D3He fuel would be acquired by either collecting solar

wind deposited or scavenging in situ major planetary

atmospheric deposits (if cost effective), which would

alleviate .-:he 3He supply issue. Processing 40 km 2 of

lunar regolith to a depth of 2 m, for example, could

yield 1 mid 6,100 mt of 3He and H2 respectively _°. If

these high orbit basing facilities were not available (or

prove too expensive) and all propellant had to be

supplied from terrestrial sources for entire round trip

missions, then operations could become overwhelming.

Although the mission analysis assumed

starting and stopping outside the effective gravity wells

of the origin and destination planets, the sub-hyperbolic

AV propellant requirements were calculated for the

Saturn mission. For the vehicle concept to go out of and

in to deep gravity wells, the additional propellant mass

would be as much as 51 mt for Earth escape and 22 mt

for Saturn orbit capture at Titan. The corresponding

spiral out and spiral in trip times were estimated to be

10 days and 5 days respectively. Thus escape and

capture maneuvers would have a minor, though not

insignificant, effect on propellant loading and trip time.

PAYLOAD SYSTEM

There have been several thorough engineering

studies of human spacecraft for Mars missions during

the last few years. The Transit Habitat from the recent

NASA Human Exploration of Mars study _ represents a

compilation of much of these engineering design

analyses. Consequently, it was decided to largely adopt

the Transit Habitat for the design of the crew habitat

payload system. The Transit Habitat's gross dimensions

and mass were then scaled-up to accommodate the

longer trip times for outer solar system missions.

Fhe original Transit Habitat for Mars was a

7.5 m diameter, 7.5 m total height cylinder, two level

structure able to accommodate a crew of six for a lg0

day trip to Mars followed by provisions for up to a 600

day surface excursion _m.The total habitable volume was

- 265 m _. This design was scaled-up for an up to one

year misfion by matching the diameter to that of the

other m_jor systems (10 m) and increasing the total

internal height proportionately to 6.75 m (9.2 m

external), where the total habitable volume then

equaled twice the original Habitat's (530 m3). This

enabled _ice the provisions for a crew of six for one

year, or the same amount of provisions for a crew of

twelve for a trip of six months. In addition, the volume

associated with the provisions for the original 600 day

Mars suface excursion could either be maintained as

backup c,r traded for other cargo. A potential significant

mass ccntributor not included in the original Mars

Habitat :,/as shielding for galactic cosmic rays and solar

flares. Its; mass impact projection remains premature.

8
Table 3: Payload System Mass Properties

Life support 12

Crew accommodations 45

Health care 5

Primary structure 20

EVA 8

Electrical power distribution 1

Communication and data management 3

Thermal control 4

Spares/growth/margin 7

Science 2

Crew 1

Total (mt) 108

The two story payload system included

primary structure, at least two docking mechanisms

mounted at the top and side, power distribution

systems, life support system, consumables (oxygen,

water, food, etc.), stowage, waste management,

communication, science instruments/experiments, and

fitness/recreation materials. It was a zero gravity

concept with no provision for artificial gravity due to

the less than one year trip time. The original Transit

Habitat was designed to be functional in 3/8 gravity

after it separated and landed on the Martian surface.

Table 4: Payload System Power Usage

Life support 12.

Thermal contract system 2.2

Galley 1.

Logistic module 1.8

Airlock 0.6

Communications 0.5

Personal quarters 0.4

Command center 0.5

Health maintenance facility 1.7

Data management system 1.9

Audio/Video 0.4

Lab 0.7

Hygiene 0.7

SC/Utility power 5.

Total (kWe) 29.4

The scaled-up total mass of the crew payload

was 108 mt (twice that of the original design). Table 3

illustrates the payload mass properties. The primary

contributor was crew accommodations (45 mt) of

which - 80% were consumables. Other major items

were primary structure (20 mt) and the life support

system (12 mt) H. Table 4 provides the nominal power

usages. The leading contributor was the life support

system (12 kWe). The total power requirement for the

payload was 30 kWe It

Defming and engineering the interface

between the payload and the vehicle represents one of

the most important, time consuming, and complex tasks

in preparing for launching today's space missions. In

advanced concepts such as this, the payload interface is

somewhat ambiguous. The structural attachment

(adapter) was a simple truss assembly of negligible

mass. Interface hardware (auxiliary power connections,

sensors, etc.) was also of negligible mass.

CENTRAL TRUSS

The primary structure linking the major

systems of the vehicle was chosen to be a truss network

in order to minimize mass while retaining strength. The

light weight truss material was Aluminum

Graphite/Epoxy (A1 GrEp). This material and the

strut/joint/node design was adopted from a tested,

earlier concept for the International Space Station. The

truss cross section was changed from a square to a

hexagon in order to accommodate the six propellant

tanks. A structural strut from a prior design was used,

where its length had been reduced from 5 m to 3.5 m,

yet the overall cross-bracing arrangement (one diagonal

per section) was retained. Figure 7 illustrates one truss

section. Table 5 contains the mass properties of a single

section, where the struts were scaled based on uniform

lengths and the connecting hardware mass properties

were maintained t2. Despite a central truss length of

-225 m (requiring 65 sections), the total mass was only

5 mt. Accommodating axial and lateral loading,

however, necessitated modest redesign.

Figure 7: Central Truss Section

9
The axial loading was assumed to be greatest

at the aft end, where the propulsive thrust would have

to be applied to the vehicle. A separate thrust structure

was designed to take the 5,567 lb r thrust load and

distribute it axially into the six longitudinal struts. Its

mass was minimal, with the 18 truss struts of radii 1.36

times greater than the baseline.

Table 5: Central Truss Section

Mass Properties

Longerons (6 + 6) 21

Diagonals (6) 15

Diagonal on hexagonal face (3) 5

Strut end joint (2* 19) 20

Node joint half (2* 19) 13

Nodes (6) 2

Total (kg) 76

Given the low bending moment limit of the

struts (scaled from the original 5 m strut cantilever

failure at 1,024 ft lbf) u, the lateral loads were of more

concern. The radiators with a mass of 59 mt each, at a

average total moment arm of 75 m, had to be attached

at each node and at each strut midpoint to reduce the

bending moment to a level comparable to the scaled

limit value. This case was of greatest concern at the

maximum acceleration, 1.514 milli-g's, encountered at

end of mission. The other systems produced lower

moments due to lower masses and/or small moment

arms. Ample attachment hardware was needed and its

mass was conservatively estimated as twice that of the

entire central truss network. Thus, the total mass of the

central truss network was 15 mt.

FUSION REACTOR SYSTEM

Plasma Modeling and Characteristics

Modeling of the plasma conditions was

performed through a I-D plasma power balance

computer program _3. It was designed to analyze

generic, small and large aspect ratio, tokamak fusion

reactors, inductive and non-inductive heating, driven

and ignited operation, burning DT, DD, or D3He fuels.

By pursuing peaked temperature and number density

profiles within the core of a plasma, a relatively small

fusion producing region was established, satisfying

Lawson and ignition criteria without necessitating large

beta (13) throughout the plasma. The lower temperature

and number density outer regions would contribute to a

volume-averaged 13value within MHD stability limits.

This approach is tremendously attractive for space

propulsioa applications where compact size, thus

reduced mass, is of paramount importance. Profile

shape factors (8) for temperature (T), number density

(n), and current density (J) were of the functional form

given in equation (4), and were integrated along the

minor radius (r) (where 0 ___r _< a) with a concentric

ellipse approximation used.

x(r) = < x > f (r) where

x= (T,n,J)

f(r) = (1 + 6r,,,,j)(1 - r 2/a2) a_"_

x = --_ x(r) rdr (4)

Charged particle and neutron power density,

including DD side reactions, were integrated as

functions of number density and radius (along with

temperature dependent reactivities). Bremsstrahlung

and synchrotron radiation power densities were also

integrated as functions of temperature and number

density (and thus radius). From these quantities, other

primary reactor characteristics such as plasma current,

magnetic field, confinement time scaling, etc., were

solved for while satisfying constraints such as critical

beta and plasma power balance. Volume averaged

quantities such transport power and radiation loss were

used to determine initial available charged power for

propulsion and available waste power for conversion to

auxiliary electrical power respectively. Plasma

transport loss power, which included convection and

conducti_m loss, represented the primary source of

fusion leaction charged products for propulsion

application and thus the quantity to be maximized.

D3He (1:1 ratio) was chosen as the reactor fuel

in order to maximize the charged transport power

output and minimize neutron output power fraction. It

was decided that in the time frame of this concept,

reactor cperation at a plasma temperature of 50 keV

would r._present only an incremental technological

challengt_ over that of a DT-based concept operating at

10 keV (and a fuel significantly more conducive to

space prapulsion application). Also, as was discussed

earlier, solar system-class operation presupposed

propellant and fuel supply availability in the hydrogen

and helium-rich outer planet atmospheres and satellites,

mitigating supply issues surrounding 3He.

10
Table 6: Selected Reactor and Plasma

Characteristics

Major radius (m) 2.48

Minor radius (m) 1.24

Aspect ratio 2.0

Elongation 3.0

Plasma volume (m 3) 225.8

Safety factor (edge) 2.50

Safety factor (axis) 2.08

Fuel ion density (102°/m 3) 5.0

Electron density (102°/m 3) 7.5

Plasma temperature (keV) 50

Volume averaged beta 0.318

Confinement time (sec) 0.552

Average neutron wall load (MW/m 2) 1.03

Average radiation wall load (MW/m 2) 5.20

Ignition margin 1.235

Toroidal magnetic field (centerline) (T) 8.9

Maximum magnetic field (coil surface) (T) 52.1

Gain factor (Q) 73.1

Plasma current (MA) 66.22

Bootstrap current fraction 0.934

Wall reflectivity 0.98

Number density profile shape factor 1.0

Temperature profile shape factor 2.0

D3He fuel with a spin vector polarized parallel

to the magnetic field was used to capitalize on the up-to

50% enhancement in fusion reactivity cross section 14'15,

tremendously improving the charged output power. The

methods of creating and utilizing polarized fuel have

been developed through theory (DT and D3He) and

experiment (DT) _4'15.Although much work remains on

maintaining fuel polarization, it is well known that

ionization and atomic collision processes cannot result

in depolarization 15. However, there are potentially

significant design impacts to the first wall, neutral beam

injection, and fuel injection that will be discussed later.

Table 6 illustrates many other reactor and

plasma characteristics. The total energy confinement

time (x) was 0.552 sec and was in good agreement with

the 1992 International Thermonuclear Experimental

Reactor (ITER) H-mode scaling law (0.565 see) that is

a function of plasma current (I,), toroidai magnetic

field (Bx), number density (n), total plasma heating

power (PL), average atomic mass (A), major radius (R),

minor radius (a), and elongation (_:) (equation (5)) _6

r = u.uz lit,'° 55/_T'n091 n°17 /-'I.'*-°55"°5At/_:n/ a)-0A9 R2.3_c0.7

(5)

A somewhat shorter confinement time was obtained

from the 1997 ITER L-mode scaling law 17of x = 0.216

sec. The primary driving terms in the x scaling were

found to be the large values of plasma current, applied

magnetic field, total plasma heating power, and major

radius. Although it is not clear which sub-ignited,

experimental database-derived scaling law would be

more representative of ignited plasma conditions for a

propulsion system, it was reasonable to conclude that

the proposed concept was consistent with what is

currently known.

