NASA Technical Reports Server (NTRS) 20040170473: Technology Innovations from NASA's Next Generation Launch Technology Program

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by  NASA Technical Reports Server (NTRS)

IAC-04-\'5.01
TECHSOLOGY INNOVATIONS FROM NASA's NEXT GENERATION LAUNCH
TECHNOLOGY PROGRAM
Stephen A. Cook
Stephen.A.Cook@,nasa.gov
Charles E. K. Moms, Jr.
[email protected]
Richard W. Tyson
Richard. W.Tvson@,nasa.g ov
"0 1, NASA Marshall Space Flight Center
Huntsville, Alabama, USA 35812
ABSTRACT
NASA's Next Generation Launch Technology Program has been on the cutting edge of technology,
improving the safety, affordability, and reliability of future space-launch-transportation systems. The ar-
ray of projects focused on propulsion, airframe, and other vehicle systems. Achievements range from
building miniature fueVoxygen sensors to hot-firings of major rocket-engine systems as well as extreme
thermo-mechanical testing of large-scale structures. Results to date have significantly advanced technol-
ogy readiness for future space-launch systems using either airbreathing or rocket propulsion.
INTRODUCTION orbit transportation concepts and architectures
options that range from man-rated, partially re-
usable systems to fully expendable launchers. In
general, NGLT addressed technologies that were
not yet technically or economically
feasible.
Fig. 1: Airbreathing-propulsion first stage.
NASA formulated the Next Generation Tech-
nology (NGLT) Program to advance the state-
of-the-art in critical and high-payoff technolo-
gies for future generations of space-launch-
Fig. 2: Expendable rocket-powered booster.
transportation systems. NGLT technologies
were to provide for lower costs, more reliability,
NASA organized the NGLT program into sets of
and greater safety for the full range of earth-to
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projects for propulsion, airframelvehicle sys- cryogenic-fuel tanks made with composite mate-
tems, operations, and flight demonstrations. rials. In short, the total list of NGLT technolo-
(Tie iarcer primariiy focused on hypersonic, air- gies was exnemeiy varied.
breathing propulsion.) Dedicated efforts for
systems analysis and integration supported the NASA’ s overall plan involving future space-
planning and evaluation processes. launch systems changed in January of this year
when President Bush announced a new vision
The NGLT Program derived its content largely for NASA space exploration (ref. 1). While the
from prior programs: the Second-Generation achievements of the NGLT program contributed
(2nd-Gen) Reusable Launch Vehicle (RLV) to the foundation for the new vision, the NGLT
Program and the Advanced Space Transporta- program itself has concluded. Several of the
tion Program (ASTP). The 2nd-GedRLV Pro- projects will continue to complete products that
gram’s technologies supported development of directly support the President’s vision.
reusable, rocket-powered, space-launch systems
anticipated to replace the U.S. Space Shuttle. In This paper provides an overview of hardware-
contrast, ASP’S vision was for vehicle concepts related accomplishments of the projects which,
with air-breathing propulsion for first-stage taken collectively, constitute the NGLT pro-
boosters; program emphasis was on the long gram. There is no overview of the entire pro-
term (e.g., for a new mode of space launch to gram’s organization or history. This paper first
become operational within several decades). As addresses technologies for: rocket-propulsion,
a combination of these two precursor programs, then air-breathing propulsion, and, finally, air-
NGLT incorporated all of NASA’s major, frame systems and vehicle operations.
longer-term investments in space-launch tech-
nologies. ROCKET- PROPULSION
NASA’s plan was to assess NGLT technology NGLT modified and then continued a series of
status periodically to determine if a specific projects already developed under the 2nd-
space-launch vehicle and associated architecture Gen/RLV program. That program had primary
could be selected for implementation in the fu- requirements derived from the needs for reus-
ture. If the technologies were ready and the ar- able systems. These requirements typically ad-
chitecture met projected needs, the selection dressed both reusability and reliability, so as to
would lead to a new program for that particular increase safety and enhance operability. While
vehicle and architecture. The new program the four projects in this set included plans for
would incorporate any on-going NGLT work ground-based testing, each project’s emphasis
that directly supported the new program’s goals. remained on technology, not development of
The new system could have been an expendable operational hardware.
