Airbreathing Hypersonic Technology Vision Vehicles and Development Dreams

17 Pages·1999·0.54 MB·English

by  McClinton, C. R.

AI A A 99 - 4 9 7 8
Airbreathing Hypersonic
Technology Vision Vehicles
and Development Dreams
C. R. McClinton, J. L. Hunt, and R. H. Ricketts
NA S A Langley Research Center, Hampton VA
P. Reukauf
NA S A Dryden Flight Research Center, Edwards, CA
an d
C. L. Peddie
NA S A John Glenn Research Center, Cleveland, OH
9th International Space Planes and Hypersonic
Systems and Technologies Conference
an d
3rd Weakly Ionized Gases Wor k s h o p
November 1-5, 1999/Norfolk, VA
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics
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AIAA 99-4978
Ai r b r eathing Hypersonic Tec h n o l o g y
Vision Vehicles and Development Drea m s
C. R. McClinton, J. L. Hunt, and R. H. Ricketts
NASA Langley Research Center, Hampton VA
P. Reukauf
NASA Dryden Flight Research Center, Edwards, CA
C. L. Peddie
NASA John Glenn Research Center, Cleveland, OH
ABSTRACT
Significant advancements in hypersonic airbreathing vehicle technology have been made in the country’s research cen-
ters and industry over the past 40 years. Some of that technology is being validated with the X-43 flight tests. This
paper presents an overview of hypersonic airbreathing technology status within the US, and a hypersonic technology
development plan. This plan builds on the nation’s large investment in hypersonics. This affordable, incremental plan
focuses technology development on hypersonic systems, which could be operating by the 2020’s.
IN T R O D U C T I O N over 40 years. Within the U.S. alone, NASA, DOD
(DARPA, U.S. Navy and USAF), and industry have par-
Man’s search for higher speeds, as for flight itself, is lim- ticipated in hypersonic technology development. Over
ited by the propulsion system required for the task. The this time NASA Langley Research Center continuously
hypersonic (Mach number greater than 5, or 3,400 miles studied hypersonic system design, aerothermodynamics,
per hour) rocket powered X-15 aircraft demonstrated propulsion, high temperature materials and structural
flight up to Mach 6.7 (6.7 times the speed of sound, or architectures, and associated facilities, instrumentation
about 4600 miles/hr.) in the 1960’s. The fastest aircraft and test methods. These modestly funded programs were
propelled by an air breathing engine, the SR-71 substantially augmented during the National Aero-Space
Blackbird, only reaches speeds slightly over Mach 3 Plane (X-30) Program, which spent more than $3B
using a turbojet engine. Ramjets have been utilized for between 1984 and 1995, and brought the DOD and other
missile propulsion at speeds up to about Mach 5. Winged NASA Centers, universities and industry back into hyper-
rocket powered vehicles, such as the Orbital Sciences sonics. In addition, significant progress was achieved in
Corporation Pegasus, have been utilized for hypersonic all technologies required for hypersonic flight, and much
flight within the atmosphere to improve launch efficien- of that technology was transferred into other programs,
cy. Not unlike the challenge facing Orville and Wilbur, such as X-33, X-37, X-43, etc. In addition, technology
dramatically improved engine performance is required for transfer impacted numerous other industries, including
hypersonic flight. In fact, efficient hypersonic flight with- automotive, medical, sports and aerospace.
in the earth’s atmosphere requires a different engine, one
that uses the oxygen within the air for combustion of the Recently, NASA initiated several hypersonic technology
fuel. Hypersonic airbreathing propulsion also provides programs: the LaRC/DFRC Hypersonic X-Plane
the option to “fly” to orbit. This air breathing engine Program, Hyper-X, in 1996; the GRC Trailblazer in
option has been considered and studied for over 40 years, 1997; and the MSFC Advanced Reusable Transportation
but not realized because of low technology maturity as ART technology program in 1997, Bantam in 1997,
compared to the rocket. Recently, as hypersonic air- Spaceliner-D and finally, just “Spaceliner” in 1999. Of
breathing technology matured, and space access require- these programs only Hyper-X and ART build on the
ments continued to grow, the world started seriously con- technology gains of the X-30 program. The Hyper-X
sidering airbreathing propelled vehicles for space access. Program focus is to extend scramjet powered vehicle
technology to flight, elevating as much technology as
NASA, DOD, the U.S. industry and global community possible, and validating, in flight, the design systems,
have studied scramjet-powered hypersonic vehicles for computational fluid dynamics (CFD), analytical and
Copyright © 1999 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the
United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under
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experimental methods required for this complex multi- system. The impact on the overall system is the only
disciplinary problem. The smaller ART program focus is adequate measure of goodness. An example of this is the
on RBCC wind tunnel testing of alternate airframe inte- development of a combustor performance index, namely
grated scramjet flowpath concepts. thrust potential (ref. 1). This parameter was developed
during the 1990’s, as scramjet engine designers realized
Likewise, within the DOD several hypersonic programs that combustion efficiency was not an adequate measure
are emerging. The USAF AFRL Hypersonic Technology of the combustor design. A method of quantifying the
(HyTech) program, the Defense Advanced Research combustor impact on overall engine performance was
Projects Agency (DARPA) Affordable Rapid Response required, and exergy, as applied in the literature, did not
Missile Demonstrator (ARRMD) Program, The USN provide an optimum design.