The critical beta (13crit)constraint was satisfied

by requiring the total plasma current to be (equation

(6)):

lp = tic,i, a B r/fin (6)

where the Troyon coefficient (fiN) of 0.05 (a somewhat

greater than typical value (0.035 < 13N< 0.04), based on

a recent analytic study _s) was used for a low aspect

ratio tokamak. The large 13c,t and magnetic field

required for space propulsive applications suggested

that great leverage existed with maximizing 13Nin order

to minimize large Ip and its impact on reactor design.

A density weighted, volume averaged plasma

T of 50 keV and an n_ of 5*1020 m 3 were chosen.

Representative n and T profile shape factors of 1.0 and

2.0 respectively were chosen to enhance fusion power

production.

I_,,,,,, = 2.86"10 -4 f(v) a

- 0.5

(7)

The significantly greater fusion product nTx

for a D3He fueled reactor led to the requirement for

large plasma current (_ 66 MA) in order to obtain

sufficient confinement. To provide for this current, a

diffusion-driven bootstrap current, first experimentally

observed in large tokamaks, was relied upon. The large

T_ and n_ present in the core of D3He fueled reactors

could benefit from significant, arbitrarily (90%+)

large _9, bootstrap fractions, where only modest seed

currents provided by external heating would be needed.

A 93% bootstrap current (lboot) was used based on

equation (7), derived from neoclassical transport theory

of tokamaks 2°. f(v) was set equal to 0.3 to approximate

the T_ and n radial profile shape factors consistent with

11
the I-D model used. Equation (7) is valid so long as the

bootstrap fraction is above 75% and safety factor is

greater than 2 for MHD-stable operation. With a large

I_o, of- 62 MA, only a small seed current (I_,d) of-- 4

MA was necessary (equation (8)).

I_eea = Ip- Ihoo, (8)

An injection power (Pa,j) of 108 MW, based on the

estimated 2_ current drive efficiency (Y,b) of 0.75"1020

A/Wm _ , and an electron number density (no) of 7.5

* 102°/m 3 was calculated from equation (9).

n _R I ,._a

P_,,/ - (9)

_'nb

A sufficiently high 13(-30%) was chosen to

efficiently use the strong magnetic field and reduce

synchrotron losses. An edge safety factor of 2.5 was

chosen. A high wall reflectivity of 0.98 was chosen to

mitigate synchrotron power loss.

An ignited reactor mode of operation was

chosen, in addition to high bootstrap current, in order to

minimize the re-circulating power fraction required and

the concomitant conversion system mass for generating

injection power. It was thought that a continuously

thrusting propulsion system would be better served by

this mode of reactor operation, where charged transport

power was maximized and used exclusively for

propulsio_ purposes.

Power Generation and Utilization

Figure 8 illustrates the power input, output,

and utilization. Of the 7,895 MW of fusion power

produced, 96% was in the form of charged particles

with the remainder in (largely 2.45 MeV) neutrons.

Total power input, which included 108 MW of

injection power, was 8,003 MW, yielding a Q-value

(fusion power/injected power) of >73. More than 3A of

the power out of the reactor (6,145 MW) was charged

transport power (D and He ions, protons, and electrons)

used solely for direct propulsion. Radiation in the form

of Bremsstrahlung (1,016 MW) and synchrotron (535

MW) power, along with neutron power (307 MW) and

other waste heat was used to produce 400 MW of

electrical power.

The choice of fuel and design of some of the

reactor components was intended to maximize the

fraction of useful power-out while minimizing the

radiation

Brem

synch

108

neutral beam

fusion

charged

neutron

7,895

7,588

307

injected

total

1,551

1,016 i neutron[ 35

535

1,551 _ 1272

I _ 1power conversion [ 1,682

[distribution I

injector 1_ 367 [ 1 _ 32

t

108

8,003

6,145 _[ transport/propulsion J

to space

_1 reaction control]

Figure 8: Power (MW) Input, Outp at, and Utilization

12
fraction of power lost to (and the mass devoted to

managing) unrecoverable radiation. This permitted the

dedication of all charged power to propulsion, while

scavenging only waste Bremsstrahlung and synchrotron

radiation from the first wall to produce electrical power

through a heat cycle. Since the electrical power that

could otherwise be produced at high efficiency was in

excess of requirements, the reactor was deliberately

designed to permit as much of the neutron radiation as

possible to escape directly to space (35 MW). Electrical

power requirements were dominated by the negative

ion neutral beam injector (N-NBI) at 367 MW which

provided the 108 MW of injection power needed for

current drive. Most of the remaining power was needed

by the arcjet reaction control system. All other auxiliary

power requirements were two or more orders of

magnitude lower and included superconducting coil

refrigeration, propellant tankage refrigeration, fuel

injector, start/restart battery recharge, communications,

avionics, and crew operations. All waste power was

rejected via the radiator arrays (1,682 MW), with the

RCS power (32 MW) exhausted/radiated to space.

To sustain a total fusion power of 7,895 MW,

D3He fuel with a specific energy of 3.52* 10 _4J/kg must

be consumed at a rate of-22 mg/sec. The fraction of

fuel bum up fib) was estimated from equation (10),

where <_v> is the Maxwellian averaged fusion

reactivity for D3He fusion reaction (0.536 < (<av> 10.22

m3/sec) < 1.726 for 50 < Ti (keV) < 100).

fh = I1 + 2 1-1 (10)

nT <o'v>

The average D3He burn up fraction was estimated to be

only 1%, requiring a fuel mass flow rate (thf_e_) of 2.2

g/sec. At this consumption rate, a 235 day trip to Saturn

required 45 mt of D3He (27 mt of 3He). Note that - 18

kg/yr represents a rough estimate of the potential U.S.

annual production capability of3He 22

Reactor Components

A small major radius (2.48 m), small aspect

ratio (2.0) device was chosen to minimize size and

mass. The scaling for 13 favored elongated (3:1),

compact devices with large plasma currents (- 66 MA)

and moderately large centerline magnetic fields (8.9 T).

These magnetic field requirements led to very large

toroidal field (TF) coil currents (9.2 MA). Twelve coils

were used to generate the toroidal magnetic field; seven

poloidal field (PF) coils were used to provide the

necessary plasma stability.

Figure 9 illustrates the radial build, including

an upper half of one TF "D-shape" coil. Beginning with

the major axis centerline, a 7.35 cm radius annulus

provided a flow channel for the slush hydrogen

propellant. A 21.8 cm thick Titanium alloy provided

structural support for the current-induced coil structural

loads. Twelve 13 cm diameter high temperature

superconducting TF coils carried the currents to

generate the toroidal magnetic field. The cylindrical

LiH blanket, 45 cm thick, protected most of the central

core from the 2.45 MeV neutrons. A toroidai shell of

gaseous helium (GHe), 23 cm thick at 25 atm, served as

the heat transfer fluid for the first wall, which was

made of a high strength (at very high temperature)

Molybdenum alloy. The first wall's heat was due to

average neutron and radiation loadings of 1 and 5

MW/m 2 respectively. The helium flowed through

toroidal shells of the Molybdenum alloy that were 0.5

cm (outer) and 0.75 cm (inner) thick. A scrape off layer

of ash was assumed to be 10% of the minor radius, with

the minor radius being 1.24 m. The outer radial

dimensions were the same, except for a slightly thicker

blanket (50 cm) to better attenuate the neutron flux.

The cross-section of the TF coils is illustrated

in figure 10. The coils were 75% by area YBa_Cu307

(YBCO, a high temperature superconductor)

surrounded by a helium refrigerant, followed by an

aluminum lithium outer shell. YBCO was chosen for its

extremely high critical current density at moderate to

large external magnetic fields, even at temperatures as

high as 77 deg K 23. These high densities permitted

greatly reduced central coil radii compare to Type I1

superconductors, more efficient use of central core

volume for neutron attenuation, and permitted

incorporating a central propellant flow channel. The

minimum coil radius was calculated by setting the sum

of all magnetic field components from the other coils 24

(including that of the far end of the same coil) equal to

an empirically derived equation for critical current

density (J_0 as a function of external field (Boxt) 23.The

result was equation (11), where Ic is the current in a

single TF coil, N is the number of coils, R_ and R2 are

the distances from the major axis to the centerline of

the inner and outer TF coils respectively, rc is the radius

of the TF coil, fc is the cross sectional area fraction

carrying current in the TF coil, _ is the permeability of

free space, and A and B are the curve fitted exponent

and scale factor of the exponential empirical equation

of YBCO Jerit VS. Bex t.

13
1.736 m

/ /

///

/

/

--• 0.50m

0.0735m _ _-

Slush H2

flow channel

Cylindrical

Ti-6AI-4V

strengthener

YBCO/He/AI Li

superconducting

coil

Cylindrical

LiH blanket

2.48 m

0.2407 m

0.0075 m.

_4-- 1.24m

2.916 m

1/4 sector

Ti-6AI-4V

strengthener

\

I

1

TZM-shelled

first wall

Scrape-off

layer

r

Figure 9: Reactor Radial Build

The limiting B_t at the coil was 20.4 T with a

corresponding rc of 6.57 cm. The total field at the coil

(self plus external) was 48.4 T; though as will be

explained later, the coils were centrally rcpositioned,

driving the final field at the coil to 52.1 T. The J,,_ was

conservatively evaluated at 77 dog K, although the

conductor was actually cooled to 70 deg K.

The load carrying assessment was made by

utilizing a reasonably accurate (99%+) approximation

for sizing constant tension D-shape magnets for fusion

reactors 24. By calculating the force per unit length (f0

given by equation (12) (where radius of curvature

angle(_b) and conductor radius (q)) then substituting

into equation (13) defining the radius of curvature (p)

of a flexible conductor under constant tension (T), the

differential equation (14) could then be solved

describing the geometric shape of the coil 2s (where z is

the axi: of toroidal symmetry).

(

/ZoI_ N

2z R

1+

1 1

+

1_ 0 A LLzr .B

(11)

where R = R, + rc

14
0.0657 m

"_[-- 0.5 m _

0.24 m

\ Li. I\

_ Superconductor

Reactor

Center

Figure 1O" TF Coil Cross-Section

Ti-6AI-4V

Strenglhener

fl = _ 1 + cos_b + Rln 1.284R

P crN

T

p, = -- (13) f,

(12)

pl dz---5-=+ I

(14)

The geometry of the D-shaped TF coils were

determined by curve fitted polynomials to the solutions

of equations (12) through (14) of the form of equation

(15). Equation (15) was parameterized for four

essential coil dimensions (P_), where the outermost coil

location (p2/R2) for the d_ = 0 position was set as the

independent variable.

