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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) 



^_ 6.00E+0S 




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+ 4 + + + ++++ + + + + + + + + + + +++ + + + + + + + + + + + + + + + + 
UJLUIULULU UJLUIULU UJLUIULULULULIJIIILULU LUUJLU UJlUmtiJLiJLUlUUJUJLLJLUlUUJ UJUiLU 



Tima: seconds 
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.