Rapid Response Missile Program and the Army Scramjet
Technology Development Program. In addition, the Formal system analysis procedures are required for vehi-
USAF Aeronautical Systems Center, in collaboration cle design and performance analysis. The LaRC design
with the Air Combat Command, is conducting a Future process is illustrated in figure 1. Engine and aerodynam-
Strike study, which focuses on hypersonic aircraft. ic performance, structural requirements, weights, and
flight vehicle performance (mission or trajectory) are
With this renewed interest in hypersonic vehicles, evaluated, always “closing” on take off gross weight for
requirements are being developed which can only be met the specified mission. This design process can be execut-
with hypersonics systems. These include the USAF ed using four basic levels of analysis.
CONUS-based Expeditionary Aerospace Force con-
cepts, and reduced cost to orbit. The lowest level, designated “0” in Table 1, does not
require a physical geometry. The level zero analysis uti-
This paper discusses the potential of hypersonic airbreath- lizes ideal engine cycle performance, historical L/D and
ing technology for endo- or exo-atmospheric vehicles (air- Cd values for aerodynamic performance, design tables (or
planes and space planes). The status of hypersonic tech- weight fractions) for structure and components weight,
nology, the significance of the X-43 flights to technology “rocket equation” for flight trajectory, and estimates for
advancements, and a method of filtering vehicle propo- packaging. This analysis does not require a specified
nents’ claims are also discussed. Finally, a plan to effi- vehicle, engine flowpath or systems definition. All higher
ciently demonstrate hypersonic technology is presented. levels of analysis require a vehicle, engine flowpath
shape and operating modes, system definition, etc.
HYPERSONIC TECHNOLOGY STATUS
The next level of system analysis, referred to herein as
This section discusses the status of hypersonic technolo- Level 1, utilizes uncertified cycle performance and/or
gy—with the goal of showing significant advancement; CFD, impact theory, unit or uncertified finite element
thus, justification for continuing to push hypersonic tech- model (FEM) weights, single equation packaging rela-
nology development to flight. tions, and energy state vehicle performance.
(Certification is discussed in the next paragraph). This
System Analysis and Conceptual Designs level of analysis does not capture operability limits, and
thus has large uncertainties.
The key to any hypersonic vehicle development or tech-
nology program is a credible preliminary system analysis Level 2 analysis utilizes “certified,” methods; i.e., the
to identify the technical requirements and guide technol- user has sufficient relevant experience. This level uses the
ogy development. The X-30 program provided an excel- same methods for propulsion, aerodynamics, structure
lent training ground for system design, analysis and and weights (but certified), trimmed 3-DOF (degree of
development of hypersonic technology. The complexity freedom) vehicle performance analysis and multiple
of the hypersonic airbreathing system and the small equation, linear or non-linear packaging relations.
thrust margin dictate that a thorough system analysis be Certification is only achieved by demonstration that the
performed before any focused technology development methods used work on the class of problems simulated
is started. Over the past 40 years, many bright individu- (this relates to the method, as well as the operator apply-
als and companies brought forward vehicle, engine or ing that method). For example, at level 2 analytical mod-
structural concepts, which at first blush appeared to be els utilize corrections for known errors, such as inlet mass
an excellent solution. However, due to the highly inte- spillage, relevant empirical fuel mixing models (ref. 2),
grated nature of this class of vehicles, an excellent com- shear and heat flux models (ref. 3), etc. This empirical
ponent solution is not always beneficial to the overall approach is based on experimental (wind tunnel tests,
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structural component tests, etc) data. Higher level meth- Boeing Phantom works a clear trend in vehicle configura-
ods (CFD, FEM) are used to refine the vehicle closure. tions emerged (ref. 4). These trends are consistent with
trends noted by the X-30 contractor team. Hypersonic air-
The highest design level (level 3) is achieved only by hav- craft configurations have converged to either lifting body
ing a significantly large fraction of the actual vehicle man- or wave rider configurations, as illustrated in figure 2.
ufactured and tested. Wind tunnel and other ground testing Lower speed, Mach 5 cruise vehicles tend to favor the
provide less verification than flight tests. Although numer- wave-rider configuration, whereas higher speed cruise
ous components have been built and ground tested, flight configurations favor the lifting body. This is partially due
data is required for the highest level of design. This has to fuel selection and operation range. The lower speed
not yet been done for a hypersonic airbreathing vehicle. vehicle operates with hydrocarbon fuel, resulting in high-
er wing loading, favoring the wave rider configuration.
Whatever the level of system analysis, closure is
achieved by sizing the vehicle so that the propellant frac- Numerous classes of vehicles have been evaluated for
tion required (for the mission) is equal to the propellant horizontal take-off and landing, single or two-stage
fraction available (packaged within the sized vehicle). access to space (ATS) missions. These include cone,
However, the reported closure weight is only as good as wing-body, high-fineness, lifting body and other “novel”
the lowest level of analysis used in the “closure.” concepts, such as inward turning (“funnel”) inlet designs.