4

RZ--

(15)

The radial build resulting from equations (15) produced

a design with an excessive R_ (for the H2 propellant

channel and central strengtheners) and insufficient

volume for the cylindrical LiH blanket. As a result, R_

was decreased and the other primary dimensions were

adjusted accordingly to maintain the general D shape.

The tension loads, however, were recalculated for the

actual TF coil positions for consistency, even though

this produced greater loads and thus required more

massive strengtheners. In addition, both vertical

dimensions had to be increased 50% to accommodate

the large elongation (3.0) of the reactor design. The

original algorithm (equation (12)) was not found to be

conducive to highly elongated toroids. The end result of

this approach to sizing and placing TF coils may not

have been rigorously correct, but was reasonably self

consistent and accounted for the primary current-driven

tension forces on the TF coils.

The high current (9.2 MA) carried in the coils

that produced the tremendous tension loads necessitated

strengtheners made from the highest yield strength

material suitable for cryogenic applications. A

Titanium alloy, Ti-6AI-4V, was selected for its high

yield strength (19 *l0 s N/m 2, solution treated and aged

ELI) even at 32 deg K 26. The Ti-6AI-4V cross section

design chosen was that of a I/, sector cylinder (figure

10), mounted onto the outer (opposite the reactor core)

TF coil surface, to efficiently carry the tension load

with minimum mass. This design was used for the

strengtheners of the top circular and outer elliptical coil

sections, while a cylindrical design was used for the

strengthener of the inner linear coil section to more

efficiently share the greatest loads. Despite its

enormous yield strength, the thickness of the outer and

top 1⁄4 sector cylinder strengtheners, and inner cylinder

strengthener were calculated to be 24.07 era, 35.7 cm,

and 21.85 cm respectively. The strengtheners were the

second greatest massive structures in the entire vehicle.

The attractive Ampere's law-driven forces between the

TF coils were also calculated but were found to be less

than 6% of those of the tension loads. Cross-bracing for

these as well as the "overturning" forces were included

in the design, with the minor additional mass penalty

assumed to be accounted in the PF coils. The total 12

TF coil mass was calculated to be 221 mt.

The LiH neutron blanket also used a 1⁄4 sector

cylinder design so that the coil would be protected from

neutron radiation no matter where in the plasma the

neutrons were created. The orientation of the blanket

was directly facing the reactor core (i.e. on the opposite

side of the coil from the strengtheners, figure 10). This

design minimized blanket mass while also permitted

neutrons that were not impacting the coils to pass

directly to space. This less than 4x shielding design

meant that human operations could not be conducted

everywhere around the reactor without protection while

15
the reactor was operating. However, the convergence of

the blankets at the forward (crew module-facing) and

aft (magnetic nozzle-facing) poles of the reactor would

provide sufficient protection for human operations

within certain angular constraints. A minimum 50 cm

blanket thickness was recommended for D3He fueled

reactors 27'2s. Packaging constraints within the central

core driven by structural loading required that the LiH

shielding there he reduced to 45 cm. This increased

neutron activation/heating and remains an issue to be

resolved. The calculated LiH blanket mass was 56 mt.

Over 88% of the 307 MW of neutron power

was intercepted by the LiH blankets; the balance of the

power was permitted to escape into space. The neutron

heating consisted of 92 MW of power into the

cylindrical blanket and 180 MW of power into the 1⁄4

sector blankets. This heat was removed by

supplemental GHe from the heat transfer system used at

the first wall. Though not shown on figure 9, a 3 cm

diameter GHe channel adjacent to the coils within the

LiH was included in the GHe first wall thickness to

provide the heat transfer mechanism. This heat was

assumed to be of comparable quality to the other waste

heat in the system. An attractive alternative would be to

examine if the space facing blanket sides would be able

to re-radiate heat to space as gray bodies while

remaining within material thermal limits. Initial

geometric analysis indicated that the arrangement of the

blankets was such that most neutron radiation was

intercepted in the volumes immediately forward of and

aft of the reactor, thereby protecting the vehicle

systems from activation and structural damage.

Minimal analysis was done on the PF coils.

Based on ITER and similar test reactors, seven PF coils

were assumed, with coil and shielding designs and

masses scaled to 60% of the TF coils as recommended

by similar advanced small aspect ratio reactor studies 2t.

The cross sectional makeup was assumed to be

identical to the TF coils, including strengtheners, but

without LiH blankets, since much of the PF coils would

be shadowed by the TF coil blankets. The total 7 PF

coil mass was calculated to be 133 mt.

The thickness of the first wall radiation shield

and GHe pressure vessel was based on evaluation of the

pressure and energy deposition within the structure.

Neutron and Bremsstrahlung/synchrotron radiation wall

loadings were 1.03 and 5.2 MW/m 2 respectively. These

values are representative of other power reactor designs

for first walls and divertors. To accommodate these

Ioadings, the molybdenum alloy "TZM" (Mo-0.5Ti0.1Zr) was selected based on its suitability for very

high temperature applications such as radiation shields.

TZM is also expected to provided high synchrotron

radiation reflectivity, which is mandatory to facilitate

its re-absorption into the plasma. TZM's high yield

strength t f 62 MPa (at 1650 deg C) 29 with a mass

density hrif that of tungsten, made it a good choice to

withstand the circumferential stress due to the 25 arm

pressure of GHe heat transfer fluid within the pressure

vessel. The initial thickness based on stress limits (0.5

cm) had to be increased for the inner (first wall) to

permit sufficient absorption of the estimated 50 keV

gamma radiation. Using an extrapolated mass

absorptiort coefficient 3° (p,/p) for molybdenum of 2.22

cm:/g, the required shield thickness (x) was estimated

from equation (16):

po_ =p., a+As+x e _PJ

a+ As

(16)

A thickness of 0.75 cm was sufficient to reduce the

radiation power (P_,) to under 100 W (Po_,), where As

was the scrape off thickness and p was the Mo mass

density. The mass of the first wall/pressure vessel was

calculated to be 18 mt. In a related issue, TZM could

have a detrimental effect on spin polarized fuel. It has

been suggested from today's research reactors that

recycling of fuel off a metallic first wall surface could

cause significant depolarization _5.This remains an issue

to be resolved should this effect prove to be true.

A 7.35 cm radius channel through the center

of the reactor was provided for the thrust augmentation

mass of tile slush hydrogen propellant. Although some

propellam warming was expected to take place, the

primary reason for channeling the propellant in this

way was to align the bulk (96%) of the propulsive mass

flow velocity and downstream magnetic nozzle thrust

vectors. As will be discussed, this approach has the

potential to cause adverse impacts to reactor operations.

Reactor 1=ueling

"Re preponderance of fuel injector design,

fabrication, and operation experience has been with :D,

and DT faelling. Two stage, light gas gun devices have

accelerated 1 - 3 mm frozen 2D pellets up to 2.5 km/sec

3t. 3:. Alternate concepts such as electrothermal guns

have been tested with plastic (Lexan polycarbonate)

pellets to similar velocities 33. The gross dimensions and

masses fi ,r most concepts were small compared to other

major sy:tems (<1 m lengths and <1 mt masses) 34._5.

_/ery little effort, however, has be devoted to

issues sarrounding injection of frozen D pellets

encapsulating (liquid) 3He. It is thought that injecting

such pellets during ignited, high temperature operation

will be very difficult ,5. An earlier conceptual design

16
study of a mirror reactor propulsion concept suggested

a two stage mechanical (centrifugal) and

electromagnetic railgun injector be used for D3He

pellets 36. This study, as well as others, suggested a

potential approach involving frozen light metal shelled

(lithium deuteride) pellets encapsulating 3He fluid _5'36.

Estimated final velocities of up to 30 km/sec with

acceleration lengths limited to up to 20 m forced one

study to use light metal shelling to mitigate barrel

heating and stress on the pellet. The light metal shelling

did have a detrimental effect (17% loss in jet power) 36

on reactor performance, even at small (< 6%) pellet

mass fraction. Pellet evaporation, however, .is presently

not thought to be a major potential contributor to fuel

depolarization _5.

Given the paucity of experimental data, it was

hypothesized that a D3He fuel pellet injector would be

no greater than 1 mt in mass (including injector, feed

lines, knife, piston), with a modest (< 5 kW) power

requirement. The ample central truss length should

permit a sufficiently low pellet acceleration so as not to

require light metal shelling. Tankage mass for the D3He

fuel was scaled from the slush hydrogen propellant

storage system. The 45 mt of fuel required a 3 m radius

cylindrical tank, 17 m long and a "wet" mass (including

reserves, residuals, and losses) of 7 mt.

Divertor

By diverting the outer plasma layers to the

propulsion system, fusion energy served to directly heat

propulsive mass without first going through

intermediate, material heat flux-limited, low-efficiency

power conversion equipment. An efficient divertor thus

enabled a direct thrust approach to space propulsion by

transferring high energy charged particles out of the

reactor and into the reservoir of the magnetic nozzle.

As was shown, 99 % of the particles exiting the reactor

will be un-reacted D and 3He ions and electrons. The

balance of the particles (neutral impurities, 4He, and p

ash) must be kept to low densities to maintain good

conf'mement and maximum propulsion performance.

As these particles cooled and migrated to the outer field

lines in the scrape-off layer, they eventually were

pulled into the divertor and out of the reactor. Plasma T

and n in the scrape-off were calculated to be T < 4 keV,

and ni < 1.95* 102°/m 3 (figure 11).

The divertor's geometric design (single null,

located in the reactor aft) was patterned somewhat after

ITER's 37. Three PF-like divertor coils (carrying

reduced current through a smaller coil radius) were

within the reactor's minor radius, between the TF coils

to divert the magnetic field outside of the reactor. The

100

>

80

ca_

E

_ 6O

._ 4o

_ 20

Z

I I

I I

I I

I I

m

Scrape Off

Layer

Number Density (electron)

Number Density (ion)

0

0.0 0.2 0.4 0.6 0.8 1.0 1.2

Minor Radius (m)

Figure 11: Reactor Plasma Temperature and Number Density Radial Profiles

1.4

17
field lines exited and re-entered the reactor through the

double annulus, carrying the exhaust plasma to the

magnetic nozzle. The estimated combined width of the

double annulus and middle coil was 0.5 m, scaled from

ITER geometry 3_. The divertor coils' plasma facing

material must be able to withstand high temperatures

and high erosion rates. Carbon fiber composites, such

as Aerolor A05 39, are likely candidates to tolerate

surface temperatures up to 1,500 deg C with peak heat

loads of at least 5 MW/m 2 and transient heat loads

potentially as great as 20 MW/m 237. Total mass of all

three coils plus additional system mass (structure,

power lines, etc.) was set to 4 rot.

The divertor's operational concept was

fundamentally different than ITER's or any of those

used in today's experimental reactors. Today's reactors

are not designed to exhaust vast quantities of transport

power, operate in steady state, accommodate immense

wall loading without assistance of radiation from

injected inert gas, or exhaust plasma without collision

with a target. In addition, today's divertor structural

designs and resulting masses are driven by

requirements that are not expected to be a primary issue

with a space propulsion concept. ITER's divertor

consists of 60 cassettes weighing 25 mt each 37"39"_.

