Fusion Propulsion has an enormous potential for space exploration in the near future. In the twenty-first century, a usable and efficient fusion rocket will be developed and in use. Because of the great distance between other planets and Earth, efficient use of time, fuel, and payload is essential. A nuclear spaceship would provide greater fuel efficiency, less travel time, and a larger payload. Extended missions would give more time for research, experiments, and data acquisition. With the extended mission time, a need for an artificial environment exists. The topics of magnetic fusion propulsion, living modules, artificial gravity, mass distribution, space connection, and orbital transfer to Mars are discussed. The propulsion system is a magnetic fusion reactor based on a tandem mirror design. This allows a faster, shorter trip time and a large thrust to weight ratio. The fuel proposed is a mixture of deuterium and helium-3. Helium-3 can be obtained from lunar mining. There will be minimal external radiation from the reactor resulting in a safe, efficient propulsion system

Buchholtz, Brent

Freuh, Jim

Hedrick, James

Jensen, Eric

Ward, Paul

English

NASA Technical Reports Server

MARS MANNED FUSION SPACESHIP
James Hedrick, Brent Buchholtz, Paul Ward, Jim Fr/_h, and
Eric Jensen
University of Wisconsin-Madison
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Fusion Propulsion has an enormous potential for space exploration in the
near future. In the twenty-first century a usable and efficient fusion
rocket will be developed and in use. Because of the great distance
between other planets and Earth, efficient use of time, fuel, and payload
is essential. A nuclear spaceship would provide greater fuel efficiency,
less travel time, and a larger payload. Extended missions would give
more time for i'esearch, experiments, and data acquisition. With the
extended mission time, a need for anificial environment exists. We
address the topics of magnetic fusion propulsion, living modules,
artificial gravity, mass distribution, space connection, and orbital
transfer to Mars.
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The propulsion system is a magnetic fusion reactor based on a tandem
mirror design. This allows a faster, shorter trip time and a large thrust-
to-weight ratio. The fuel proposed is a mixture of deuterium and
helium-3. Helium-3 can be obtained from lunar mining. There will be
minimal external radiation from the reactor resulting in a safe, efficient
propulsion system.
INTRODUCTION
The use of fusion power will be a major development in space travel.
Last year, a design project was started on a manned Mars mission
using fusion propulsion. This year's design team has continued the
analysis on the spacecraft with a general emphasis on the living
quarters design.
Due to the extended length of a Mars mission, there is a need for
artificial gravity. A state of zero gravity tends to weaken the body
by bone decalcification, diminishing bone marrow, reduced tone and
atrophy of the heart and lungs. We have chosen to create a 0.8g
environment in order to reduce design implications and yet have a
healthy environment.
The initial design, Figure 1, incorporates a ring of modules about the
magnetic fusion reactor supported by four radial shafts. To produce
artificial gravity, the modules will be spun about the axis of the ship
to produce constant gravity around the ring. Spinning the modules
too fast will cause motion sickness of the astronauts. A study done
by the Navy 1 reports that for spin rates less than 5 rpm, people will
lose all effects from motion sickness within 24 hours. A radial
velocity of 4.2 rpm will produce the 0.8g environment.
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BELT RADIATOR
ONE OF 25 SETS SHOWN
MASS ELEVATOR
MODULES FOR:
A_locks
Barrooms
Bedrooms
Cat'eteria
Controls
I._r_y
U/e Support
R,_-rlat;on
Recydin g
Research
Storeroom
St_l[e"
Fig. 1. D-3He Fusion Propulsion Reactor Design with Living Quarters
Due to cosmic radiation, the living modules needed to be constructed
out of a material to shield the crew while also having a high
strength-to-weight ratio. For this purpose, the modules are made out
of an aluminum-lithium alloy.
Each individual module was chosen to be an enclosed capsule
consisting of 3 joined cylindrical pieces. They are 4 m in diameter
and approximately 7.3 m in length. When coupled together, the 36
modules form a ring with a 80 m diameter. The modules were not
chosen to be curved since disorientation occurs when one walks on
what feels to be a fiat surface, but appears curved.
In this paper, we present most of the design work our team did along
with a brief description of the power system.
POWER SYSTEM
The power source chosen as the basis of the present design is a linear
magnetic fusion reactor. The fuel cycle, deuterium and helium-3
(D-3He), produces over 95% of its power as charged particles, which
are confined by a magnetic field and guided to provide direct thrust.
Most of the present fusion research effort focuses on a different fuel,
deuterium-tritium (D-T), which gives 80% of its power as neutrons.
The copious neutrons from D-T fuel lead to more massive radiation
shielding requirements and to the use of a thermal cycle for energy