The lifting body configuration is the clear choice for
Quantifying the uncertainty associated with the “level” these missions (see figure 3). These vehicles are some-
of closure is difficult. Clearly the higher-level design what different from hypersonic cruise configurations;
methods have less uncertainty—level 3+ would be near mostly in the relative size of the engine package.
zero. On the other extreme, level “0” analysis includes so Vehicles designed for accelerating ATS missions require
much uncertainty as to be non-quantifiable. Level 0 as much air capture as possible.
analysis simply indicates that the objectives can be
achieved, but does not indicate how. Any configuration “An early lesson of high speed flight was that proper
shown with this level of analysis is meaningless. aerodynamic integration of the ramjet or scramjet with
the remainder of the vehicle is critical to success…” (ref.
Level 1 analysis is the first step toward a real vehicle 5). This comment is generally associated with podded
definition. However, the methods used are not validated, engines. However, it applies equally to engine integra-
so the uncertainty in each discipline is large, and the tion which ties the scramjet engine mass capture to wing
closed TOGW uncertainty is huge. The potential for over area, such as the X-30 “Government Baseline” wing-
predicting the TOGW is negligible. However, because of body configuration, or to fully wrapped configurations
the small thrust margins associated with hypersonic air- which utilize ineffective conical compression surfaces
breathing systems, and potential operability limitations, (providing high surface area/unit capture, and excessive
the potential for under prediction by a factor of 10 is not engine weight. The two-dimensional high-fineness ratio
unusual. lifting body configurations provide one-sided compres-
sion, which efficiently pre-compress the inlet air. Top
Level 2 design methods retain significant uncertainty— and side panels add little to the profile drag, and low
again because of the overall level of this technology local surface pressure results in relatively low heating
development and small thrust margin. Therefore, it is and shear drag. This approach has been successful for a
prudent to carry 15% dry weight margin, 5 to 10% thrust wide range of vehicle applications (figures 2 and 3).
or ISP margin and +/-10 to 15% on the predicted aerody-
namic forces. The resulting impact on TOGW is signifi- Engine selection for air breathing hypersonic vehicles is
cant, but quantifiable. Estimates of +20 / -5% on TOGW another mission dependent design parameter. If the vehi-
uncertainty are considered reasonable. cle is truly hypersonic, a scramjet is required (except for a
Mach 5 vehicle, which can operate with a classical ramjet
The color codes presented on table 1 will be used herein and possibly with a Revolutionary Turbine Based Engine
to identify the quality of vehicle closure when discussing (RTBE, see below). For this discussion, the term scramjet
current configurations. refers to a dual-mode scramjet, or a scramjet that can
operate in supersonic, subsonic and mixed
Vision Vehicles— supersonic/subsonic combustion modes, without the use
of a second minimum, or nozzle throat. This is the type of
Based on over 30 years of detailed hypersonic system scramjet which has been studied for 40 years in the U.S.,
analysis at NASA Langley Research Center and the and is traditionally referred to simply as a scramjet (SJ).
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For single stage to orbit (SSTO) access to space (ATS) tem provides a substantial benefit over the ESJ/SJ, R
missions, a rocket is required for the final stage of combined cycle engine system. It is interesting to note
boost, orbital insertion, orbital maneuvering and deor- that lightly loaded two-stage to orbit (TSTO) vehicles
bit. SSTO is highlighted because of the perceived are not as sensitive to low speed engine selection.
reduction in complexity and operating cost vis-à-vis a
two or more stage system. For airplane configurations, which are required to oper-
ate over a large speed range within the atmosphere, the
Several options are available for low-speed propulsion, low speed system requirement dictates utilization of an
and are defined by the following notation in ref. 6: efficient airbreathing engine. Therefore, based on
• Air-augmented rocket (AAR), today’s technology, a turbine-based engine is mandatory.
• Rocket ejector with ramjet after burning (ERJ), for
configurations with a second minimum (nozzle throat). Structural concepts for the vision vehicles also con-
• Rocket ejector scramjet (ESJ), for configurations verged as the analysis fidelity improved. Current con-
without a second minimum. cepts include cold integral graphite/epoxy LH2 tanks
• Pure airbreathing, such as (developed under X-30 and used in the X-33) with a car-
- Turbojet (TJ) bon/silicon-carbide “insulated multi-wall insulation”
- Turbofan (TF) (IMI) thermal protection system (TPS) on the windward
- Turboramjet (TRJ) which is a combined cycle in its side, and tailored advanced blanket insulation (TABI) on
own right. It includes turbofan engines operating the lee side. This provides a light weight TPS with
with “wind milling” compressors at high speed. durable external skin. Wing and tail structure is titanium
• Air-turborocket (ATR), where a compressor or fan is metal matrix composite, developed for the X-30.
driven by a rocket or monopropellant-decomposition.
The engine primary structure is graphite-polyimide
The propulsion system required for SSTO operation can (being demonstrated on the X-37). Regeneratively
then be defined as either a combined cycle engine or cooled copper, aluminum, and high temperature superal-
combination engine. Combined cycle engines are single loy panels are utilized in the engine, and the engine and
flowpath, integrated engines capable of operation in two vehicle sharp leading edges are cooled by the impinge-
or more modes, such as: ment process developed and verified for the X-30.