Their mass is attributed to both withstanding the

immense structural loadings initiated by disruptions and

also shielding the vacuum vessel from neutron

radiation 37. It was assumed that steady state operation

of the reactor concept would at least minimize if not

eliminate disruptions altogether. And since there will be

no vacuum vessel, the other reason was mute. As a

result, the entire divertor was envisioned as a series of

TZM structures covered with Aerolor A05 plasma

facing material protecting the divertor coils. The coils

were assumed to be the same cross-sectional design as

the TF/PF coils, although at a much lower, to be

determined current density. Since the divertor coils

were inboard of the TF coils, a difficult task awaits in

designing them to be of sufficient field strength to

extract plasma scrape off without having a detrimental

affect on the primary field and the fusion process.

Resolution of this problem is fundamental to the

feasibility of this concept.

MAGNETIC NOZZLE

The conversion of the reactor's transport

power into directed thrust was accomplished by the

magnetic nozzle in two steps. In the first step, the

nozzle mixed high enthalpy transport plasma from the

divertor with the injected hydrogen propellant in order

to reduce the excessive temperature and increase total

charged propellant mass flow. In the second step, it

convertet l the propellant enthalpy into directed thrust

by acceh_.rating the flow through diverging magnetic

field lines. In addition, its magnetic field prevented the

high tem_rature plasma from coming in contact with

the nozzle's coils and structural members that make up

the thrust chamber. Thus for a fully ionized flow, the

lines of magnetic flux served as the containment

device, minimizing heat transfer loses and the need for

actively cooled structure.

The l,p's of 20,000 to 50,000 lbr sec/lbm and

corresponding a's required for multi-month travel to

the outer planets required ion reservoir temperatures

(T_) of 100's eV. As was discussed, the too great

temperature and too small number density plasma

layers that entered the diverter had to be adjusted prior

to acceleration through the nozzle so as to produce the

mission appropriate Isp. This was accomplished by

heating/ionizing slush hydrogen thrust augmentation

propellant by the escaping reactor plasma, thereby

producing the desired values of bulk plasma

temperature (thus lsp) and mass flow rate (thus thrustto-weight). The propellant was preheated by its pass

through the reactor core with residual heat from

absorbed neutron radiation. The propellant was then

injected into the reservoir along the nozzle centerline.

The "reservoir" of the magnetic nozzle was

somewhat analogous to a conventional liquid chemical

rocket engine's combustion chamber. Adjacent to the

reactor's divertor, it consisted of two small radius

superconducting coils of the same design and materials

as the Tt" coils. Forming an "effective" I0 cm radius

solenoid, they provided the meridional magnetic field

to confine the converging propellant and reactor plasma

streams until their temperatures equilibrated. The

reservoir was in large part a "virtual" chamber due to

its mostly skeletal design where magnetic field lines

defined the flow boundary for charged particles passing

through. This design minimized mass and heating

concerns but also placed a premium on rapid, effective

ionizatior., and enthalpy equilibration (i.e. neutrals were

lost through the sides). The second small radius coil

also constituted the "throat" for the nozzle, where

choked f_ow (sonic) conditions existed. An arbitrarily

larger current loop radius third coil in the downstream,

diverging section provided additional curvature to the

magnetic field. The entire length of the assembly was

somewhat arbitrarily set to 12 m with a total mass

estimated at 4 mt. The nozzle coils were considerably

less mas::ive than the TF/PF coils due to negligible

neutron r ldiation, significantly lower required magnetic

field (thus lower coil current and structural loads), and

lightweight AI-GrEp composite structure. The entire

layout, including number of coils, geometric and

magnetic field curvature, was largely notional and was

18
Figure 12: Magnetic Nozzle and Power Conversion System

intended to foster discussion, analysis, and

experimentation into its salient aspects. Until mature

design concepts with experimental data are available,

magnetic nozzle designs, operations, and rb for fusion

propulsion will be largely speculative. Figure 12

illustrates the layout of the magnetic nozzle.

Much of the potential jet power could be lost

unless the internal energy can be efficiently converted

into accelerated axial flow. Estimates of the relative

importance of r h on piloted interplanetary travel have

been recently reported i. It has been shown that low rlj

is particularly detrimental to payload mass fraction,

since decreased rb (at constant thrust) significantly

increased propellant consumption. An example is

illustrated in figure 13 for a case where c/V, = 0.7,

representative of a fast, piloted outer planet mission.

Although at a greater payload ratio than the design

concept (17.7 % for a 100% efficient nozzle), the

payload ratio is seen to be a strong function of rb,

vanishing at an qi less than 50%. Thus efficient

magnetic nozzle designs appear to be essential for the

class of systems envisioned for fast interplanetary

travel, where r h < 70% may not be tolerable.

20%

15% 10%

[ J Payload loss due to

¢_ 5% "['/ increased propellant

T consumption

0% ; • ,' ' l ' t '

60% 70% 80% 90% 100%

Nozzle Efficiency

Figure 13" Nozzle Efficiency vs. Payload Fraction

19
Magneticnozzleexperimentation applicableto

fusion-classpropulsion has been recently initiated 4_ by

a team at NASA LeRC, Ohio State University (OSU),

and Los Alarnos National Lab (LANL). A series of

three small, proof of concept experiments are focused

on providing key experimental data for nozzles

intended for GW jet power level, quasi-steady plasma

flows. The first of these experiments is focussed on

how the plasma would detach from the magnetic lines

of force upon exiting a magnetic nozzle and produce

net thrust. This problem lacks both theoretical and

experimental definition and may represent the leading

obstacle towards developing a practical fusion

propulsion system. Of primary importance is obtaining

plasma number density (thus mass flow rate) and

velocity distribution data as functions of axial and

radial position downstream of the throat. By integrating

these quantities, P)_ and F can be calculated, inferring a

value for rlj. A variety of operating parameters should

be investigated, such as reservoir temperature (T,,,),

density (Pr,_), magnetic field strength (B_), choice of

propellants, and others. Initially using helium

(eventually hydrogen) ions to mimic the exhaust

plasma, these experiments will enable estimation of a

nozzle's performance that can be correlated to fusion

propulsion systems. Further experimentation will

include designing and testing a rudimentary sub-scale

magnetic nozzle similar to the one used for this

concept, as well as designing and testing propellant

injection for thrust augmentation. Until experimental

data is available to correlate theoretical models, the

performance of direct plasma-propelled fusion

propulsion systems will be to a considerable extent

conjecture. Some theoretical analysis projecting and

defining their operation is available however.

The theory of how a magnetic nozzle would

operate and the primary obstacles anticipated in its

development were the subjects of a lengthy analysis 42.

Idealized I-D, MHD-modified fluid dynamic equations

for a magnetic nozzle with an applied meridional

magnetic field were derived in that work. Various

simplifying assumptions were made such as neglecting

charged particle-neutral collisions, assuming a fluid (as

opposed to a kinetic) model, and neglecting various

loss mechanisms such as transport, radiative, resistive,

and Hall effect. A simple, single turn coil was assumed

to form the throat of the magnetic nozzle. Using this

approach, along with other supporting analysis 43,_,4_,

the flow conditions at key points through the nozzle

were estimated. Isentropic flow was assumed

throughout, with an arbitrary nozzle efficiency of 80%

assumed. For simplicity, it was assumed that the 20%

nozzle loss took place at propellant-fuel enthalpy

transfer.

"'he required downstream conditions for the

exiting fl)w (rljct = 3.136 kW/kg and 1,, = 40,485 lbf

sec/lbm) were used to initiate the flow calculations. By

working back up through the nozzle, the req.uired fuel

entry conditions from the divertor were found. These

condition_ were required to match the transport power

available from the reactor. The 1-D model revealed that

the propellant exit velocity (V_) was approximately

twice the throat (sonic) velocity (V,h,oat), shown in

equation (117) 42.

Vex 2 V,h,.o,,,

I,p - - (17)

gc g.

The related throat state variables (where T,h,_.,, P,h,o.,,

(PR)a_mt, B_t, and Ath_, are the temperature, density,

pressure, magnetic field, and cross-sectional area in the

throat, and R is the gas constant using monatomic

hydrogen (1.0078 AMU)) are given by equations (18)

through (21):

T,n,_ , - V'h_" (18)

yR

(PR)a,,o,,,

F

2y Albion,

(19)

(PR),h,_,

P:,,.o,,, - (20)

RT, h,,,,,,

B,h,,,,,, = _[21.to(PR),n,.,_, (withfl= l) (21)

The total thrust (F) was previously calculated to be

5,567 lb,.. An "effective" 10 cm throat radius was

created by the magnetic field lines, with the structural

throat radius at 2 m. The 10 cm radius was chosen so

that the throat cross sectional area power flux of

1.56"10 s MW/m 2 was comparable to proposed NTR

designs (1.16"I0 s MW/m2) 46 and within extrapolation

of current cryogenic engines such as the Space Shuttle

Main Engine (5.14'104 MW/m :) 47. A 13= ! condition

was assumed to calculate the magnetic field strength

and required current for throat and upstream reservoir

magnet (oils. A specific heat ratio (y) of 5/3 was used

for a mo]tatomic gas of three degrees of freedom.

rhe reservoir conditions were solved using

equations (22) and (23), and equations (20) and (21)

with reservoir values (M is throat Mach number - 1).

The T_, _vas 329 eV, an order of magnitude lower than

20
Prop Flow Rate = 0.0601 kg/sec

\,

\.

Reservoir

Temperature = 329 eV

Pressure = 4.9 E+05 N/m 2

Density = 9.2 E+21/m 3

Magnetic Field = 1.I T

Exi._.._t

Flow Velocity = 397 km/sec

Isp = 40,485 lbf secflb m

Total Flow Rate = 0.0624 kg/sec

Divertor

Temperature < 4,000 eV

Density (ion) < 2.0 E+20/m 3

Fuel Flow Rate = 0.0023 k_sec

........ ??-

Throat

Temperature = 247 eV

Pressure = 2.4 E+05 N/m 2

Density. = 6.0 E+21/m 3

Magnetic Field = 0.8 T

Flow Velocity = 199 km/sec

Figure 14: Plasma State Conditions Through Magnetic Nozzle

that at the scrape off boundary. The Br,_, similar to

nthmat, WaS a comparatively modest 1. ! T.

charge exchange, and other losses. Figure 14 illustrates

the plaSma state variables through the nozzle.

