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IECEC-98-012
33rd Intersociety Engineering Conference on Energy Conversion
Colorado Springs, CO, August 2-6, 1998
ELECTWCAL POWER SYSTEMS FOR NASA's SPACE TRANSPORTATION PROGRAM
Louis F. Lollar
NASA/Marshall Space Flight Center
Preliminary Design Office/PDl 1
Huntsville, AL 35812
(256)544-3306
(256)544-4225
Louis.Lollar@msfc.nasa.gov
ABSTRACT
Marshall Space Flight Center
(MSFC) is the National Aeronautics and
Space Administration's (NASA) lead
center for space transportation systems
development. These systems include earth
to orbit launch vehicles, as well as vehicles
for orbital transfer and deep space
missions. The tasks for these systems
include research, technology maturation,
design, development, and integration of
space transportation and propulsion
systems.
One of the key elements in any
transportation system is the electrical
power system (EPS). Every transportation
system has to have some form of
electrical power and the EPS for each of
these systems tends to be as varied and
unique as the missions they are supporting.
The Preliminary Design Office
(PD) at MSFC is tasked to perform
feasibility analyses and preliminary design
studies for new projects, particularly in the
space transportation systems area. All
major subsystems, including electrical
power, are included in each of these
studies. Three example systems being
evaluated in PD at this time are the Liquid
Fly Back Booster (LFBB) system, the
Human Mission to Mars (HMM) study,
and a tether based flight experiment called
the Propulsive Small Expendable Deployer
System (Pro SEDS). These three systems
are in various stages of definition in the
study phase.
Louis C. Maus
NASA/Marshall Space Flight Center
Preliminary Design Office/PDl 1
Huntsville, AL 35812
(256)544-0484
(256)544-4225
lou.maus@msfc.nasa.gov
The goal of this paper is to
describe the goals, missions, and system
requirements of each project and then to
focus on the unique EPS requirements that
flow down for each of the three projects.
Finally, we will discuss potential new EPS
technologies that could be used to better
meet the project requirements.
L BACKGROUND
Marshall Space Flight Center
(MSFC) is the National Aeronautics and
Space Administration's (NASA) lead
center for space transportation systems
development. These systems include earth
to orbit launch vehicles, as well as vehicles
for orbital transfer and deep space
missions. The tasks for these systems
include research, technology maturation,
design, development, and integration of
space transportation and propulsion
systems.
The Preliminary Design Office
(PD) at MSFC is tasked to perform
feasibility analyses and preliminary design
studies for new projects, particularly in the
space transportation systems area. All
major subsystems, including electrical
power, are included in each of these
studies. The final study reports for these
projects are costed, scheduled, and then
presented to the customer for possible
follow-on funding. A few projects that
have passed through PD have been the
Hubble Space Telescope, the Advanced X-
ray Astrophysics Facility, and the
redesigned International Space Station
(ISS).
II. INTRODUCTION
The Preliminary Design Office at
MSFC is investigating over twenty-five
projects at this time. One can group these
projects into various themes. For example,
one theme could be "airplanes in space"
which would include the liquid fly back
booster (LFBB) and the single stage to
orbit rocket based combined cycle (RBCC)
projects. Another theme could be
"transportation to the planets" which
would include the Mars sample return
mission and the human mission to Mars
(HMM) studies. Another theme might be
"tethers for transportation" which would
include the many boost/deboost tether
projects for the ISS and Mir, plus the
propulsive small expendable deployer
system (ProSEDS) flight experiment.
One of the key elements in all of
these projects is the electrical power
system. Each one of these projects has to
have some form of electrical power and
the EPS for each of these systems tends to
be as varied and unique as the missions
they are supporting. To view as many EPS
requirements as possible, three example
projects currently being evaluated in PD
will be reported. These are the LFBB, the
HMM, and the ProSEDS. The goal of this
paper is to describe the goals, missions,
and system requirements of each project
and then to focus on the unique EPS
requirements that flow down from each
project.
III. THE LIQUID FLY BACK BOOSTER
A. Overview
The concept of retrieving the first
stage booster of a multi-stage rocket
system has been around since early in the
Space Shuttle program. Requirements and
funding levels changed which led to the
recoverable solid rocket booster concept
presently being used in the Space
Transportation System (STS).
Since then, partially as a result of
the Challenger accident, various
replacement options for the solid rocket
boosters have been studied. Some of these
studies have resulted in the redesigned solid
rocket motor (RSRM) system now flying,
yet none of the more ambitious redesigns
have been implemented. A 1993 study
concluded that from a life cycle cost and
safety improvement perspective, the most
competitive booster design would use liquid
rocket engines and be fully recoverable by
flying back to the launch site.