conversion, with consequent low efficiency, large radiator mass, and
a separate system for converting electricity to thrust. The choice of a
linear design, the tandem mirror, provides the benefit of allowing
direct thrust by guiding the fusion plasma along magnetic field lines
and out the end of the magnetic 'bottle.' Linear configurations,
however, are less developed than their toroidal counterparts, such as
the tokamak, which more easily confine a plasma, but are more
difficult to configure for direct thrust.
The specific configuration used is based on the D-3He tandem mirror
reactors designed for use either on Earth 2 or in space3. The basic
principal of the designs is that plasma losses out the end of a linear
magnetic field configuration are greatly reduced by a combination of
increasing the magnetic field (creating a 'magnetic mirror') and by
generating an electrostatic potential to enhance ion confinement. A
key critical issue is verifying tandem mirror physics for reactor
conditions. The main parameters for the design are given in Tablel 3.
Note that 77% of the fusion power is converted to thrust, and that the
specific power of the fusion propulsion system is 1.2 kW/kg,
including radiators.
Table1. D-3He Tandem Mirror Fusion Propulsion System
Design Parameters.
Specific power
Fusion power
Thrust power
Thrust efficiency
Input power
Thermal power
1.2 kWthrust/k g
1960 MW
1500 MW
77%
115 MW
574 MW
(Bremsstrahlung and synchrotron radiation,
neutrons, plasma not usable for thrust)
Total mass
Total length
Midplane outer radius
Main magnetic field
3He to D density ratio
Neutron wall loading
1250 Mg
113 m
1.0 m
6.4 T
1
0.17 MW/m2
LIVING MODULES
Stress Analysis
The thickness of the module walls was earlier determined to be 4 cm
thick9 to shield the living quarter from cosmic radiation present in
the solar system. The stress calculations of the module shell have
resulted in the fact that stress is not a design constriction for the
module shell. This thickness is almost twice as that needed to safely
contain the environment of the modules.
The modules were modeled as a beam rigidly fixed at both ends in
cylindrical coordinates. A continuous load developed by the module
and cargo masses per radian multiplied by 0.8g is applied to the
beam. From standard calculations, the maximum shear and normal
stresses can be found. By using a Mohr's circle analysis the
maximum shear stress was found to be on the order of 22 MPa. This
value is well below the maximum allowable shear stress for the
aluminum-lithium alloy.
Spaeedoor
A door is needed between modules to isolate the module in case of
emergency such as wall rupture, toxic gases, or a fire. There will be
a door present between each of the 36 modules. The door must
operate in low temperatures, hold 14.6 psi pressure, and close
automatically at high speed. A sandwich panel is used for the door
because of its light weight and high strength. An aluminum
honeycomb 13.5 cm thick is sandwiched between two titanium
sheets, each 0.25 cm thick. The door is 187 cm high and 93.1 cm
wide having a total mass of 10.7 kg. The bottom and top edges are
rounded in shape to prevent the corners from curling and to press
equal force around the edge on the seal.
The door slides on tracks and can be closed by a spring or electric
motor. The door can also be opened and closed manually. Sliding in
tracks allow the door to be close faster in response to emergency. A
control system would be needed to release and close the door. A
pressure difference between modules will push the door against an
O-ring seal resulting in an effectively sealed doorway. The door can
also be sealed manually if there is no pressure difference.
Pressure Loss
An important calculation for the module door design is the time for
the pressure to fall from one atmosphere to a minimum value of 55
kPa for human habitation 4. If this time is known, then we can
design the door drive system to close the door in that time. As a
'worst case' assumption, we chose the. puncture size to be greater
than door size, so we can use the door size which gives a nozzle area
of 1.67 m2.
The calculation begins by determining the critical pressure ratio for
sonic flow through the opening. The ratio of space pressure to
internal pressure will always be less than critical since space is at 0
kPa. With sonic flow, the weight flowrate equation5 gives a flowrate
of 3000 N/sec.
To obtain a loss time, we used initial and final conditions to find the
specific volume of the air and from there find the mass of air which
left the spacecraft. For one module, the fastest time of pressure loss
is 0.03 sec and for the entire living quarters, the time is 1.14 see.
Thermal Radiation Loss
The question was raised as to whether the living quarters could be
used to dissipate some of the heat generated by the reactor. To
determine the emissive power of the living quarters, we took the
maximum emissivity e of aluminum (0.84) and found what kinds of
power values we can expect to emit at different locations. We chose
three cases for exchange: the spacecraft at Earth orbiting at
Lagrange Point 1, the ship orbiting at Mars, and the ship radiating to
empty space.
From Incropera
equation:
and DeWitt 6, we can use the radiation exchange