or a reusable launch system - NGLT was to
enable virtually any feasible architecture. Rocket Engine Prototyue
As a result, the various NGLT projects had no- The Rocket Engine Prototype (REP) Project was
tably different objectives and content. One pro- conceived in the 2nd GedRLV program to ad-
ject sought to advance to prototype-engine vance technologies for a million-pound-thrust-
development for a large rocket engine that burns class, oxygedkerosene engine. Technologies
kerosene and liquid oxygen (LOX). Another for advanced design tools and oxygen-rich,
sought to achieve a ground-based demonstration staged combustion were anticipated to yield high
of a turbine engine capable of Mach-4 flight, reliability at reasonable cost; project plans called
while yet another focused on in-flight demon- for validation through ground-based testing of a
strations, at Mach 5 to 7, of airframe-integrated full-scale, fully integrated, engine system.
scramjets. Other projects addressed specific Boeing Rocketdyne, prime contractor for this
challenges for subsystems or components - work, designated the resulting prototype engine
such as space-based, range-safety operations or as the RS-84 (ref.2).
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temperature, oxygen-rich environment. The hot-
flow tests at NASA White Sands Test Facility
I cdciicd cxiieiiit: condiiioris (of iO,OOO yvuiids
per square inch) with gaseous oxygen. This
demonstrated significant potential for improved
reliability at reasonable cost.
NASA Glenn Research Center (GRC) completed
heated-tube testing that pumped standard-grade
Rp-1 (kerosene) through conventional copper
tubing to simulate fuel cooling of the combus-
tion chamber and nozzle. The deposition of
copper sulfides inside the tubes led to significant
Fig. 3: RS-84 Prototype Engine pressure drops -- confirming the need for opera-
tions with a low-sulphur grade of RP- 1.
Successful cold- and hot-fire tests of the first of
several configurations of one-fifth-scale pre- The REP Project also tested two types of turbine
burners produced data (and 40,000 pounds of design. NASA accomplished the first test, with
thrust) at NASA Stennis Space Center. The Rocketdyne participation, in the turbine-airflow
data set included combustion performance and test rig at NASA Marshall Space Flight Center.
stability, system durability and combustion- Results characterized blade loading and dynamic
response for an instrumented, rotating-turbine test
product uniformity.
section that was representative of the turbines in
both of the RS-84 high-pressure turbopumps.
Next, the project completed whirligig test in-
volved spinning a bladed turbine disk, modeled
after RS-84 turbopump designs. While it spun at
speeds up to 34,000 rpm, streams of high-
pressure nitrogen excited the turbine-blade dy-
namic modes. Those results demonstrated the
level of blade-motion damping required for tur-
bine reliability and, hence, long life.
Integrated Powerhead Demonstrator
A partnership between NGLT and the Air Force
Fig. 4: RS-84 Preburner Tests
Research Laboratory (AFRL) is demonstrating
new technologies for a full-flow, staged-
A Rocketdyne trade study helped define the final
combustion rocket engine using oxygen and
engine configuration that had two separate tur-
hydrogen as propellants in the Integrated Pow-
bopumps, which allowed for optimization with erhead Demonstration (IPD) project. The tech-
different turbine speeds for each pump. The
nologies include those for long-life components,
supporting analyses led to having two separate a wide range for throttling and oxygen-rich tur-
units, instead of one shaft linking two turbines, bine systems -- all integrated into a demonstrator
thus avoiding the danger of a single-shaft failure engine with 250,000 pounds of thrust.
mode (e.g. single-pump runaway).
NASA Marshall Space Flight Center manages
REP achieved notable success in demonstrating NGLT responsibilities, which include acting as
compatibility of an uncoated, high-strength alloy the government’s technical lead organization
with operations in a high-pressure, high- and testing at NASA Stennis Space Center.