(M = 1) (22) P.. = mtota • ! h.. = m,ol,,i " Y RT.,

y-1

(24)

y

(PR).s= 1 + M2 r I(pR),h,.o,,, (M = 1)

(23)

1 . 2

P.. = rb P._ = rlj 2 mY"etv a'_`

(25)

1 . 2

= -_ mtota/Vres

(26)

The reservoir flow power (PJ and thermal

velocity (v_,_), which are related by the stagnation

enthalpy (hr_), can be estimated by equations (24) and

(25). The flow power can then be related to the

available transport power (Pt_) from the reactor by

equation (26). Thus the jet power and state conditions

throughout the nozzle can be correlated to the power

delivered to the reservoir by the divertor. The 20% loss

aSsociated with the enthalpy transfer was aSsumed to

exit the nozzle in the form of neutrals, ions with

velocity vectors not strictly aligned with the thrust

vector, collisions with support structure, line radiation,

The augmentation propellant maSs flow rate (th p_oo)of

0.06 kg/sec was eaSily found by subtracting th ¢o,jfrom

thtot, j . The augmentation propellant waS - 27 times

that of the fuel, emphasizing that a significant portion

of the magnetic nozzle system must be dedicated

towards accommodating the injection of augmentation

propellant. This great a mass infusion into the flow

stream (compared to the relatively low flow rate from

the reactor) may be great enough to adversely affect the

plaSma conditions back into the reactor. This could

represent a significant operational problem and

warrants future assessment.

21
POWER CONVERSION

The primary function of the power conversion

system was to utilize the Bremsstrahlung and

synchrotron radiation by "thermalizing" the energy flux

so that it could be converted into electrical output

power. A closed Brayton (gas) thermodynamic cycle

was selected on the basis of proven design and

fabrication experience, significant performance and

reliability database, and the ability to use an inert gas

(gaseous helium (GHe)) as the working fluid to

transport heat energy (cooling the reactor first wall and

TF coils) at high temperature (up to 2,000 deg K)

directly to the turbine.

As shown in the power flow diagram (figure

15), two sources of reactor radiated power of 1,823

MW were at 50 keV and 2.54 MeV. In addition, 259

MW of waste heat from the N-NBI was also recycled as

thermal energy. Thus from the total input heat of 2,082

MW, 400 MWe of electrical power was produced for

the N-NBI and other on board power requirements. The

remainder of 1,682 MW was thermally radiated to

space via the heat rejection system.

To neutralize gyroscopic torque generated by

the power generating system, two 200 MWe

"turbosets" were used (figure 12), configured to rotate

in opposite directions. The conceptual design of the

power conversion system was patterned after helium

turbine designs for nuclear power plants advanced in

Germany three decades ago 4s'49. Some units, however,

were actually constructed up to 25 MWe power levels 5°.

"l-he salient design and performance

parameter_ for the two power conversion systems were

regenerated cycles with >80% effectiveness (eRo)

operating with a peak temperature (turbine inlet) of

1700 K. The cycle temperature and pressure ratios of

2.52 and 2.15, respectively, established the need for

each unit to accommodate half the input heat 1,041

MW while generating the required output of 200 MWe.

This set the required cycle efficiency at approximately

20%. With the peak cycle pressure (i.e. compressor exit

pressure) set at 25 atm, the resultant inlet pressure was

11.7 arm, and the total mass flow rate for each unit was

437 kg/sec.

A flat plate, square radiator design was chosen

using light weight carbon-carbon, parallel duct heat

To

Parallel

CCGT Unit_ [

)

3600 RPM

T = 1700K

P = 24.5 at

:'_Re zcto(

Tucbino

T = 1327K

P = 12.2 at

Ti 1256K

P = 24.8 at

tl

I l

Alternator

RPM

I

T = 675K I

P= 11.7at_

I

"I= 954K

IF =25at

;_.11==437 kg/sec

I I I I I

I I I I

I I I I

I I I I

l[ ! | I

Radiators

Area = 78,540 m2

I I I I

I I I I

I I I I

I I I I

I I I I I

T = 1025K

P = 12.15 at

Figure 15: Single 200 MWe Closed Cycle Gas Turbine Power Conversion System

22
pipes. This radiator technology has been proposed as

the appropriate match for closed Brayton cycle with a

high temperature gas reactor heat source 51.The radiator

parameters were derived from successful laboratory

demonstration of specific mass up to 1.5 kg/m _ (based

on radiation from both sides), surface emissivity up to

0.9, and radiating temperatures up to 650 deg K. 5_With

GHe entering the radiator at 1,025 deg K, the total

radiating area for each of the turbosets was 78,540 m 2,

necessitating a radiating area of two 150 m square

planar heat rejection surfaces, each rejecting heat from

both sides. Thus for the two units, four radiating

surfaces arranged in a "cruciform" configuration was

required. Although this configuration resulted in

radiative interchange between the surfaces, the overall

view factor to space was only decreased from a value of

1 to 0.8. The radiator area calculation has been adjusted

to account for this penalty, as well as additional small

waste heat dumps.

Table 7: Power Conversion System

Mass Properties

Recuperator 23

Compressor 20

Turbine 9

Alternator 23

Conditioning 40

Radiator 236

Power conditioning radiator 30

Total (mt) 381

The turboset casing diameters were 2.6 m and

the length, including the alternator-generator were-18

m. With the turboshaft rotational speed of 3,000 to

3,600 rpm, the alternator could be driven at

synchronous speed to generate 50 to 60 Hz AC power.

It is advantageous to generate most of the power at the

voltage demanded by the largest loads. The power

consumed by the N-NBI (367 MW) represented over

90% of the total power output of the power conversion

system. As a result, the power would be generated in

the appropriate form (voltage, amperage, phase) for the

N-NBI in order to minimize heavy power conversion

transformers and associated systems. Almost 90% of

the N-NBI power was required in the form of high

voltage (3 steps of 150-170 kV each) s4and respectable

current (64 A). Voltage transformers, however, were

included in the power management and distribution

contribution to the total mass budget of 381 mt for the

power conversion system (Table 7).

NEUTRAL BEAM INJECTOR

Negative ion neutral beam injection (N-NBI)

appears to be the most promising method of noninductive plasma heating, at high number density, due

to its greater neutralization efficiency than positive ion

NBI 52. One of the leading, operational N-NBI is part

of the Japan Atomic Energy Research Institute's

(JAERI) JT-60U reactor. It was this N-NBI that was

used as the basis for scaling a system capable of heating

the non-bootstrap driven current of the design concept.

The JT-60U N-NBI is capable of providing 500 keV

negative D ions, at n = 5"1019/m3, with a total beam

power of 10 MW 52

An injected power of 108 MW was shown to

be required from the N-NBI to heat the non-bootstrap

(seed) current of 4.35 MA, an order of magnitude

greater power than that delivered to the JT-60U. The

system scaling approach taken was to retain the same

system efficiencies, increase by an order of magnitude

the input ion power from the ion sources, and adjust the

N-NBI mass properties to reflect the space vacuum

environment and additional ion source tanks.

To supply 108 MW of neutralized D ion beam

heating power, 367 MW of power had to be supplied to

the N-NBI. These considerable values resulted even

after a deliberate effort was made to minimize the

heating power required by maximizing the bootstrap

current fraction. The various components of the N-NBI

power consumption are illustrated in Table 8, and were

scaled up from the JT-60U. Almost 90% of the power

was needed by the acceleration power supply (315

MW), the system upon which the power conversion

system would be designed to accommodate. The

significant power losses were associated with the ion

source tanks (37 %) and the neutralizer (24%) and

represent the majority of the waste heat (259 MW) that

was sent to the power conversion and rejection system.

Table 8: Neutral Beam Injector Power Usage

Cathode 3

Arc 15

Bias < 1

Plasma grid filter 2

Extraction 16

Acceleration 315

Bending coil 16

Total (MW) 367

One of the chief concerns in such a system

scale-up was whether an order of magnitude increase in

23
beamline power density could be achieved. The JT-60U

configuration was thought to be able to accommodate a

power density increase up to a factor of two and at most

four s3. Current research at JAERI on a 1 MeV N-NBI

device is focused on increasing current density to these

levels _. It is also pursuing a merged beam extractionacceleration system with a shorter length, smaller

diameter, multiple channel neutralizer, enabling a more

compact (thus less massive) design. It is anticipated that

this research could eventually enable the N-NBI

postulated for this design concept.

Table 9: Neutral Beam Injector

Mass Properties

Ion sources (20 @ 2.5 mt each) 50.

Ion source steering mechanism 1.

Beam scraper 0.6

Cryopumps 1.

Beam limiter 0.4

Ion source tank 1.

Bellows 0.6

Neutralizer 0.5

Bending coils 4.

Ion dumps (positive & negative) 1.2

Ion dump tank 1.

Calorimeter 1.5

Beam limiter 0.4

NBI bellows 0.3

Isolation valve 0.5

NBI port 0.6

bellows 0.2

Total (mt) 64.8

The total mass of the N-NBI was 65 mt. The

N-NB1 mass properties are given in Table 9. A total of

20 ion source tanks (at 2.5 mt each) dominate the NNBI mass. These 3 m diameter, 4.8 m high structures

produce the 22 A, 500 keV negative D ions by passing

cesium through an arc discharge, then on through a

transverse magnetic field to enhance negative D ion

yield, to a three stage multi-aperture extractoraccelerator that electrostatically increases the ion's

energy to 500 keV sz, _4.Its total mass was estimated by

accounting for an order of magnitude increase in

required ion yield coupled with deletion of structure

associated with vacuum generation 53.Other significant

mass reductions from the JT-60LI device included

removal of hardware providing vacuum conditions in

the ion source and dump tanks (from 25 to 1 mt each),

and length shortening-driven reductions of the

neutralizer (from 6 to 0.5 mt) and NBI port (2 to 0.6

mt) due to test site physical layout rather than

experimer_ tal requirements19.

"Ihe N-NBI was placed within the aft central

truss, adjacent to the reactor to shorten the beam line

length. The entire system length was re-scaled to be

approximately 32 m. Injection of the neutral beam at

right angles to the reactor plasma flow has the potential

to depolarize spin polarized fuel 15.To avoid excitation

of magnetic fluctuations about whatever ion cyclotron

frequencies might be encountered, one or more right

angle turns in the beam line conduit may be required to

mitigate this effect.

PROPELLANT CRYO-TANKAGE

The slush hydrogen propellant cryo-tankage

was based on a pre-existing conceptual design _, itself

predicated on operational or previously designed

conceptual liquid hydrogen propellant tanks sS. The

large quantity of propellant needed for the vehicle

rendered even the largest past, current, or proposed

liquid hydrogen tank designs unsuitable. The largest

liquid hydrogen tank ever built and flown is the

hydrogen component of the Space Shuttle's External

Tank (ET), a 27.5 ft (8.4 m) diameter, 96.7 ft (29.5 m)

long structure which can accommodate 103 mt of

propellant 56.The Saturn V's S-II hydrogen tank, a 33 ft

(10 m) diameter, 53 ft (16 m) long structure

accommodated 70 mt of propellant 57. In order to carry

the signit icant quantity required for the design concept

(1,292 m0, multiple tanks of an even greater capacity

were obviously needed. Therefore, a conceptual design

was made of the largest tank that appeared reasonable

on the grounds of experience, engineering judgment,

and ground transportation concerns. An existing

conceptual design of liquid hydrogen tankage done in

part by t ae co-authors _ and based on an NTR vehicle

concept s', was adopted for slush hydrogen tankage.