qm =(Ebm-Jm) x(exAm)/(1-e)
where: Ebm = blackbody radiation from a module
Am = surface area of a module
The radiosity of the module is:
Jm = (e x Ebm)+ (1 - e) x Gm
The irradiation to the module is:
Gm = (Ap x Ebp x Fpm + As x Ebs. x Fsm) / Am
where: A p = surface area of planet
Ebp = blackbody radiation from planet
Fpm = view factor of module by planet
As --- surface area of the sun
Ebs = blackbody radiation from the sun
Fsm = view factor of module by the sun
The view factors are calculated after reading the view factors of the
planet or sun by the ship off of a graph from Chapman7.
We have determined the radiation from the living quarters at Earth
to be 1.15 MW8. At Mars the radiation is 1.21 MW and in empty
space the heat loss is 1.22 MW. This is not enough heat loss to be
used as an effective radiator and insulation of the modules is not a
major design consideration.
RADIATION SHIELDING
Shadow shielding
Since there will be some external radiation present from the reactor,
the living modules must be shielded to protect the astronauts. A
dose of 2.5 millirem/hr 10 is an amount of exposure that can be
exposed to humans. Shadow shielding was the first idea since it
seemed logical that a shadow shield close to the module would
provide the required protection and minimize the mass of the shield.
This method of shielding takes advantage of the l/r2 fall off over the
distance from the reactor to the modules. Boric acid (H3B03) was
chosen as the shielding material because of its shielding effectiveness
and low density. The boric acid shield includes a steel tank in
calculations (5% of the weight, about 4-5 cm thick). A shield
thickness of 1.75 meters and a total mass of 3.89 x 106 kg is needed
to shield the modules in this way.
Reactor Cover Shielding
The shielding mass for the shadow shielding seems to be a quite
large number. Because of the large mass of the shield, a second
shield calculation was done for a shield around the reactor. The
shield thickness is 1.37 meters and mass of this type of shielding is
1.97 x 106 kg about half of the shadow shield requirement.
Diagrams of these two shielding methods are shown in figure 2.
Although it would seem that the reactor cover shield would seem a
better alternative, it was initially neglected because of the heat
produced by the reactor. A complex method of cooling the magnets
would be needed, if even possible, to cool the magnets so they do not
melt. The shield may need to be a few centimeters thicker because
to a cooling system. However, if possible, this would still be the
lightest shield.
Shadow
Fusion
reactor
Reactor cover
sh eld
93m
•100m
37m
Figure 2. Diagram of reactor radiation
MASS ELEVATORS
One problem with rotating the living quarters is that the stability of
the spacecraft is offset with changes in the distribution of mass
around them. To account for these small changes, we incorporated
masses in the four radial support shafts which move with crew
movement so that the mass center remains at the rotational center of
the ship.
In order to determine how large the moving masses have to be, we
must find the worst mass distribution problem. As an estimate, we
assumed that 52.5% of the mass is evenly distributed on one side of
the living modules while the remaining 47.5% is distributed on the
other 8. The corresponding mass elevator is moved out to its furthest
position to keep the mass center at the longitudinal axis of the
spacecraft. For a total living quarter mass of 13000 Mg, each mass
elevator is calculated to be 481 Mg. We decided that the elevators
must be as dense as possible for them to fit into such a small volume,
so we decided to make them out of cast iron. The mass would be
cylindrical with a radius of 1.8 m and length of 6.15 m.
w° Since the moving mass introduces a Coriolis acceleration, we must be
careful not to move them too fast so that the force exerted on the
support shafts is tolerable. A maximum velocity of 0.1 m/s was
chosen, however, no stress analysis was done on the shafts.
ELEVATOR SHAFTS
The living quarters is held in place by four radial spokes or shafts.
These shafts are also used for housing the 'mass elevators' or
movable masses which correct for uneven mass distribution.
These shafts were designed to be 38 m in length and and outside
diameter of 4 m with a 13.2 cm thickness 8. The design needs to be
analyzed with the Coriolis force mentioned earlier and does not
represent a final decision.
We did choose to make the shafts out of AI 2014 for its ability to
remain tough in cold temperatures and for its high strength-to-
weight ratio.
MASS ESTIMATE
As in any space mission, mass reduction is a critical part of design.
We have started with an allowable mass of 13.000 x 106 kg for the
life support system. The life support masses were determined by
adding the masses of several significant components: Mass elevators,
module shell, and radiation shielding. This was subtracted from the
total allowable mass. The remaining amount is considered to be the
payload. The radiation shielding is by far the heaviest component of
the five. The masses of each of the components are given in table 2.
Note that the pay load is more than half _ of the mass of the life
support system. The reactor and radiator masses will be on the