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AFRL manages the contracts of the two prime Under the IPD project, the Aerojet Corporation
contractors. which are Aerojet and Boeing tested the first large-scale. flight-weight. oxy-
Rocketdyne. gen-nch preburner in the United States at its fa-
cility in Sacramento, California. The preburner
design allowed for throttling to low-power set-
tings while minimizing the temperature varia-
tions in the preburner's oxygen-rich steam -
coupled with material selection to minimize the
risk of ignition in the presence of the hot, high-
pressure oxygen gas. Successful tests covered
start sequences, various propellant-mixture ra-
tios and throttling.
Several key technologies will be tested initially
during the upcoming engine test series. Among
the component technologies, are: the first U.S.
Fig. 5: IPD engine system in facility mount
large-scale, gas-gas main injector, a channel-
wall nozzle and advanced main-combustion-
Hot-fire testing of the oxygen turbopump in
chamber concept. (The channel-wall construc-
2003 at Stennis proved the feasibility of key
tion promises to increase durability over the
technologies for improving reliability and oper-
conventional tube-type nozzle designs.)
ability for next-generation, LOX, rocket tur-
bomachinery that uses hydrostatic bearings and
Auxiliary ProDulsion Project
clutching bearings plus oxygen-compatible ma-
terials at high pressures. Later, the IPD Project
The Auxiliary Propulsion Project (APP), as ini-
performed component testing of an advanced
tiated under the 2nd-GenKV program, ad-
liquid-hydrogen turbopump at Stennis. In addi-
dressed technology gaps and risks in the
tion to validation of the hydrostatic and clutch-
development of non-toxic and cryogenic pro-
ing bearing designs, materials in the turbine
pellants for future auxiliary-propulsion applica-
proved resistant to hydrogen embrittlement as a
tions. The applications include reaction-control
way to avoid use of protective coatings. These
engines and orbital-maneuvering engines, as
turbines were designed to allow greater mass-
well as the storage, control, and transfer tech-
flow rates, resulting in lower turbine-operating
nologies associated with on-orbit storage of
temperatures for a given level of
cryogens.
engine thrust.
On-orbit propulsion systems and reaction-
control systems for launch vehicles and space-
craft provide for attitude control andor maneu-
vering at very high altitude or in space.
Traditional on-orbit propulsion systems, in-
cluding those of the NASA Shuttle, use toxic
propellants that are dangerous, expensive and
environmentally unfriendly. The APP project
has primarily focused on oxygen-based systems
but also performed some tasks on peroxide-
based systems. The project has taken a systems
approach - not only focusing on engines but
also on tanks, feed-systems and propellant-
management techniques relative to long-term
Fig. 6: IPD hydrogen-pump test operation in microgravity.
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nical knowledge database relative to the safe,
efficient use of hydrogen peroxide in rocket-
propuision appiications (ref. 3j .
Fig. 7: APP vernier-engine hot-fire test
In late 2003, APP began testing a LOdethanol,
reaction-control engine (RCE) at Gen Corp
Aerojet facilities in Sacramento, California. The
engine was designed to operate in a dual mode,
with separate sets of valves used for vernier
mode (limited to 25 pounds of thrust) and for the
primary mode (at 870 pounds of thrust). One Fig. 8: Peroxidehypergolic injector test
significant finding of the vernier testing was
confirmation of the ability to start and operate a Cryogenic-fluid management for on-orbit pro-
small thruster using liquid oxygen without the pulsion systems presents many challenges rela-
need for pre-injection gasification. Results de- tive to propellant storage and delivery. In order
termined the need for further design trades on to minimize the size and weight of the tank, it is
steady-state versus pulsing performance for the desirable to store propellants at high density and
vernier mode. low pressure. Many non-toxic, on-orbit propul-
sion concepts utilize cryogenic propellants, such
In mid 2004, testing at Gen Corp’s Aerojet as LOX, and typically require the undesirable
facilities resulted in over 3640 pulses and i847 option of active cooling and venting in space.
seconds of hot-fire operation for long duration, (Active refrigeration adds complexity and
pulses and throttled thrust representative of weight, whereas venting not only imparts un-
typical ascent, on-orbit, and re-entry operations. wanted thrust on the vehicle but also wastes
The LOx/ethanol RCE test demonstrated the propellants.)
ability to operate in a pressure-fed application
without turbo-pumps or separate heat On-orbit microgravity complicates efforts to
exchangers. deliver propellants uniformly to the tank outlet
or even to vent only vapor overboard.