Support systems, net mass, and power required to

accommodate slush hydrogen above and beyond those

for liquid hydrogen were for the most part negligible sS.

One concern, however, was allowing sufficient

additional tank volume and strength to accommodate

the slush that would liquefy during launch and up to

docking with refrigeration systems on the vehicle

concept. Due to its ascent dependency and lack of

launch s'enario definition, no design provisions were

made fo:' this concern at this time. The original NTR

tankage design was for a human Mars mission. A

graphite-epoxy (GrEp) composite hydrogen tank

material was used to obtain considerable mass savings

over advanced aluminum alloys _. The 10 m diameter

was maintained in order not to significantly impact

manufacturing and ground transportation limits. The

24
3:1 aspect ratio (cylindrical barrel section length-todiameter) yielded a total tank length of 37 m. The net

slush hydrogen propellant available for main impulse

per tank was calculated to be - 207 mt.

Table 10: Cryo-Tankage Mass Properties

Structure 16

upper ellipsoidai dome

cylindrical barrel section

lower ellipsoidal dome

forward skirt

ring

cylinder

aft skirt

ring

cylinder

insulation

Fluid system < 1

feedlines and insulation

valves and manifolds

autogenous bleed

zero g vent

purge

Electrical and power system 2

avionics and power

instrumentation

telemetry

range safety

shielding

harnessing

Reaction control system 2

500 Ibf (2) thrusters

50 lbf (24) thrusters

propellants

Interface hardware 2

OMV-derived interface

tank attach central truss

..................................................................

Total (mr) 22

Table 10 lists the subsystem masses for the

designed cryo-tankage, which totaled to a dry mass of

22 mt. The subsystems were based on and modified

from three studies, all of which were based primarily on

the two Centaur upper stage configurations flying on

the then-current Atlas/Centaur and Titan IV/Centaur

expendable launch vehicles. In the resulting tank

design, over 46% of the total mass was attributed to the

composite tank (upper and lower domes, plus barrel

section), and 22% attributed to the 5 cm multi-layer

cryo-insulation plus aluminum micro-meteoroid

shielding (sized for Mars orbit "mean space

temperatures" to approximate the range of Earth to

Saturn environment). A significant amount of mass and

system complexity was required by considering the

means for delivering the propellant to the departure

orbit via an HLLV. The fully loaded propellant tankage

must be launched with all the requisite ETO-related

systems (avionics, telemetry, range safety, etc.). Upon

separation from the HLLV, the cryo-tankage must be

capable of controlling its own attitude, rendezvous, and

dock with the vehicle concept, thus requiring a reaction

control system and interface hardware.

Table 11: Cryo-Tankage Mass Summary

Propellant

Stage Dry

Adapter

main impulse 207

flight perf. reserve 2

residuals/losses 6

Contingency (30% of dry mass)

........................................

GLOW (mt)

215

22

6

8

251

Table 11 illustrates the mass summary for the

fully loaded cryo-tankage in its ETO launch

configuration (i.e., including adapter and associated

contingency). Added to the useable, main impulse

propellant was a flight performance reserve (FPR) of

1% (of main impulse propellant), consistent with past

mission experience to accommodate in-flight

dispersions. Estimates of residuals and chill-down

losses were also included and made up 3% (of the total

tankable propellant). The adapter was sized to

accommodate launch loads and was chargeable to the

propellant tankage payload, although it would only be

used from ETO and not retained on the interplanetary

vehicle concept. A 30% weight growth allowance was

assessed on the tankage dry mass and adapter. The

gross liftoff weight (GLOW) of a fully loaded tankage

payload would be 251 mt.

REFRIGERATION

Two Garret, reverse Brayton refrigeration

systems with helium working fluid were used for

cooling the propellant/fuel tankage and the

TF/PF/divertor/magnetic nozzle superconducting coils.

The first system used a ! 4 deg K low temperature sink

(TL) to accommodate the slush hydrogen's boiling

point. A second system was used to cool the YBCO

superconductor in the coils, where TL was set to 70 deg

K. Even though other higher temperature, more

25
appropriate refrigeration systems (neon, nitrogen, etc.)

might have been used for the coils, commonality with

the propellant/fuel system was pursued for the sake of

simplicity.

The electrical power (P_t,g) required to operate

the system was calculated by equation (27), where the

heat rejection temperature (Tu) was chosen to be 350

dog K, and was only a weak function of the quantity of

heat rejecting for the values of interest. The

refrigerators' coefficient of performance (COP) was set

at 15% of the Camot COP, consistent with cooling

loads above 500 W 59. Using a Mars orbit heat flux to

approximate an average Earth to Saturn thermal

environment, and the tankage insulation and shielding

described earlier, the total thermal power to be removed

(Q) from the seven propellant/fuel tanks was 0.923 kW.

The heat to be rejected from the coils (produced by the

neutron power flux intercepted by the blankets and not

removed by the GHe heat transfer fluid, plus ancillary

heating) was not explicitly calculated however. Instead,

an arbitrary 500 W of residual heat was assumed to

reach the outer aluminum-lithium coil surface. The total

thermal power to be removed from the 12 TF, 7 PF, 3

divertor, and 3 magnetic nozzle coils was 12 kW. Thus

the total electrical powers required for the tankage and

superconducting coil refrigeration systems were < 0.15

MWe and 0.32 MWe respectively.

P'_/';g - 0.15 (27)

The mass of the refrigeration (M,ang) system

was determined from equation (28) and included all

subsystems except power source and radiator 59. The

additional radiator mass was determined to be small

compared to the main array. Thus, the total

refrigeration power (operation and rejected heat) to be

rejected was merely added to the main array. The total

masses for the tankage and superconducting coil

refrigeration systems were 1.4 and 52.4 mt respectively.

Mref_tc = 91.9 * l O°°46s_l°g(O))_" I T28:L 1

(28)

AVIONICS /COMMUNICATIONS

The concept vehicle's avionics suite would be

composed of primary/backup computers, guidance,

navigation, and control system (GN&C), tracking, data

display, sequencing, and instrumentation. The

exponential rate of growth in speed and capability of

electronics and computer technology will no doubt

enable future systems to be vastly superior to today's

systems at only a fraction of the mass and power

required. Thus minimal attention was devoted towards

trying to extrapolate avionic system capabilities and

requirements. A sufficient avionics suite of the future

should be available with a total mass < 1 mt with an

arbitrarily small power consumption.

The communication system was presumed to

be a derivative of the recently developed Ka band (20

to 30 GHz) technology. This NASA LeRC digital

processing and storage communication technology is

currently capable of integrated data, voice, and video

with typical throughput rates in excess of 108 bits/see

and up _o 101° bits/sec 6o.6t.62 Two 15 m diameter

deployabie Ka band dish antennas were used for the

audio, video, and data communication system. Dual

antennas were used to enable two way simultaneous

communications with the departing and arriving

destination planets. Mounted on 18 m truss booms aft

of the crew habitat payload, the dish antennas were

positioned at 45 dog angles with respect to the radiator

arrays to minimize heating (figure 16).

The total avionics/communication system

mass was set at 2 mt. The power for the avionics suite

was assumed to be comparable to the Space Shuttle

orbiter available power 630.02 MW. The power required

by the communication antennas was indeterminate,

since the planetary antennas' diameters and powers

were undefined. Therefore a value was set at 0.2 MW.

STAl_T/RE-START REACTOR & BATTERY

I'he startup system consisted of a 1 MWe

nuclear fission reactor power system and a nickel

hydrogen (NiH) bipolar battery bank. Weeks prior to

departure, the startup reactor was used to gradually

refrigerate the TF/PF/divertor/magnetic nozzle coils,

initiate and ramp up their current, provide auxiliary

power, m_d charge-up the battery. When all systems and

crew wele ready, a - 10 sec plasma startup sequence

was initi:tted with battery discharge providing the - 1

GJ ofemrgy and 100 MW of power needed for startup.

._ior to current startup and neutral beam

heating to ignition conditions, _ 1 to 2 MW of auxiliary

radio-frequency (r0 heating at the electron cyclotron

frequency was used to create a small volume of high

conducti',ity plasma (T_100 eV and n_ _1019/m 3)

outboard of the plasma major radius which assisted in

the curre_at startup process 64. This plasma conditioning,

referred _o as preheating, permitted a small radius (ao

0.2 to 0A m) current channel to be established with a

relatively low initial loop voltage (<25 V as opposed to

26
Figure 16: Dual Ka Band Antennas and Crew Payload

-100 - 200 V without rf assist). With the onset of

current initiation and establishment of the desired

safety factor q in the small current channel, the startup

major radius near the outboard midplane was gradually

shifted inward to R = 2.48 m. During this "expanding

radius startup ''65, new layers of plasma were added to

the warm core through ionization of a regulated gas

feed. Major radius compression permitted minor radius

expansion and a simultaneous increase in plasma

current while a constant q was maintained. As the

plasma minor radius grew in size, sufficiently high

levels of current and plasma density were achieved to

ensure adequate confinement of energetic protons from

D3He fusion. Neutral beam heating to ignition

conditions could then commence.

A high temperature gas cooled reactor (at - 1

MWe) power system would supply auxiliary/standby

power during emergency re-start of the main propulsion

system. The reactor heated 1.5 kg/sec flow of GHe to

1,500 K in order to drive a gas turbine. Designed for

minimum mass, total system was estimated at < 5 mt.

The NiH bipolar batteries were derivatives of

devices designed and tested at NASA LeRC. These

82% efficient, high peak power systems were capable

of specific energies of 180 kJ/kg and energy packaging

densities of over 80 Whr/liter 66,67. For the required

startup energy pulse, a bank of volume 3.36 m 3 with a

mass of 5.6 mt would be needed. State of the art high

energy capacitor-batteries 68 were also considered, but

despite their attractive energy densities, their specific

volumes were not competitive with NiH batteries.

Table 12: Power Usage Summary

(Nominal and Re-start Battery Re-charge)

NBI 367.

RCS 32.

TF/PF/div/mag noz (req'ed during re-charge) 0.320

Prop/fuel tankage ...... 0.148

Battery re-charge ...... 0.278

Communications ...... 0.2

Payload ...... 0.03

Avionics ...... 0.02

Fuel injector 0.004

Total (MW) 400.

Should the fusion reactor need to be re-started

during the interplanetary transit, the same startup

sequence would be followed. The startup reactor power

would be used to maintain the refrigeration to the

TF/PF/divertor/magnetic nozzle coils (to prevent them

from "going normal") and propellant/fuel tankage, and

to provide for crew payload accommodations,

communications, and maintain vehicle avionics. Should

the battery bank fail to re-start the reactor, sufficient

startup reactor power would be available to re-charge

the bank in 35 minutes for another attempt while

maintaining power to the other essential systems. Table

12 illustrates the essential power requirements during

the re-charge of the re-start battery bank, as well as a

summary of the nominal power usage.