order of about 2 x 106 kg. Payload will then be about 45% of the
total mass, which is very good for a trip of this length.
Radiation shielding
Elevator masses
Module 'shell'
Elevator shaft
Payload
3.890 x 106 (kg)
1.920 x 106
3.202 x 105
1.780 x 105
6.692 x 106
Total mass 13.000 x 106 (kg)
Table 2. Component masses
°SPACE COUPLING
The modules and spaceship will be assembled is space before
operation. Initially the modules will be brought together and held
magnetically. Magnets are ideal for space, since they require no
mechanical force to implement. Magnets can also be used to orient
the modules correctly before permanent coupling is made.
The modules will be more permanently secured with some type of
mechanical device which can be operated from within the modules
themselves. This will save time and money from having to secure
the living quarters with spacewalking operations.
LOW THRUST TRANSFER
Much emphasis has been placed on the use of low thrust propulsion
systems, such as fusion power, to propel spacecraft to other planets.
Missions beyond the moon need faster trip times and with larger
payload fractions than chemical rockets 3. The key to satisfying these
needs is high exhaust velocity, since payload fractions depend
exponentially on - (vf- vi)/vex. A low thrust fusion propulsion
system also has many other advantages over a chemical system:
• High specific power values
• Direct conversion of energy to thrust
• Thrust-mode flexibility over a wide range of thrust-
to-weight ratios and specific impulses
FLIGHT"rIME FOR SAME PAYLOAD
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E=r_-U=r= Ezru't.Jul=t_r
PAYLOAD FOR SAME FL1GHTTIME
'L0'
U
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I; ell
==
'll
L4
u
an
Sc_ _c_ tmo_
Fig. 3. Flight Time and Payload Comparisons
As with any transfer, we would like to optimize payload and transfer
time. For a trip to Mars, our low-thrust rocket will use about 70%
less fuel than a high thrust chemical rocket with comparable transfer
times.
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Ilat,._iM A
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= 10'¢3,.
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,_ 10'
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10"i 10" 10.= 10" 10"
Thrust to Weight Ratio
Fig. 4. Thrust to Weight Ratio vs. Specific Impulse
The analysis of transfer is broken up into three stages: the escape,
the transfer, and the capture. In the transfer, one can obtain two
differential equations from the equations of motion. To solve, one
needs to use computational techniques using 6 boundary conditions
and selecting acceleration components which produce these
restraints 11. Another analysis technique which we could use is to
manipulate energy equations and differentiate to maximize payload
or minimize transfer timel2.
Several calculations for the trip have been done by Stuhlinger 12 for
motion in a central gravity field. For one particular trip, he showed
that, using variable thrust, and Earth to Mars trip would take 178
days at an acceleration of 1.6 x 10 -4 Earth gravities with a resulting
payload fraction of 0.40.
CONCLUSIONS
We have covered a wide range of topics in our work. This is one
indication of the enormous size of such a project and how much work
is left to do. We have laid the basic foundations for future design
and have not run across any major design problems.
Our work has covered the areas of module dimensions and stresses,
module door analysis, initial coupling processes, mass distribution
correction procedures, radiation shielding requirements, and the low
thrust analysis used for travel to other planets.
Future work may include design of the drive system for the doors,
the drive system for the mass elevators, finding coupling power
requirements, and look at possible windows for launch to Mars or
Jupiter.
vp
ACKNOWLEDGEMENTS
Special thanks to Dr. J.F. Santarius and Ron Thomson for their help
and support throughout this semester. Funding provided by NASA
through NASA Lewis Research Center.
References
1. Space Resources and Space Settlements, NASA SP-428, U.S.
Government Printing Office, 1979.
2. Santarius, J.F., "Ra: A High Efficiency, D-3He, Tandem Mirror Fusion
Reactor," 12th Symposium an Fusion Engineering, (IEEE,NY,1987)
3. Santarius, J.F., "Lunar3 He, Fusion Propulsion and Space
Development", Fusion Technology Institute, 1988.
4. Dreyfuss, Designing for People, 1966
5. Vennard, Elementary Fluid Mechanics - 4 th Edition, John Wiley
and Sons, New York, 1962.
6. Incropera, F.P. and DeWitt, D.P., Introduction to Heat Transfer, John
Wiley and Sons, New York, 1985.
7. Chapman, A.J., Heat Transfer , Third Edition, Macmillian Publishing
Co. Inc., New York, 1974.
8. University of Wisconsin-Madison EM 569, 1990 Fusion Rocket
Team
9. University of Wisconsin-Madison EM 569, 1989 Fusion Rocket
Team
10. Khater, H., Radiation Dose Tables, provided by Dr. J.F. Santarius.
11. Irving, J.H. and Blum, E.K., "Comparative Performances of Ballistic
and Low-Thrust Vehicles for Flight to Mars", Vistas in Astronautic5,
Pergamon, New York, 1959.
12. Stuhlinger, E. Ion Propulsion for Space Flight, (McGraw-Hill, NY
1964)


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