In 2003, APP concluded work on component Additionally, a “gravity-enabled,” automobile-
technologies for Rocket-Grade Hydrogen Per- type, gas gauge cannot function for propellant
oxide (RGHP), which was investigated because disbursed throughout a tank in microgravity.
of its potential use as a non-toxic, room- Future test at NASA Marshall will address ap-
temperature, storable oxidizer to replace LOX or proaches to address pressure control, while tests
toxic/carcinogenic, room-temperature oxidizing at Glenn Research Center will use capillary or
compounds. Early work explored peroxide- surface-tension devices for liquid delivery and
detonation sensitivity (to pressure/temperature will test optical systems for gauging liquid
combinations), materials compatibility, JP-8- volume.
based hypergolic-fuel characterization, and the
applicability of propellant-management tech-
niques. These efforts added to the general tech-
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Propulsion Technology and Integration The PT&I Project also demonstrated friction-stir
welding between two plates of the copper-
A series of sub-projects within the Propulsion chrome-niobium aiioy, GRZop-84. Tie weids
Technology and Integration (PT&I) Project proved to have ninety percent of base-metal
advanced subsystem and component level tech- strength at both room and elevated temperatures.
nologies for rocket-propulsion systems. As compared to concepts designed with NARloy-
Z (the currently used alloy), the new materials and
Northrop Grumman Space Technology, Inc. processes could improve durability and tempera-
successfully completed testing of a reaction- ture tolerance, especially for reusable engines.
control subsystem with a 120-second firing of a
thruster using a new Platinum-Iridium (Pt-Ir) Boeing carried a small-scale cross-feed valve
combustion chamber at their Capistrano Test from design to water-flow testing. The cross-
Site in California. The use of LOdethanol pro- feed valve passively allows the transition of pro-
pellants provided for technology improvements pellants from one tank set to another, while the
for reusable, non-toxic, small-thruster applica- engines are burning, by simply opening the iso-
tion. The use of the Pt-Ir alloy avoided the need lation valves on the second set of tanks. Such
for lining the chamber with coatings, which passive systems required no complicated control
typically lack durability. These advancements circuitry, improving both safety and reliability.
should lead to increased safety and decreased
operational costs (ref. 4). PR&T also focused on fabricating combustion-
chamber liners using the Vacuum Plasma Spray
(VPS) process - creating liners in a variety of
shapes and sizes with GRCop-84. The tech-
nique’s application also extended to producing a
cross-sectional gradient of material, avoiding the
need for a separate protective liner for the
chamber and nozzle. After more than 100 hot-
fire tests, the subscale, 5,000-pound-thrust-class,
thrust-chamber assembly showed no signs of
degradation. The same technology is also being
evaluated for coatings on injector faceplates.
Fig. 9: PT&I LOdethanol RCS thruster
Demonstrations of miniaturized leak-detection
sensors supported objectives of significant im- Fig. 10: GRCop-84 rocket-engine components
provements in safety and operations. The self-
contained sensor system, approximately the size AIR-BREATHING PROPULSIOX
of a postage stamp, allows for easy installation of
a device to detect the fuel/oxygen ratio, thereby The NGLT projects for air-breathing-propulsion
defining the explosion hazard at a given location. technology originated with ASTP. Two projects
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had plans to culminate in ground-based testing rametric testing proved the viability of the in-
of engine concepts that could have led to the jectorirocketlstrut concept.
selection of one of propulsion systems tor in-
flight demonstration. The third project provided
support for both of these through investments for
relevant technology challenges.