- Low speed: ejector-ramjet/mid speed: ramjet/high
speed: rocket. This is referred to as an ERJ/RJ/R SSTO vision vehicle capabilities are summarized in fig-
combined cycle engine (popularly called a RBCC). ure 4. This figure illustrates the TOGW for several vehi-
- Low speed: ejector—scramjet/mid speed: scram- cles, for either “LEO” or Space Station Freedom. The
jet/high speed: rocket. This is an ESJ/SJ/R com- small (2000 lb.) payload X-30 vehicle which started as
bined cycle engine (popularly called a RBCC). the 50,000 lb. TOGW “Government Baseline” is shown
- Low speed: air augmented rocket/mid speed: as the red symbol. The X-30 program started with a
scramjet/high speed: air augmented rocket/orbital: vehicle analyzed with a combination of level 0 and 1
rocket. This is referred to as a AAR/SJ/R (popular- methods. After 8 years and $3+B of technology develop-
ly called a RBCC). ment and configuration refinement the final X-30 config-
uration TOGW was about 500Klbs, determined with a
Combination engines can also be used for SSTO, such as: predominately level 2 analysis. The final X-30 vehicle is
- Low speed: turboramjet/mid speed: scramjet/high shown as light green because of one non-verified critical
speed: stand alone (tail) rocket. This is referred to as structural element. The other SSTO vehicles from figure
TRJ, SJ, R combination engine (popularly, but incor- 2 are also shown on figure 4. The vehicles with liquid
rectly called a TBCC—the TRJ is in fact a TBCC). oxygen are heavier than current runway limits.
- Low speed: ESJ/mid speed: scramjet. This is
referred to as an ESJ/SJ combined cycle engine. Hypersonic Propulsion
When used with a high-speed external (tail mount-
ed) rocket, the total package is referred to as a Hypersonic propulsion has been studied continuously at
ESJ/SJ, R combination engine or propulsive system. NASA Langley Research Center for the past 40 years.
Technology development has focused on design meth-
The effect of low-speed engine selection on TOGW is ods, experimental methods, experimental databases, air
illustrated for the SSTO and TSTO configurations in fig- vehicle configuration and flowpath designs. Although
ure 3. For the SSTO heavy lift (25,000 pounds to space engine study emphasis was on the scramjet flowpath, it
station) mission, the TRJ, SJ, R combination engine sys- also included structural concepts, designs and tests.
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Design methods range from cycle analysis (ref. 7) to formed detailed structural design of two second-genera-
complex 3-D Navier Stokes analysis (ref. 8). This work tion regeneratively cooled concepts (the three-strut
included development of full and reduced finite rate engine, ref. 12, and the CIAM-NASA axisymmetric
chemistry models for scramjet operating conditions (for scramjet (ref. 13) which was wind tunnel and flight test-
both combustor and nozzle flow), evaluation and selec- ed by CIAM). In addition, NASA LaRC was involved in
tion of turbulence models (ref. 9), and other CFD meth- third generation engine design activities, with the X-30,
ods required to accurately model the flow. including design and tests of advanced regeneratively
cooled, film cooled, and transpiration cooled panels and
Scramjet test capability also has been continually improved leading edge cooling concepts to Mach 15 flight heat
at LaRC over the past 40 years. Combustion, arc and shock loads (refs. 14 and 15).
heated facilities (figure 5) were developed for testing com-
ponents or engines from Mach 1 to about Mach 20, over Hypersonic Aerothermodynamics
the scramjet operating flight trajectory. Both component
(inlet, combustor and nozzle) and integrated flowpath tests Hypersonic aerodynamics and aero heating has been
have been performed. Facility, instrumentation and test studied at NASA Langley Research Center for over 40
techniques were tailored to scramjet testing requirements, years. Technology development focused on design meth-
and understanding the scramjet flow physics (ref. 10), ods, experimental methods, and experimental databases
including airframe integration effects. Between 1975 and (ref. 16). Design methods range from simple analytical
1999, tests were performed on 21 scramjet engine flowpath to complex 3-D analysis. This work included: develop-
models. More than 3500 tests were performed, providing ment of full and partial finite rate chemistry models, for
about 30 hours of testing, equivalent to about 4 trips operation in continuum and non-continuum flows; evalu-
around the world at Mach 5. Results from these tests veri- ation and selection of turbulence models and other CFD
fied scramjet powered vehicle performance and were solution requirements to accurately model the heat trans-
instrumental in the 1984 initiation of the X-30 Program. A fer; and, development of CFD and analytical models for
summary of some of the work leading up to the X-30 pro- complex flowfield issues such as shock-shock interaction
gram is included in ref. 10. heating, wing gaps, etc.