In 1994, a NASA study team
representing Johnson Space Center (JSC),
Kennedy Space Center (KSC), and MSFC
performed a pre-phase A study assessment
on a LFBB.^ In 1996, MSFC, along with
the Boeing Company and Lockheed-
Martin, began a feasibility and cost study
into using the LFBB as a phase IV upgrade
to the STS. One of the primary goals of
this study is to develop a set of level I
design requirements. The EPS design
concepts are derived from this study's
design requirements.
B. Mission Groundrules and
Requirements
In order to bound the design space,
several groundrules were established for the
LFBB. The first groundrule was that the
boosters would be designed for use in the
STS, therefore Orbiter, External Tank
(ET), and launch and processing facilities
modifications must be minimized. Next,
launch loads, maximum Q points, mission
requirements, and environmental impacts
would be better or no worse than the
present STS levels. Finally, all LFBB
designs will focus on lowering the
operations and processing costs as much as
possible.
Based on these groundrules and
other derived requirements, the following
basic concept has emerged. At T -
TBD(few) seconds, the three Space Shuttle
Main Engines (SSME) explode into life.
Shortly thereafter, the eight to ten Hquid
booster engines fire. Now, unlike the
RSRM's, if an engine problem occurs, the
system could be shut down or launched
into a known, safe trajectory. If all
systems are operating nominally, the STS
operates almost the same as it does now
which includes ascent guidance and control
being governed by the orbiter. At
approximately T+2.5 minutes, 150,000
feet, 31 nautical miles(N]V[) downrange,
and 5500 feet per second, the LFBB
separates from the orbiter and ET. The
boosters coast to an apogee of
approximately 260,000 feet and 100 NM
downrange while all deployables (wings,
canards, etc.) deploy. The boosters coast
at a 40 degrees angle of attack and
perform a large bank turn towards the
launch site. At about T + 8 minutes,
31,000 feet, and 215 NM downrange, the
air breathing engines(ABEs) perform a
cold start and the boosters autonomously
fly back to the KSC landing strip where
they complete a safe autolanding at
approximately T + 52 minutes. The
boosters are then rolled back to a
processing area to be prepared for
integration into the STS for the next
flight.
C. EPS System Requirements
The basic requirements for the
EPS combine the redundancy and
reliability requirements of a spacecraft
EPS with the maintainability requirements
of an aircraft EPS. During ascent, the
LFBB EPS will have to provide power to
the LFBB engine controllers and avionics
as well as interface with the orbiter' s EPS.
After separation, LFBB control functions
reverts to its own avionics suite with the
energy being supplied by its own space
qualified power source. After reentry, the
bank turn, and the ABEs cold start, the
EPS can now obtain energy from a power
take-off on the ABEs. After landing and
the ABEs shut down, the EPS will provide
any power needed for vehicle health
monitoring (VHM) until the ground
support equipment arrives.
Orbiter interface
(if applicable)
(AgZn Batteries)
£
120 or 28 Vdc bus
dc to dc
converter
(270 to 120
or 28 Vdc)
Energy
Storage
(270 Vdc)
Power
Distributor
Subsystem
Loads (- 1 kW)
Power
Controller
turbo-
alternator
■— I unit
270 Vdc bus
H
Power
Distributor
EMA Loads
(~ 49 kW)
Figure 1. Preliminary Electrical
[One Boo
A conceptual EPS design is shown
in figure 1. The system is two fault
tolerant with a 270 Vdc bus for the
electromechanically based actuator
systems loads and a 120 Vdc or 28 Vdc bus
for the remaining avionics loads. The
power sources include three 270 Vdc, 60
Ahr silver-zinc (AgZn) batteries and a
turbo-alternator unit for each ABE. The
power controller monitors and controls
the flow of energy to the busses and the
Power System For LFBB
ster]
power distributors provide protection and
monitoring to each load. This proposed
system should be versatile enough to
handle the varied load and source power
profiles of the LFBB, yet simple enough
to be reliable and serviceable.
IV. ProSEDS
A. Introduction
The tether project, ProSEDS, is
a continuing effort in the research and
development of an electrodynamic tether
system that has operational applications
for future spacecraft scenarios and
missions. Electrodynamic tethered flights
of the recent past, such as the Tethered
Satellite System (TSS-1 & -IR) and the
Plasma Motor Generator (PMG)
Experiment, has provided experimental
data to further develop practical systems.