Rocket Based Combined Cycle
A Rocket-Based, Combined-Cycle (RBCC) en-
gine has the potential to provide thrust from one
integrated, singleduct system for a vehicle that
takeoffs like an aircrafl and ultimately deliver a
payload to orbit. An RBCC engine relies on
rocket engines buried within a duct to fire as rock-
ets to produce augmenting, induced-flow thrust at
low speeds and then fire fuel-rich through tran- Fig. 11: RBCC strut-injector test
sonic speeds to produce induct ramjet combus-
tion. Adjustments to the RBCC duct geometry Turbine-Based, Combined Cycle
and fuel-injection system allow it to operate as a
scramjet from about Mach 5 to flight in the upper The Turbine-Based, Combined-Cycle (TBCC)
atmosphere, where it reverts to operate purely in engine was conceived as an accelerator engine
rocket mode. The primary advantage of this con- for propulsion from takeoff through Mach 4 (ref.
cept is that “one engine” serves throughout the 6). TBCC application is projected for a winged,
entire speed range; the challenge is to achieve the first-stage vehicle with a scramjet for higher-
requisite levels of performance and operability speed propulsion; that first-stage vehicle would
throughout all modes of operation. carry a rocket-powered second stage to its high-
speed, high-altitude launch point from which to
The RBCC Project selected one type of RBCC- place a payload into orbit.
engine concept to be carried to a high state of
technology development - through large-scale, The project would have culminated in a system-
ground-based demonstration - as the “Integrated level ground test of a TBCC engine at Mach-4
Systems Test of an Airbreathing Rocket” flight conditions. The test results would not
(ISTAR). The ISTAR concept selected 90-percent only have demonstrated turbine performance and
hydrogen peroxide and hydrocarbon fuel as pro- durability for space access but also matured
pellants for ‘’thsters” integrated into struts (e.g., higher-risk, turbine-engine technologies. The
strut-rockets) internal to the combustion chamber. ground-test results were expected to validate
(ref 5). The ISTAR configuration evolved to en- performance while providing an initial, system-
able airhndengine integration for a stub-winged, level database for dealing with performance,
lifting-body vehicle with externaVforebody com- operability, and durability issues.
pression and external/aft expansion.
TBCC Project planning assumed progressing to
hsk reduction for the thrusters led to designing flight-tests with a flight-qualified version of the
and testing a full-scale injector. The last of a TBCC engine. The resulting X-43-class of
series of thruster tests addressed injector-face- flight vehicle could explore critical develop-
heating, performance and combustion-chamber mental issues -- especially challenges associated
heat flux. The figure shows testing of long- with propulsion/airframe integration and with
duration unit representing one full-scale thruster transition from a “lower-speed” propulsion sys-
out of a flight demonstrator scale strut-rocket - tem (e.g., TBCC) to a scramjet engine.
with a calorimeter installed. Results of the pa-
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The TBCC project passed successfully through a and seals focused on meeting mission-based re-
preliminary-design review with General Electric quirements while minimizing thermal-
Aerospace Engines (GEAEj as the prime con- management issues (ro yieid weighr savings ar
tractor. Computer-based analyses of engine de- the vehicle level). The successful tests of mag-
sign supported the earlier projects of netic bearings at in-flight temperatures and ro-
performance. This engine concept was referred tating speeds were important since that concept
to as the Revolutionary Turbine Engine (RTA). avoids direct-contact transmission of loads ex-
cept briefly at engine start-up and shut-down.
Propulsion Research and Technolorn
The Propulsion Research and Technology
(PR&T) Project has explored new technologies
for materials, structures and propulsion-system
concepts. PR&T addressed those technologies
to support a full-scale-vehicle demonstration of
a hypersonic, air-breathing, propulsion system
by year 2015.