The NASA Langley Research Center played a key role Construction of the hypersonic facility complex at LaRC
in the generation of the large X-30 database for dual was started in the late 1940s. These facilities (figure 6)
mode, hydrogen-fueled scramjet operation from Mach 3 supported numerous programs. Some of these include
to 16 and low speed engine operation from Mach 0 to 3. the X-15, the Space Shuttle, X-30, Pegasus, X-33, X-34,
Component investigations included dual-mode scramjet X-38, X-43 and the X-37, starting this year. These facili-
combustor tests from Mach 3 – 20, ducted rocket tests at ties (figure 6) span the Reynolds number range required
Mach 15, inlet tests from Mach 1.5 to 18, and nozzle for Space Shuttle, X-38, and X-43 vehicles from Mach 6
afterbody tests from Mach 0 to 14. This large database to 20. Much of the work performed in these facilities,
was used to verify and improve both analytical and CFD and the aerodynamic databases developed, were validat-
based design methods. ed with traditional blunt lifting body flight data. The
upcoming X-43 flight will provide the first flight valida-
The Hyper-X scramjet engine is currently being tested in tion of these facilities for sharp, high-fineness hyperson-
all the engine test facilities shown in figure 5 to provide ic lifting body configurations.
pre-flight scramjet performance data for flight test risk
reduction, and to provide data for comparison to flight LaRC led the X-30 aerodynamic program which provid-
data and for validation of wind-tunnel test methods. ed extensive databases for numerous hypersonic configu-
Continuous improvements in scramjet engine flowpath rations including the final X-30 configuration.
designs are demonstrated by these tests; the X-43 engine Unpowered models of the X-30 final configuration were
has the highest performance and operability, and shortest tested over the Mach range of 0.25 to 17. Aerothermal
length (weight) of all dual mode scramjet flowpaths test- models were tested at Mach 3.5, 6, 10, 10-17, and 18.
ed to date. The largest uncertainties following these tests Powered models were tested over the Mach 3.5 to 10
are hypervelocity scramjet performance, flight vs. wind range at NASA LaRC in the Unitary Plan Wind Tunnel,
tunnel performance and operability, integration effects 20" Mach 6 and 31" Mach 10 tunnels. Following the X-
and geometric scaling. 30 program, the largest aerodynamic uncertainties are
boundary layer transition, powered effects and external
NASA LaRC also tested a first generation, “flight- burning. The Hyper-X program is resolving powered
weight” liquid-hydrogen cooled engine (ref. 11) and per- effects and providing additional information on natural
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and forced boundary layer transition. The Program is Hy p e r - X
also evaluating the possibility of generating base burning
data, with the X-43 low-pressure remnant hydrogen. The primary goals of the Hyper-X Program (ref. 18) are to
validate the airframe integrated, dual-mode scramjet pow-
Materials and Structural Architecture ered vehicle in flight and provide databases for validation
of design methods and tools. This will be accomplished
Many hypersonic structural concepts and materials were using data from the X-43 vehicle (figure 8) under powered
developed and/or tested at NASA Langley Research conditions at Mach 7 and 10, and unpowered conditions
Center (see figure 7) over the past 30 years. down to subsonic flight. In preparation for these X-43
flights, refinement of the vehicle design using optimization
Primary airframe structures for advanced space trans- methods was required to assure that the small, compact X-
portation vehicles, such as, X-33, X-34, and X-37 are 43 vehicle accelerates. In addition, every detail of the
being constructed of graphite composite materials. hypersonic system was evaluated, including the high Mach
Graphite epoxy is currently being used; however, higher number, high dynamic pressure stage separation. The most
temperature bismaleimides and polyimides are being extensive hypersonic aerodynamic, propulsion and thermal
considered for other vehicles. These are “cool” struc- database ever generated for this class of vehicle is being
tures, which must be protected with a thermal protection used to develop autonomous flight controls, size TPS and
system (TPS). Large airframe structural sections fabri- reduce risk for this first ever scramjet-powered hypersonic
cated and tested for the X-30 were used to anchor the flight. This flight will usher in the hypersonic century.
FEM models (Ref. 17) which are used to design the
vision vehicles illustrated in figures 2 and 3. TECHNOLOGY DEVELOPMENT PLANS
Thermal protection systems (TPS) must be used to keep The Hyper-X Program Office (HXPO), under the direction
the airframe structure within allowable temperature limits. of NASA Code R, is developing a plan to continue the far-
Tailored advanced blanket insulation (TABI) is being used ther advancement of hypersonic vehicle technology after
for the lower temperature leeward side of the airframe; the X-43 flights. This plan focuses on elevating the
alumina enhanced thermal barrier (AETB) is being used Technology Readiness Level (TRL) of all airbreathing
on the higher temperature windward side. Metallic TPS is hypersonic vehicle related technology to TRL 6 (system
being developed at Langley Research Center where it has demonstration in flight environment) as fast and efficiently
been tested to high temperatures and Mach 7 flow in the as practical. The technology to be addressed is applicable
8-ft High Temperature Tunnel. (ref. 14) to both endo- and exo-atmospheric missions (represented
by the configurations that are shown in figure 2 and 3).