With conductive tethers, induced
voltages were measured in both TSS-1 & -
IR flights. In the TSS-IR flight, induced
vohages in excess of 3 kV were measured
in a 20 km length tethered system. The
PMG experiment verified that tether
current will flow in both directions which
represented both the generator and motor
modes of operation.
For future space based system,
recent engineering studies have shown that
practical implementation of conductive
tethers can be achieved in the following
applications: (1) electrical power
generation, (2) orbital reboost of space
based systems if excessive electrical power
is available, and (3) deorbit pay loads or
space debris utilizing electrodynamic drag
forces.
The ProSEDS mission will
demonstrate continued development in the
following areas: (1) deboost or deorbit a
payload through the utilization of
electrodynamic drag forces, (2) collect,
test and validate the current collecting
capability of a "bare wire" tether which
will enhance the current collecting
capabilities, and (3) convert the electrical
energy generated in the tether system into
a more useable form.
B. The Space Plasma Environment
The basic or fundamental
principle of the electrodynamic tether
system is that a conductive wire tether
cutting the earth's magnetic field will act
like a generator and induce an emf into the
conductor that will cause a current to flow
in a closed loop. The magnitude of this
induced emf (v x B • L) is a function of
velocity (v), magnetic field strength (B),
and tether length (L). These parameters
are functions of the orbit definition (such
as altitude and inclination), solar flux
activity and time of year.
As defined by the ProSEDS
mission, the payload will be a secondary
payload on a Delta II upper stage. The
orbit will be 400 km circular with an
inclination of 32 degrees. Figure 2 depicts
the predicted open circuit voltage induced
into a 5 km conducting tether. From the
graph, it can be seen that an average
induced voltage, of approximately 140
volts per kilometer can be expected. Peak
voltages, however, can be as high as 200
volts per kilometer.
EMF® (400 km, IncI 32 Dsg)
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Figure 2: Induced Tether Voltage (Open Circuit)
Non-ESS
Bus
Figure 3. ProSEDS Electrical
One of the more challenging
aspects of the pre-mission phase is to
predict the magnitude of the current flow
through the "bare wire" tether system and
through the plasma of the ionosphere.
Within the constraints of the mission and
the available hardware, peak orbital
average currents of two amperes are
anticipated. Peak current may be as high
as 4.5 amperes.
C. The Electrical Power Subsystem
From a conceptual standpoint,
one of the aspects of the ProSEDS
mission, and subsequently the electrical
power subsystem, is to collect data over at
least three orbits. This will be
accomplished with a primary battery sized
for the deployment of the 20 km tether
system plus the energy required for the
loads during this minimum period of three
orbits of data collection. Following this,
the power source for the loads will be the
tether system which will supply energy for
secondary battery recharge and power
conversion.
A secondary mission objective is
to demonstrate the storage, conversion
and regulation of the tethered-generated
electrical power. The basic concept, shown
in Figure 3, is to provide regulation and
conversion of the very high induced
tethered voltages for further conversion
into a usable system voltage. Lacking a
Power System
high input voltage "buck" power
converter, the input voltage to the
converter is regulated using a low cost,
secondary nickel-cadmium battery. This
battery clamps a portion of the high input
system voltage with the remaining voltage
being dropped across the tether resistance,
the impedance of the ionosphere, and the
plasma contactor (not shown).
A group of front-end high
voltage vacuum relays (SW1,SW2, SW3)
are designed to operate in a sequence to
measure parameters of the tether and of
the ionosphere such as open circuit
voltage, short circuit current, and
characteristics at a specified resistive load.
The recharge logic for the secondary
battery will be controlled by relays SW3
and SW4.
V. HUMAN MISSION TO MARS
A. Overview
Working with the science and
exploration community, NASA (MSFC,
JSC, LeRC) is developing a design
reference mission (DRM) to be used in a
planning exercise to send humans to Mars
during the 2011 and 2013/2014 Mars
opportunities. The DRM represents the
most current approaches to completing
the mission and provides a baseline
architecture to analyze new technology
insertions.
B. Design Reference Mission
Beginning in 2011, two cargo
flights will leave the Earth toward Mars on
a low energy, long transit time trajectory.