27
REACTION CONTROL

Positive control of a vehicle's attitude requires

arresting various torques associated with natural

perturbations as well as applying torques to achieve

vehicle orientations and assist in steering. Most of the

significant natural perturbations present in planetary

orbit (such as gravity gradient-driven apsidal and nodal

drifts, atmospheric drag, magnetic field interactions,

etc.) are either significantly diminished or not

applicable to interplanetary travel. The torque

associated with solar radiation pressure was assessed

due to the considerable surface area of the radiator. It

was found that even the worst case orientation at Earth

solar distance, the radiation torque was negligible.

Main propulsion steering assistance would,

however, be needed. The single nozzle configuration

would not be able to perform roll control. Pitch and

yaw could conceivably be accomplished by selective

configuring of the nozzle's magnetic field geometry 69.

This would appear to be the viable way to steer the

vehicle using the nozzle, given that the complex

coupling between it and the reactor diverter would

make gimballing of the nozzle unlikely. However, since

reaction control was difficult to assess without a vehicle

system control model, and the need for roll control, and

the yet to be defined nature of selective configuration

of the nozzle's magnetic field geometry, a separate

reaction control system (RCS) was added to the design

concept.

RCS propulsion technology for today's

satellites is transitioning from N2H 4 monopropellants to

more advanced technology electric resistojets and

arcjets. Since the vehicle concept was significantly

more massive than today's spacecraft (requiring more

impulse) and since large quantities of electric power

and hydrogen would be available, the higher I,p

hydrogen fueled arcjet was chosen for the RCS.

Since a vehicle system control model was not

available to guide the design of the RCS from a "needs

up", a top down approach was pursued to determine

how much control authority would exist for a

reasonable impulse, power, propellant, and mass

allocation. An initial estimate of 500 ibf per engine,

requiring an excessive 50 MW input power per engine,

was sealed on the mass and thrust of the current Titan

IV/Centaur RCS system. Re-estimation for 8

MW/thruster (i.e. two jets firing per couple, with one

fore couple and one aft couple) yielded a thrust level of

over 80 lbr per thruster, implying a longer firing time to

deliver the same impulse. The RCS thruster parameters

were calculated based on the NASA LeRC 30 kW

hydrogen arcjet development program. That program

produced thrusters delivering lso's up to 1,460 lbf

see/Ibm and efficiencies of 30+% 70.Using these values,

mass and mass flow rates per thruster were extrapolated

to be 2 mt and 0.0235 kg/sec respectively.

The RCS was composed of two units of twelve

thrusters each. Each unit housed four thruster clusters

90 degrees apart. Each cluster contained three thrusters,

each aim,_d at right angles to the others. Full six

dimensions of freedom control was provided with both

units firing simultaneously. One unit was mounted on

the aft end of the truss network, forward of the reactor.

The other unit was forward of the propellant tanks. It

was decided not to place the forward unit at the front of

the vehicle since 87% of the vehicle's fully loaded

mass (and 78% of its dry mass) was aft of the radiators.

An angular velocity of 1.6 deg/min could be

produced by four RCS thrusters firing about the

minimum axis (roll) with a torque of -2,200 N-m and a

firing time of one minute. For a pitch or yaw maneuver

associated with steering, the greater torques (22,780 Nm) due te larger moment arms (mitigated somewhat by

the greater moment of inertias about those axes)

required only 20 seconds of firing time to produce the

same angular rate. In both instances, the angular

velocity and the f'wing times appeared reasonable. The

total propellant consumption, assuming five 60 second

start/stop corrections per day, every day (up to one

year) was almost 21 mt. The total mass of the 24

thruster system was 48 mt, assuming all power

processing was performed by the primary power

conversion system.

WEIGHT GROWTH CONTINGENCY

Weight growth contingency is a margin

allocatec to compensate for the inevitable growth in

mass experienced by aerospace systems as designs

mature and construction takes place. Underestimates

become apparent and technical problems are solved by

incorporating solutions requiring additional mass. All

new launch vehicle development programs carry such

an allocation, though the percentage allowable varies as

a functi_ _n of component maturity. Experience with the

developnent of eighteen major aerospace vehicles has

demonstrated that from the point of initial contract

proposai through acquisition of first unit, the total

average weight growth experienced by military

aerospace vehicles has been 25.5% 71. For more than a

half dozen major NASA manned and unmanned

vehicle., and spacecraft, from the point of phase C/D to

first vehicle flight, most programs have experienced a

similar 20% to 30% weight growth 7z. Given the

immaturity of the overwhelming majority of this

technol)gy, past aerospace experience would suggest a

pruden_ minimum value of 30% weight growth

28
allowance be assessed on the total dry mass of the

propulsion system (M_). Thus, a weight growth

contingency of 345 mt was carried, representing a

major mass property component.

FUTURE AREAS OF ANALYSIS

Due to time constraints, several potentially

superior design solutions were not pursued. Further

study is expected to be performed and possibly

incorporated into a future upgrade of the design

concept. Some potential design alternatives include:

inclusion of co-axial helicity injection to reduce the

requirement for N-NBI and a significant power

conversion system; reduced absorption of waste

radiation together with a more efficient power

conversion system to reduce radiator size and mass; a

more rigorous heat transfer analysis; assessment into

the feasibility of gray body heat rejection from the LiH

blankets directly to space; replacement of most of the

RCS by altering the strength and geometry of the

magnetic nozzle's field; and a liquid nitrogen or neon

refrigeration cycle for cooling of the TF/PF coils.

CONCLUSIONS

A conceptual vehicle system design predicated

on a small aspect ratio spherical torus nuclear fusion

reactor has the potential for enabling relatively fast

outer solar system travel. The requirements for a human

mission to Saturn were satisfied with a 108 mt payload

mass, a 235 day one way trip time, and a IMLEO of

2,941 mr. The same concept was found to deliver the

same payload to Jupiter requiring a 150 day one way

trip time. Both missions presuppose the availability of

in situ planetary refueling capability. High orbit space

basing was assumed, but not required. A requirementdriven approach provided guidance during the design

of the concept, though much of the design decisions

were driven as much by the desire to maintain a balance

between performance capability and reasonable

extrapolations from current technology expectations. A

complete vehicle concept was produced with analysis

to a sufficient level to make certain assessments on

general concept viability and research requirements.

Systems analysis, design, and assessment was

performed on in-space operations, payload, central

truss, nuclear reactor (including divertor and fuel

injector), magnetic nozzle, power conversion (turbine,

generator, radiator, conditioning), neutral beam

injector, refrigeration, tankage, avionics, start-up

reactor and battery bank, communications, and reaction

control systems. Detailed assessment was performed on

reactor operations, including plasma characteristics,

power balance, power utilization, and component

design. Overall feasibility of nuclear fusion propulsion

systems must include assessments of support systems

such as heavy lift launch vehicles, space based orbit

transfer vehicles, and in situ resource utilization with

associated cost of their operations. Critical areas of

research upon which the feasibility of this concept rests

include the demonstration of ignited, long term steady

state, D3He fueled nuclear fusion reactor operation, the

successful incorporation of a divertor able to transfer

large quantities of transport power to a propulsion

system, the determination whether large downstream

propellant infusion adversely impacts fusion reactor

operation, the identifying of operations incompatible

with spin polarized fuel, and the successful

demonstration of a high thrust magnetic nozzle. Once

such issues are addressed, a judgement can be rendered

as to the practicality of a solar system-class, nuclear

fusion-based transportation system of the 21 stcentury.

ACKNOWLEDGMENTS

The authors wish to thank many who have

been very helpful in providing guidance and expertise:

to Gerald Hale of Los Alamos National Lab (LANL) in

the area of spin polarized D3He fuel, to Stanley Kaye of

Princeton Plasma Physics Lab (PPPL) in the area of

fusion confinement time scaling laws, to Joseph Warner

of NASA Lewis Research Center (NASA LeRC) in the

area of YBCO superconductor material properties and

prospects, to Ronald Moses of LANL in the area of Dshape magnet designs and forces, to Lawrence Green of

Westinghouse Science and Technology Center and

Daniel Driemeyer of Boeing Co. in the area of ITER

divertor design drivers, to Peter Turchi and Hani

Kamhawi of Ohio State University in the area of

magnetic nozzle design and state conditions, to Larry

Grisham of PPPL and Masaaki Kuriyama of the Japan

Atomic Energy Research Institute in the area of

negative ion neutral beam injectors, to Richard Kunath

of NASA LeRC in the area of Ka band space

communications, to Robert Cataldo of NASA LeRC in

the areas of start-up NiH batteries and Mars human

payload systems; to John Miller of JME, Inc. and Mary

Ellen Roth of NASA LeRC in the area of high energy

battery capacitors, to John Sankovic of NASA LeRC

and R. Joseph Cassady of Primex Technologies Corp.

in the area of high power hydrogen arcjets, to Meilissa

McGuire of Analex Corp. in the area of reaction

control, to Judith Watson of NASA Langley Research

Center in the area of Space Station truss network, and

to Mohamed Bourham of North Carolina State

University and Stanley Milora of Oak Ridge National

Lab in the area of fuel injectors.

29
REFERENCES

t Williams, C.H. and Borowski, S.K., "An Assessment

of Space Propulsion Fusion Concepts and Desired

Operating Parameters for Fast Solar System Travel",

AIAA paper 97-3074, July 1997.

2 Williams, C.H. and Borowski, S.K., "CommerciallyDriven Human Interplanetary Propulsion Systems:

Rationale, Concept, Technology, and Performance

Requirements", paper # 58, 13* Symposium on Space

Nuclear Power and Propulsion, Albuquerque, NM,

January 1996.

3 Williams, C.H., "An Analytic Approximation to Very

High Specific Impulse and Specific Power

Interplanetary Space Mission Analysis", NASA TM

107058, NASA Lewis Research Center, Cleveland,

OH, September 1995.

4 Borowski, S.K., "A Comparison of Fusion /

Antiproton Propulsion Systems for Interplanetary

Travel", AIAA paper 87-1814, July 1987.

s Williams, S.N., "An Introduction to the Use of

VARITOP", Jet Propulsion Lab, JPL doc. D-11475,

January 1994.

6George, J., NASA Lewis Research Center, personal

communication, June 1997.

7 Hyde, R.A., "A Laser-Fusion Rocket for

Interplanetary Propulsion", 34 thInternational

Astronautical Federation, Budapest, Hungary, October

1983.

s lsakowitz, S. J., "International Reference Guide to

Space Launch Systems", AIAA, 1991.

9 Borowski, S. K. et al., "Nuclear Thermal

Rocket/Vehicle Design Options for Future NASA

Missions to the Moon and Mars", NASA TM-10707 i,

September i 995.

_oKuicinski, G. L., "Importance of Helium-3 for the

Future", Report of the NASA Lunar Energy Enterprise

Case Study Task Force, NASA TM-101652, July 1989.