A carbon-carbon wing elevon was developed under the X-
30 program (ref. 17). This control surface was designed to For this technology development plan, the HXPO estab-
operate in a 3000 F environment. Ceramic matrix com- lished the following guidelines: 1) expedite technology
posites are being used for the wing ailerons and rudders on development; 2) leverage existing technology; 3) provide
the X-37 vehicle. Improved coatings are also being studied incremental approaches/options; and 4) provide “test
by the Hyper-X Program for Mach 10 application to sharp beds” for other hypersonic technology and programs.
leading edges. Heat pipe concepts and designs were devel- Expediting technology development is necessary due to
oped and tested for the X-30 program, undergoing stagna- limited R&D resources. As mentioned previously, scram-
tion point heating to conditions in excess of Mach 10. jet and other related hypersonic technologies have been
extensively studied; the time is right for moving to flight.
Graphite composites are being used for the liquid H2
cryogenic tanks on the X-33 and fuel tanks on the X-34 Leveraging from existing technology is critical to expe-
and X-37 (ref. 17). Foam has been added to the tank for dite development and reduce cost. This plan leverages
cryogenic insulation. from the X-43 flight research vehicle, which in turn
leveraged from the X-30 and NASA’s 40 years of basic
Analysis methods are an integral part of structural devel- research in hypersonic propulsion. In addition, the plan
opment. Methods developed for hypersonic vehicles leverages from the ongoing USAF, DARPA, and U.S.
include detailed thermal structural coupled flow-structur- Army Hypersonic Technology Programs. The USAF and
al analysis, 3-D solution methods for shock-shock DARPA programs are maturing hydrocarbon scramjet
impingement heating, thermal modeling for active/regen- technology, which is essential in reducing the cost of
eratively cooled engines, etc. These methods have been demonstrating hypersonic systems in flight. The U.S.
verified by experimental tests and hardware fabrication. Army program is addressing Mach 8-14 scramjet tech-
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nology including full pressure and enthalpy tests of the vehicle must be boosted to dual-mode scramjet takeover
Mach 10 X-43 full forebody and engine configuration. speed (Mach 4.5). It will investigate controllability
issues associated with hypersonic operation and scramjet
This technology plan provides an incremental approach mode transition. Several solid and liquid rocket boost
to reduce risk, provide budget options, and allow for early concepts are being studied both in-house, and under a
flight tests/technology demonstration. Some of the incre- NASA—AFRL Air Vehicles funded study by a
mental steps can be skipped, without adding undue risk. Microcraft–Boeing team. The vehicle will likely use a
liquid rocket and be air launched (B-52 drop) to mini-
The proposed hypersonic development plan is illustrated mize the overall size. The vehicle will land with skids
in Figure 9. The plan builds on three ongoing hypersonic and a steerable nose wheel. Because of size limitations,
technology programs: Hyper-X Phase 1, HyTech (USAF the Phase 1a vehicle uses hydrocarbon fuel for more vol-
AFRL), and ARRMD (DARPA). Hyper-X provides umetric efficient fuel packaging required to extend the
wind tunnel, analysis and flight databases, systems, and scramjet test time.
flight controls for a hydrogen-fueled, scramjet-powered,
hypersonic vehicle. This includes an extensive wind tun- This Phase 1a vehicle will utilize the X-43 and
nel database from Mach 4.5 to 10 for the engine and HyFLITE aerodynamic databases, HyTech endother-
Mach 0.6 to 10 for the vehicle aerodynamics. HyTech mic fuel-cooled engine structure, and blended X-
will provide hydrocarbon scramjet flowpath and engine 43/HyTech engine flowpath, flight systems and data-
designs and design methods, and a wind tunnel database bases. By maintaining the same basic configuration
for the hydrocarbon-fueled engine, flight-weight engine (after boost), much of the large X-43 aero propulsion
structures, and hydrocarbon fuel control systems. The database remains applicable.
DARPA program will provide low-cost flight hardware,
manufacturing techniques, and flight validation of the Hyper-X Phase 1a will be the first recovered/reused
hydrocarbon fuel cooled engine and fuel control system. hypersonic airbreathing engine propelled vehicle. This
program element will elevate the TRL of the scramjet
HyFLITE Subsonic Remotely-Piloted Vehicle (RPV) engine over a large operating range. This will add to the
airframe integrated scramjet engine flowpath validation
One additional on-going “hypersonic vehicle” activity is achieved in Hyper-X Phase 1. This vehicle will allow
the Accurate Automation Corporation (AAC) HyFLITE durability testing of hypersonic engines not currently
remotely piloted vehicle (RPV) program. The subsonic possible in hypersonic blowdown wind tunnel facilities.
RPV (figure 9) is a 12' long, density scaled replica of the The vehicle will also be designed with a limited test-bed
dual-fuel, global-reach Mach 10 cruise vision vehicle capability. For example, methods of testing alternate
(ref. 19). The vehicle is powered by two 35 lbf thrust TPS, alternate flight controls, and using hydrogen fuel
micro turbojet engines. This vehicle was fabricated dur- are being evaluated. In addition, the vehicle will be
ing Hyper-X Phase 1 by the McDonnell Douglas capable of testing other airframe integrated scramjet-
Aerospace Corporation, and later donated to Accurate based engines, and may be able to test “RBCC” engines
Automation Corporation. Flight tests have been funded over the full range from Mach 0.8 – 7. This vehicle will
by the USAF and NASA ARC. These tests are to investi- also be capable of flight above Mach 7 with limited
gate flight characteristics and control during low speed modifications, as performed on the X-43 vehicle. Figure
takeoff and landing operations. This vehicle is sharing 9 illustrates one possible configuration, using liquid JP-
aerodynamic data with the Hyper-X Phase 1 Program: LOX rocket engines with LOX drop tanks. Mach 4
using low speed wind tunnel data from the X-43, and will stage separation of the drop tanks will be designed using
provide low speed flight data to the Hyper-X Program. the stage separation tools developed for the X-43.