The first cargo flight containing a fully
fueled Earth return vehicle will
aerocapture into a Martian orbit. This
vehicle which will eventually be used to
bring back the crew and their samples will
remain in the Martian orbit until the Mars
exploration activities are completed. The
second cargo flight will land on the
Martian surface. Its cargo contains storage
tanks, liquid hydrogen, an in-situ
propellant production (ISPP) unit, and a
nuclear surface power (NSP) unit. After
landing, the NSP unit will deploy and begin
supplying power to the ISPP unit. Thus,
the ascent vehicle's propellant will be
produced and stored on Mars before the
crew has to commit to the long journey to
Mars.
Then in 2014, a crew of six and
their exploration equipment will depart for
Mars along a more direct trajectory. After
aerocapture into a Martian orbit, the crew
and equipment will descend to the surface,
landing near the previous cargo landing
site. After 569 days on the Martian
surface, the crew and sample material will
ascend to and dock with the orbiting Earth
return vehicle for the 154 day trip back
home. [For more details, see references (3)
and (4).]
C. Transportation and EPS
Requirements of the DRM
The four major transportation
elements of the DRM are the trans-Mars
injection (TMI) stage, the descent/ascent
stage to the Martian surface, the trans-
Earth injection (TEI) stage, and the
aerobrake elements. Of these four, only
the first three will have some form of
EPS.
The TMI stage is comprised of
three nuclear thermal propulsion (NTP)
engines providing a total thrust level of
200,000 Newtons. In order to reduce the
delta-velocity budget, a TMI stage will
stay in a low Earth orbit (LEO) for up to
32 days. This requirement will force the
TMI stage to have some form of power
generation/energy storage EPS. Depending
on the number of electromechanical
actuators (EMAs) and rendezvous and
docking avionics, the EPS could be in the
lOOO's of Watts. One alternative is to use
the NTP engine as a power source in a bi-
modal concept, but this would require the
TMI stage to remain with the cargo stages
until near Mars orbit.
The TEI stage, used to return the
crew to Earth, will be similar to the TMI
stage. The primary differences involve
longer life (up to 4 years), power-hungry
cryo-coolers to minimize boil-off, and
multiple, long time between starts and
engine firings. This requires the EPS to be
highly reliable and to provide power to a
larger payload than the TMI stage. In
addition, the EPS has to operate in a
Martian orbit which is much further from
the Sun than an Earth orbit.
Finally, the descent/ascent stage
EPS could be integrated with the overall
EPS required to maintain life support for
the crew.
VI. TECHNOLOGY NEEDS AND
CONCLUSION
As we begin the 21st century, the
EPS technology requirements for our
space transportation systems will be as
diverse as our people. From unmanned
flying rockets to tethered spacecraft to
visiting our nearest neighbor planet, the
EPS design engineer will have to become
more creative as high power, high
reliability, and low mass become the
driving requirements.
To meet these requirements,
MSFC has been investigating the following
technology areas. The Department of
Defense's (DOD) "More Electric
Transportation" initiatives are producing
many new innovations in ruggedized power
generators, electrical actuation and high
power, high temperature electronics for
power management and distribution
systems. In the energy storage arena,
flywheels and Lithium-Ion batteries appear
to have tremendous potential in lowering
the EPS mass through higher Watt-hours
per kilogram levels. New lightweight, low
intensity solar cell technology, as well as
clean, safe nuclear technology could help
solve some of the challenges in sending
humans to the Martian surface and/or
more complex (read high power)
unmanned probes to Mars and beyond.
Other related technology areas such as
micro-miniaturization, composite
materials, and new thermal devices and
materials are also being investigated for
potential EPS applications.
MSFC still has more challenges
than solutions as we attempt to provide
safe, inexpensive, and reliable space
transportation systems to the United
States and the World. Thus, we will
continue to seek out new and innovative
technologies which can be used to help
solve the many challenges of space travel.
ACKNOWLEDGMENTS (ProSEDS
SECTION):
Thanks to Chris Rupp (PS04) Fred
Elliott (EL22), Kai Hwang (CSC/MSFC).
REFERENCES
1 . "Access to Space Option 1 Team
Summary Report," JSC White Paper.
August 31, 1993.
2. Peterson, W., et al, "Liquid Flyback
Booster Pre-Phase A Study
Assessment," NASA Technical
Memorandum 104801. 1994.
3. Hoffinan, Stephen J., "Human
Exploration of Mars: the reference
mission of the NASA Mars
Exploration Study Team," NASA SP-
6107. 1997.
Kos, Larry, "The Human Mars
Mission: Transportation
Assessment," 15th Symposium on
Space Nuclear Power and Propulsion
at the Space Technology and
Applications International Forum.
January 1998.