Fig. 12: RTA engine concept
By the end of FY03, the Cooled Panels and
The augmenter increases thrust, in an after- Ducts team fabricated five 2.5-by- 10-inch, ac-
burner-type mode, in addition to thrust provided tively cooled, ceramic-matrix-composite (CMC)
by the turbine-based flow. RTA tests at GEAE panels. Such panel concepts could lead to dura-
of a trapped-vortex flame holder addressed in- ble, light-weight heat exchangers as walls of a
stability problems associated with traditional hypersonic, air-breathing engine. The tests of
augmenter designs. Test results showed stable three panels accumulated more than 5-minutes
operation of the proposed flame holder over the of test time at hypersonic-propulsion-flowpath
range of conditions anticipated for flight. In conditions. The team fabricated, instrumented,
addition. post-test data analysis has shown good and tested one of the concepts as a 6 by 30-inch,
matching to pre-test prediction. thereby validat- actively cooled, CMC panel. Prior heat ex-
ing CFD modeling/design tools for the full aug- changers, built from NARloy-Z material, were
menter design. relatively heavy and demonstrated less than a
50-mission life -- a factor of ten less than the
goal.
PR&T activity for high-temperature, lightweight
structures succeeded in developing a new, high-
temperature polyimide (HFPE) that exceeds the
prior temperature limit by 200°F. This material
capability can reduce the weight of supporting
structures for propulsion systems without sacri-
ficing mechanical properties.
The PR&T instrumentation team developed thin-
film sensors that yield simultaneous readings of
Fig. 13: Flame stability rig installed in test cell surface temperature, strain, and heat flux on ce-
ramic or metallic components, for use at 590°F.
This project also invested in several long-term They achieved a reductions in minimum sensor
technologies relevant to the extreme internal insulation thickness from 100 to 10 microns plus
temperatures associated space-access missions. the capability to simultaneously measure surface
Work on high-temperature materials, bearings,
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- .
strain, temperature and heat flux, while doubling “acreage.” His paper also provides information
life to 200 (temperature-based) cycles. on propellant-tank sub-projects - one tank built
of iiietallic iiiateiki aiid aiiuiht-r uf wmposiit-
material. The list of tasks on hotlintegrated
structures: includes sensors, active cooling, con-
trol surfaces, primary structures and ‘‘multi-
functional” structures.
Airframe Subsvstems
NGLT investments in vehicle subsystems in-
clude developmental work in areas such as
power, actuation, avionics and health manage-
MEMS sensor die bonded ment. One of the main focuses in this area was
to enable an “all-electric” launch vehicle by in-
corporating electric actuators, thereby eliminat-
Fig. 14: PR&T valve leak sensor ing conventional hydraulic systems. Meeting
the peak-power demands of the electric actuators
The subproject for advanced propellants, in col- required investments in technologies for the as-
laboration with the Air Force Research Labora- sociated power system.
tory, has demonstrated the feasibility of using
Rp-1 in place of JP-7 in combined-cycle propul- One of the power technologies developed under
sion systems. (JP-7 is no longer in production.) NGLT was for Proton Exchange Membrane
Tests simulating the cooling loads required for a (PEM) fuel cells (ref. 8). The PEM fuel-cell
scramjet showed that RP-1 could provide endo- technology-development program completed
thermic heat-sink capability equal to that of JP-7. Phase I development of two 2-5kW breadboard
PEM-fuel-cell power plants; Phase I1 is pro-
AIRFRAME SYSTEMS AND VEHICLE gressing through development and environ-
OPERATIONS mental test of a 7-1 O kW engineering-model unit.
This section of the paper covers NGLT hard- PEM fuel-cell technology offers many potential
ware-based technological achievements for air- advantages over existing alkaline fuel cells: im-
frame systems and vehicle operations - proved performance, lower weight and volume,
essentially the entire array of technologies be- longer life, reduced hazardous materials, fewer
yond those for engines and closely associated critical failure modes and simpler ground proc-
propulsion systems. In a few cases, however, essing. While this work benefits from on-going
adaptations of the following technologies could fuel-cell development for terrestrial use,
find application in “propulsion” projects too. NASA’s focus is on the space-unique aspects of
PEM fuel cells and on long-life operation with
Airframe Structures oxygen instead of air.