_ Hoffman, S. J. and Kaplan, D. I., editors, "Human

Exploration of Mars: The Reference Mission of the

NASA Mars Exploration Study Team", NASA Special

Publication 6107, July 1997.

_2Watson_ J., "Design Considerations for an Astronaut

Monorail gystem for Large Space Structures and the

Structural Characterization of its Positioning Arm",

masters thesis (engineering mechanics), Old Dominion

University, August 1992.

_3Borowski, S. K. and Strickler, D. J., "A ProfileAverage& Steady State Model for Studying Ignitedand Driven-Tokamak Operations", unpublished ORNL

paper, 198,5.

14Hale, Gerald, Los Alamos National Lab, personal

communication, February 1998.

t5 Mitarai, O., et al., "Spin Polarization Effect on

Ignition Access Condition for D-T and D3He Tokamak

Fusion Reactors", Fusion Technology, July 1992.

_6Christiansen, J.P., et al., "Global Energy

Confinement H-Mode Database for ITER", .Nuclear

Fusion, vol. 32, no. 2, 1992.

_7Kaye, S.M., et al., "ITER L Mode Confinement

Database", Nuclear Fusion, vol. 37, no. 9, 1997.

_sMcCooi, S.C., et al., "A Mega-Ampere Spherical

Tokamak Design for Beta-Limit and Confinement

Studies", Transactions of Fusion Technology, vol. 27,

April 1995.

m9Grishma, L., Princeton Plasma Physics Lab, personal

communication, March 1998.

soAlava, M.J., et al., "Bootstrap Drive of D- 3He

Tokamak Reactors", Proceedings of the 5'h

Internati¢_nal Conference on Emerging Nuclear Energy

Systems, 1989.

2_Devote, R.S., et al., "Projections for a Steady State

Tokamak Reactor Based on ITER", UCID 21519,

December 1989.

22Wittenberg, L.J., et al., "Lunar Source of 3He for

Commer,;ial Fusion Power", Fusion Technology,

Septemb ;r 1986.

53Servio:, R. F., "New Superconductor Stands Up to

Magnetit Fields", Science May 1995.

24Moses: R.W. and Young, W.C., "Analytic

Expressions for Magnetic Forces on Sectored Toroidal

Coils", _Proc. Eng. Probl. Fussion Res., 6 th, IEEE, 1976.

30
zsTeller,E.,editor,Fusion,vol. 1, Academic Press Pub.,

1981.

26Boyer, R., et al., editors, "Materials Properties

Handbook: Titanium Alloys", ASM International Pub,

1994.

27Roth, J.R., Introduction to Fusion Energy, Ibis Pub.,

1988.

z8 Baker, C.C., et al., "Technology Implications of

Advanced Fusion Fuel Cycles", Proc. International

Conf. On Plasma Science, IEEE, 1979.

29Davis, J.R., et al., editors, "Metals Handbook", vol. 2,

ASM International Pub, 1990.

3oLaMarsh, J.R., Introduction to Nuclear Engineering,

Addison-Wesley Pub., 1977.

31Frattoliilo, A., et al., "Development of a Two-Stage

Pneumatic Repeating Pellet Injector for the Refueling

of Long-Pulse Magnetic Confinement Fusion Devices",

Fusion Technology, December 1997.

32Combs, S.K., et al., "Pellet Injector Research and

Development at ORNL", Fusion Technology, 1994.

33Kincaid, R.W. and Bourham, M.A., "Electrothermal

Plasma Gun as a Pellet Injector", Fusion Technology,

1994.

34Bourham, M.A., North Caroline State University,

personal communication, June 1998.

3s Milora, S.L., Oak Ridge National Lab, personal

communication, June 1998.

36Carpenter S.A. and Deveny, M.E., "Mirror Fusion

Propulsion System (MFPS): An Option for the Space

Exploration Initiative (SEI)", Space Technology, voi.

14, No. 1, Elsevier Science Ltd. Pub., 1994.

37Post, D. et al., "The ITER Power and Particle Control

System", Fusion Technology, December 1996.

38Araki, M., et al., "Analytical Study of

Electromagnetic Forces Induced By Eddy Current in

ITER Divertor", Fusion Technology, December 1996.

39Sevier, D.L. et al., "Implications of Steady State

Operation on Divertor Design", Fusion Technology,

December 1996.

,o Green, L., Westinghouse Science and Technology

Center, personal communication, May 1998.

4t Cikanek, H.A., "Lewis Research Center Revised

Submissions to the Advanced Technology Lead Center

for the Fiscal Year 2000 Budget Call", NASA official

letter, June 1998.

42Gerwin, R. A., et al., "Characterization of Plasma

Flow Through Magnetic Nozzles", Astronautics Lab

(AFSC) AL-TR-89-092, final report, February 1990.

43Kamhawi, H., Ohio State University, unpublished

notes, April 1997.

44Turchi, P.J., Ohio State University, unpublished

notes, April 1997.

45Turchi, P.J., Ohio State University, email, June 1998.

46Borowski, S.K. and Dudzinski, L.A.," '2001: A

Space Odyssey' Revisited - The Feasibility of 24 Hour

Commuter Flights to the Moon Using NTR Propulsion

with LUNOX Afterburners", AIAA paper 97-2956,

July 1997.

47Sutton, G.P. and Ross D.M., Rocket Propulsion

Elements, John Wiley & Sons Pub., 1976.

48Keller, C. and Schmidt, D., "Industrial Closed Cycle

Gas Turbines for Conventional and Nuclear Fuel", Gas

Turbine Products and Turbine Products Show,

Houston, Texas, March 5-9, 1967.

49Bammert, K., '"'A General Review of Closed Cycle

Gas Turbines Using Fossile, Nuclear, and Solar

Energy", Verlag K. Thiemig, 1975.

50Bohm, E., "Gas Turbines for Nuclear Power Plants",

off-print for Euratom Review, February 1968.

51Juhasz, A.J., et al., "Closed Brayton Cycle Power

System with a High Temperature Pellet Bed Reactor

Heat Source for Nuclear-Electric Propulsion

Applications", NASA TM- 105933, October 1992.

31
s2Kuriyama, M.,etal., "High Energy Negative-Ion

Based Neutral Beam Injection System for JT-60U",

Fusion Engineering and Design, voi. 26, 1995.

5_Grisham, L., Princeton Plasma Physics Lab, email,

May 1998.

54Ohara, Y., "Progress of Neutral Beam R&D for

Plasma Heating and Current Drive at JAERI ", Fusion

Engineering and Design, vol. 26, 1995.

s5 "Nuclear Reactor/Propulsion System Interaction",

General Dynamics Space Systems Division, final

report, November ! 993.

s6"Space Shuttle System Summary", Rockwell

International Space Systems Group, May 1980.

57"Saturn V News Reference", NASA Marshall Space

Flight Center, et al., December 1968.

58Kudlac, M., NASA Lewis Research Center, personal

communication, June 1998.

59"Nuclear Reactor/Propulsion System Interaction",

General Dynamics Space Systems Division, final

Report, November 1993.

60"Advanced Communication Technology Satellite",

NASA Fact Sheet, ACTS Pub #301.

6_Cook, D. and Bauer, B., "Advanced Communications

Technology Satellite Program Overview", April 1993.

62Kunath, R., NASA Lewis Research Center, personal

communication, June 1998.

6s "Space Shuttle System Summary", Rockwell

International, May 1980.

Borowski, S.K., Peng, Y-K. M., and Kammash, T.,

"Radio-Frequency-Assisted Current Startup in the

Fusion Engineering Device," Fusion Technology, Voi.

6, July 1984.

65Borowski, S. K., "Inductive Current Startup in Large

Tokamaks With Expanding Minor Radius and RF

Assist," Oak Ridge National Laboratory Report FEDC83 / 8, February 1984.

66Cataldo, R.L., "Test Results of a Ten Cell Bipolar

Nickel-Hydrogen Battery", NASA TM-83384, August

1983.

67Cataldo, R.L., NASA Lewis Research Center,

personal cemmunication, June 1998.

6g "Fastcharged Traction and Starter Supercapacitor

Batteries", marketing brochure, ESMA Co., Moscow,

Russia, 1998.

69Turchi, P.J., Ohio State University, personal

communication, June 1998.

7oSankovic, J.M., et al., "Hydrogen Arcjet

Technology", NASA TM-105340, October 1991.

7_Hawkins, K., "Space Vehicle and Associated

Subsystem Weight Growth", SAWE paper 1816, May

1988.

75Anonymous presentation, "MSFC Programs Weight

Growth History", April 1979.

32

Form Approved REPORT DOCUMENTATION PAGE OMB No. 0704-0188

Public reporting burden lot this collection of information is estimated to average 1 hour per response, includ Ig the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing Ihe collection of information Serd comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for information Operations and Reports. 1215 Jefferson

Davis Highway, Suite 1204. Arhngton, VA 22202-4302, and to the Office of Management and Budget, Papecwork Reduction Proiect (0704-0188). Washinglon, DC 20503,

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 1998

4. TITLE AND SUBTITLE

A Spherical Torus Nuclear Fusion Reactor Space Propulsion Vehicle Concept

for Fast Interplanetary Travel

6. AUTHOR(S)

Craig H. Williams, Stanley K. Borowski Leonard A. Dudzinski,

and Albert J. Juhasz

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

Technical Memorandum

5. FUNDING NUMBERS

WU-242-74-74-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-11442

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--1998-208831

AIAA-98-3591

11. SUPPLEMENTARY NOTES

Prepared for the 34th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE. and ASEE,

Cleveland, Ohio. July 13-15, 1998. Craig H. Williams, Stanley K. Botowski, Leonard A. Dudzinski, and Albert J. Juhasz,

NASA Lewis Research Center. Responsible person, Craig H. Williams, organization code 65 I(), (216) 433-7063.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Categories: 20, 15, and 16 Distribution: Ncnstandard

This publication is available from the NASA Center for AeroSpace Information. COl ) 621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

A conceptual vehicle design enabling last outer solar system travel wz, s produced predicated on a small aspect ratio

spherical torus nuclear fusion reactor. Initial requirements were for a human mission to Saturn with a > 5% payload mass

fraction and a one way trip time of less than one year. Analysis revealed that the vehicle could deliver a 108 mt crew

habitat payload to Saturn rendezvous in 235 days, with an initial mass in low Earth orbit of 2.941 rot. Engineering

conceptual design, analysis, and assessment was performed on all major systems including payload, central truss, nuclear

reactor (including divertor and fuel injector), power conversion (inclu, ling turbine, compressor, alternator, radiator,

recuperator, and conditioning), magnetic nozzle, neutral beam injector, tankage, stardre-start reactor and battery, refrigeration, communications, reaction control and in-space operations. D ;tailed assessment was done on reactor operations,

including plasma characteristics, power balance, power utilization, and component design.

14. SUBJECT TERMS

Nuclear: Fusion: Interplanetary; High thrust: Nuclear propulsion: Saturn; Jupiter;

Spacecraft

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

38

16. PRICE CODE

A03

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std Z39-18

298-102