Hyper-X Phase 1A Revolutionary Turbine-Based Engine (RTBE)
De v e l o p m e n t
Phase 1a is intended to be the worlds first demonstration
of a reusable, flight-weight scramjet powered vehicle, A multi-agency team (NASA GRC, LaRC, DFRC and
and is planned to operate over a larger speed range than MSFC plus the U.S. Navy) has been formed, led by the
possible with the small X-43 vehicle. This vehicle will GRC, to identify the requirements and technology sta-
allow transition from dual-mode scramjet to pure scram- tus, and develop a Mach 4 to 5 capable turbine based
jet operation. Conceptual studies indicate that a larger engine for high speed expendable and crewed flight.
(25 – 30') vehicle is required to achieve this goal of The requirements provided by the Hyper-X Program
reusability and operation from Mach 4.5 to Mach 7. This are presented in Table 2. These requirements are based
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AIAA 99-4978
on the technology development plan and conceptual foot, reusable, Mach 7 vehicle with a revolutionary tur-
studies illustrated in figure 9. Several U.S. turbine bine based engine (RTBE)– dual-mode scramjet combi-
engine companies are currently evaluating/defining nation (TRJ/SJ) engine. The primary objective of this
plans to meet these requirements. In addition, NASA flight test program is elevation of the TRL for the
Langley Research Center is evaluating the Japanese TRJ/SJ engine, not just the flowpaths as done in Hyper-
deeply cooled, hydrogen-fueled Mach 6 turbo-ramjet X Phase 1, or the scramjet engine as planned in Phase
(ATREX) in the SSTO access to space vehicle illustrat- 1a. This vehicle will be designed for horizontal take-off
ed in figure 3. and landing and acceleration through engine mode tran-
sitions (turbine to ramjet, ramjet to scramjet). This
There is reason to believe that a RTBE can be developed vehicle will be used to investigate the controllability
to meet the requirements in table 2, considering the cur- issues associated with hypersonic vehicle operation
rent status of high speed component and core testing from Mach 0-7, integrated TRJ/SJ performance and
throughout the industry. The requirements for the RTBE operability, engine mode transition, durability testing,
program are consistent with a logical engine develop- powered take-off and landing, 3rd generation scramjet
ment process: starting with small scale expendable regeneratively cooled concepts, dual-fuel operation,
engines; then small limited life, non-man rated engines; and flight weight hypersonic aircraft structures. Boeing
followed by a large prototype; and finally man rated pro- completed a conceptual study of this vehicle, called
duction engines. HySID (ref. 19). Because of the size limitations, the
only way to extend the TRJ/SJ test time is by use of
Hypersonic Aircraft-Shaped Low-Speed volumetrically efficient hydrocarbon fuel. This Phase 2
Demonstrator (HLD) vehicle will utilize the X-43 aerodynamic database, a
blended X-43/HyTech engine flowpath, HyTech-
Preliminary assessment of low-speed propulsion-air- ARRMD derived endothermic fuel-cooled engine
frame integration (PAI), performance, aerodynamics, structures, the RTBE from the GRC led program dis-
stability and control, and handling qualities can be cussed above, and X-43 systems and database. Low
assessed using a relatively inexpensive X-43 shaped speed aerodynamics, handling, and scramjet engine and
vehicle which provides risk reduction before preceding systems will be supplemented by data from Phase 1a,
to an expensive hypersonic capable vehicle. A conceptu- the AAC RPV and the piloted HLD discussed above.
al assessment of this vehicle was recently completed.