This particular paper does not cover NGLT air- NGLT is developing other power technology -
frame-structures technologies, which are the a “non-toxic” Turbine Power Unit (TPU), in
subject of a concurrent paper, given by Dr. which hydrogen-oxygen gas generator drives a
David Glass (ref. 7). Dr. Glass’s paper covers turbo-alternator. Unlike the Shuttle Auxiliary
the key NGLT areas of thermal-protection sys- Power Unit (MU), that uses toxic propellants to
tems (TPS), propellant tanks, hothntegrated drive a hydraulic pump, the NGLT-funded TPU
structures and tools for both design and integra- drives a generator to produce electrical power.
tion of airframe structures. NGLT investments Applications of the TPU technology include
in TPS technology made advances for leading power for high-horsepower, all-electric actuation
edges, seals and large external surfaces, e.g., system or electric-pump hydraulic system for
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- .
launch-vehicle flight controls. Potential advan- commands) and transmitted telemetry data con-
tages of the new concept include elimination of taining tracking data, plus safety-specific health
and in&icztsis.
-:....t~ U A I&iu..~-\.Ii..i n, -i-bd.u.--uAb bu f**i-i,u,-uA uu prGCeSSiGg, iiE- 71%11-b "D.b-.b-,U-.-IdlU -1l-la;J,W-l. -nD.r ,Y-~,.J-..rL blll,
proved reliability and reduced critical failure a Range User system, provides broadband com-
modes. The schedule of the prototype unit calls munications for voice, video, and vehi-
for performance tests and completion in clelpayload data.
September 2004.
In an initial flight demonstration, STARS met all
In the area of electric actuation, NGLT has sup- objectives, successfully demonstrating its ability
ported development of a high-horsepower, fail- to maintain satellite-communication links
ure-tolerant, electro-hydrostatic actuator (EHA). between the Tracking and Data Relay Satellite
Potential applications for launch vehicle appli- System (TDRSS) and an F-15 aircraft in dy-
cations range from main-engine, thrust-vector namic flight. The Range Safety telemetry
control to movement of flight-control surfaces. passed its data to four NASA sites (Dryden
When the 50-Kp actuator completes testing in Flight Research Center, Kennedy Space Center,
September 2004, it will be one of the largest, Goddard Space Flight Center and Wallops Flight
most powerful electric actuators ever developed Facility) for monitoring. NASA Langley Re-
for space applications. search Center, NASA Wallops Flight Facil-
ity, NASA White Sands Complex, and the Air
Force Eastern and Western Ranges also partici-
pated in the overall effort.
Aero-thermal Testing
NGLT investments led to development of im-
proved, optical-based systems for assessing hy-
personic, aerodynamic-induced heating. A
NASA team completed calibration and acquired
prototype data following the development of a
faster, high-accuracy, higher-resolution, phos-
Fig. 15: Prototype, 50hp phor-thermography system for use during wind-
electro-hydrostatic actuator tunnel testing. The team also integrated an in-
frared-image-acquisition system into data-
Launch-Range Ouerations reduction software. (The same team provided
support with these tools for aero-heating data
Space-Based Telemetry and Range Safety acquisition during wind-tunnel tests for the Co-
(STARS) is a space-based communications sys- lumbia-accident investigation.)
tem to relay telemetry between launch vehicles
and the ground. It serves as a proof-of-concept CONCLUDING REMARKS
project to determine if operational costs can be
reduced and operational flexibility increased. NASA's NGLT Program is completing research
High reliability and extensive coverage are key and development for a wide range of technolo-
parameters for judging success. gies to enable future generations of space-launch
transportation systems. Objectives for improv-
STARS is composed of two major systems. ing safety, reliability, and affordability extended
One, the Range Safety system, includes: (a) a to providing improvements for application to a
versatile, low-power, multi-channel transceiver, vehicle concepts and architectures ranging from
(b) a custom-built flight processor for command- those for reusable systems to those for fully ex-
and-data handling, and (c) a commercial Global pendable launchers. Many projects achieved
Positioning System receiver. The transceiver significant results as evidenced by results, such
received an up-link (of flight-termination-system as those of rocket-system firings, thermo-
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