This vehicle will be the first hypersonic airbreathing
Figure 9 shows a 50' long, dynamically scaled, piloted engine powered vehicle to operate from takeoff to
vehicle based on the dual-fueled, global-reach Mach 10 hypersonic speed. It will allow durability testing of
vehicle (ref. 19). The HLD is very similar in physical hypersonic engines not currently possible in blowdown,
size and performance to a T-38 Talon high performance hypersonic wind tunnel facilities. This vehicle will be
training aircraft. The aircraft will have a maximum level capable of about 5 minutes cruise at Mach 7 with on-
flight speed of approximately Mach 0.6. board fuel using the baseline TRJ/SJ engine (can not be
achieved with the Phase 1a vehicle). The vehicle will
A piloted Low Speed Demonstrator allows for safe, also be designed as a test-bed for other engine and
incremental flight envelope expansion. The inherent hypersonic systems. Hydrogen fuel tanks/systems will
ability of the pilot to do real-time test point repetition be interchangeable with the hydrocarbon system. The
and problem solving is an asset that cannot be provided vehicle will be capable of testing other airframe inte-
by autonomous or remotely piloted aircraft. A pilot-in- grated scramjet-based engines (RBCC: ERJ/RJ,
the-loop flight control system also reduces development AAR/SJ, etc.). To allow the same mission range, addi-
and operational costs. Thus, the presence of a pilot tional fuel will be required, and could be contained in
increases system redundancy and mission safety thereby drop tanks as shown for the Phase 1a vehicle. This
enhancing data return and program accomplishment. In vehicle is also large enough, and timing should be right
concert with current and planned hypersonic flight test- for a complete system demonstration of weakly ionized
ing of autonomous research aircraft, it will lead to the gas (WIG) effects. This vehicle will also be capable of
ultimate goal of a piloted hypersonic cruise vehicle Mach 10+ flight with limited modifications, as per-
which will revolutionize flight as we know it. formed on the X-43 vehicle. Figure 9 illustrates the
conceptual configuration for this vehicle. The canard is
Hyper-X Phase 2 added to the X-43 shape, to meet takeoff and low speed
requirements for the higher density of the hydrocarbon
The Phase 2 vehicle is currently envisioned as a 45- fueled system.
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The pacing technology for the Phase 2 vehicle is the Symbols & Acronyms
RTBE. The Phase 2 program can be pursued earlier
using existing turbojet engines, with transition to the AA R Air augmented ramjet combined cycle
dual-mode scramjet at Mach 2.5 to 3. This adds some (RBCC) engine
risk, and time to the Phase 2 schedule if Phase 1a is not Ac Engine frontal area
performed first. The major risks are the increased vari- AE D C Air Force Arnold Engineering and
able geometry requirements and low-speed operability of Development Center
the dual-mode scramjet. AF R L Air Force Research Laboratory
AR C NASA Ames Research Center
Hypervelocity, Structures and ARRMD Affordable Rapid Response Missile
Advanced Design Methods De m o n s t r a t o r
AR T MSFC Advance Reusable Transportation
Other technology development is required for SSTO technology program
capable vehicles. This includes: AT R Air turborocket combined cycle (RBCC) engine
• Mach 10 to 17+sc r a m j e t ; AT S Access to Space vehicle
• Mach 12 to 20 ducted rocket; CF D Computation Fluid Dynamics
• Leading edge design/validation testing; and DA R P A Defense Advanced Research Projects Agency
• Mach 15-20 capable, regeneratively cooled scramjet DF R C NASA Dryden Flight Research Center
ha r d w a r e . DOF Degree Of Freedom
ER J Ejector ramjet combined cycle (RBCC) engine
Other related technology which could affect hypersonic ES J Ejector scramjet combined cycle (RBCC)
vehicle systems, designs and operation include: engine
• Hydrogen cooled, variable-geometry scramjet; FE M Finite Element Model
• Refined design systems (fully three-dimensional and Fn Th r u s t
real-time flight simulation); GR C NASA John Glenn Research Center
• Thermal management for combined/combination HC or H/CHydrocarbon (fuel)
en g i n e s ; HL D Hypersonic aircraft shaped low speed demon-
• Pulse detonation (airbreathing and rocket) engines strator vehicle
• Plasma physics such as WIG, energy extraction from HX L V Hyper-X Launch Vehicle
hypersonic flows, etc.; and HX P O Hyper-X Program Office
• Improved facility, instrumentation, and test techniques. HX R V Hyper-X Research Vehicle (X-43)
Hy F L I T E Hypersonic shaped RPV
These technology areas are required to push hypersonic HyTech AFRL hypersonic technology program
airbreathing vehicles beyond Mach 10. Work is under- H Hy d r o g e n
2
way in most of these areas, and a sound foundation IM I Insulated multi-wall Insulation TPS
remains from the X-30 program. However, these efforts Is p Fuel specific impulse, Lbf/Lbm
need to be directed toward a particular application, and La R C NASA Langley Research Center
eventually flight validation. LH 2 Liquid hydrogen
MS F C NASA George C. Marshall Space Flight Center
NASA National Aeronautics and Space
CONCLUSIONS AND RECOMMENDATIONS Ad m i n i s t r a t i o n
NA S P National Aero-Space Plane (X-30)
NASA, DOD and the aerospace industry have invested O Ox y g e n
2
40 years, and $3+B in hypersonic airbreathing vehicle RB C C Generic name for Rocket – Airbreathing
technology development. This includes refined design Combined cycle engine
methods, large wind tunnel and ground test databases, R Non-RBCC rocket
and (with the X-43) flight validation. Additional flight RJ Ramjet engine
tests are required to elevate the technology to the point RL V Reusable launch vehicle
that prototype vehicles can be considered. This paper RP V Remotely piloted vehicle
presents a logical, affordable approach to complete the RT B E Revolutionary Turbine Based Engine
development of this unique technology. Requirements SJ Supersonic combustion or dual-mode com-
for this technology have been identified. A method of bustion ramjet engine
screening hypersonic vehicle configurations proposed to SS T O Single stage to orbit
meet these requirements was presented. ST F Scramjet test facility
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American Institute of Aeronautics and Astronautics