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EXECUTIVE SUMMARY
The Committee on Hypersonic
Technology for Military Application of
the Air Force Studies Board was formed
to:
· Evaluate the potential military
applications of hypersonic aircraft.
Assess the status of technologies
critical to the feasibility of such
vehicles.
.
MILITARY APPLICATIONS
Early in our proceedings, we found
that firm military operational concepts
do not exist for applications of hyper-
sonic aircraft. Determination of oper-
ational requirements must await a better
understanding of critical technologies.
Thus, the focus of this report is an
evaluation of the status of these tech-
nologies. From this evaluation we
concluded that:
Hypersonic aircraft technology and
ramJet/scramjet propulsion offer
potentially large increases in speed,
altitude, and range of military
aircraft, and may enable or extend
important Air Force missions.
The simplest (and probably most
feasible) hypersonic vehicle would
cruise in the range below Mach
number 8.
The most attractive missions involve
flight to orbital or near-orbital
speed above the sensible atmosphere.
These missions offer flexible recall,
en r cute redirection, and return to
base.
Any potential military advantage in
the speed range between Mach num-
ber ~ and orbital velocity would be
negated by technical difficulties in
the areas of surface heating and
thrust, and weapons carriage, aim-
ing, and release.
The global or near-global range
1
EXECUTIVE SUMMARY
.
coupled with the cryogenic fuels of
hypersonic aircraft will require
unusual base support requirements
that must be considered in any
judgment of their operational utility.
NATIONAL AEROSPACE PLANE
PROGRAM
Part of our task is to advise the
Commander of AFSC on the:
· Research and development strategy
of the National Aerospace Plane
(NASP) research vehicle designed for
single-stage- to-orbit (SSTO) capa-
bility.
Research vehicle approach for the
aerospace plane that will maximize
the acquisition of knowledge in the
technical areas most critical for
hypersonic vehicles.
The aerospace plane program is in the
technology development phase, and
decisions to build an SSTO vehicle will
not be made before 1990. Due to the
technical uncertainties, we agree with
this approach. In particular, substantial
technology demonstrations are needed,
especially in high temperature materials
and structural concepts appropriate to
them, before a commitment to build an
SSTO research vehicle can be technically
justified.
We recommend that the NASP pro-
gram office retain the ultimate goal of
demonstrating the technical feasibility of
SSTO capability, but maintain an option
of selecting less than SSTO as part of a
prudent risk management strategy. Con-
sideration should be given to incorporat-
ing design features into this research
vehicle that allow such modifications as
may be found necessary from flight
experience.
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2
The research vehicle should also
have auxiliary rocket propulsion to
enable controlled flight with some inde-
pendence from the air-breathing propul-
sion system. The auxiliary propulsion
will enable greater exploration of the
vehicle's flight characteristics and will
help ensure flight safety. The drawback
is that this auxiliary propulsion system
will make the research vehicle heavier,
thus making it more difficult to achieve
the structural weight fraction required
for SSTO capability. Further, the
program office might design a series of
flight research vehicles, each one able
to access an increment of the total
flight corridor up to orbital conditions.
TECHNOLOGY ASSESSMENT
The following is a summary of the
status of the major technical areas
critical to hypersonic aircraft. Tech-
nical reviews such as this one tend to
focus on deficiencies and problems.
Because our purpose is to help the Air
Force support solid approaches and
correct deficiencies in its program, this
report will follow that pattern. The
committee wishes to be on record, how-
ever, as strongly favoring a vigorous,
vehicle-focused hypersonic technology
program, carried on at least at the level
of the planned aerospace plane program.
Propulsion-Airframe Integration
The success of any hypersonic
vehicle will depend as much, or perhaps
more, on the integration of the various
contributing technologies into a complete
system, as on the individual technol-
ogies. Such integration is quite mission-
specific. Integration of the propulsion
system and airframe is basic to deter-
mining the aerodynamic shape of a
hypersonic vehicle. How such integra-
tion is achieved becomes increasingly
important as the maximum air-breatliing
Mach number rises toward orbital values.
HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION
Efficient operation at very high Mach
number requires configurations that pose
serious integration problems at other
Mach numbers. Consequently, the low
speed propulsion system must be inte-
grated with the hypersonic propulsion
system and overall aerodynamics in a
way that does not degrade performance
prohibitively over the entire vehicle
speed range.
Finally, the amount of hydrogen
required for engine cooling exerts an
unusually high leverage on the airplane
size. Under the most severe conditions,
the molar flow rate of hydrogen coolant
within the airframe and engine structure
is more than double the molar flow rate
of air through the engine. Conse-
quently, the cooling system must be
effectively integrated with airframe and
engine system, and coolant plumbing and
pumping losses must be minimized.
To help solve these problems in the
NASP research vehicle program, we
recommend that engine-airframe integra-
tion receive emphasis in the design
definition by teams drawn from both
engine and airframe contractors.
Propulsion Systems
The key technology for hypersonic
air-breathing propulsion at higher flight
Mach numbers is the supersonic combus-
tion ramjet (scramjet). In this device,
injection of hydrogen fuel and its mixing
with air is the most influential factor
affecting engine performance. Current
techniques for modeling related phen-
omena are inadequate. Conventional
one-dimensional or quasi one-dimensional
computation of the reacting flow in the
combustor is insufficient and often mis-
leading. Further, the stability of the
scramjet flow with hydrogen reaction is
not understood. Flow instability poses
the possibility of developing strong
shock waves and catastrophic loss of the
engine.
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EXECUTIVE SUMMARY
For the scramjet to operate at peak
performance throughout its Mach number
range, extensive geometric changes must
be made while maintaining minimum
internal aerodynamic losses. These
changes place additional demands on the
design of coolant passages and on cool-
ant management. Also, transitions
between the various operating modes of
the propulsion system present severe
design and control problems if extreme
loads on the engine and airframe are to
be avoided. Since experimental verifica-
tion of most transitional modes cannot
be accomplished in ground-based facil-
ities, the complete engine will undergo
much of its high Mach number develop-
ment during the flight program. Thus,
it will be necessary to incorporate
auxiliary rocket propulsion in the
research vehicle with fuel supply and
controls separate from the scramjet, to
be able to vary the engine operating
point at fixed flight conditions.
Additionally, some rocket propulsion
must be incorporated in the propulsion
system of vehicles with orbital capabil-
ity, for final insertion into orbit and for
the de-orbit maneuver.
Aerodynamics
The aerodynamic problems of hyper-
sonic flight can be considered in two
categories. Below Mach number 10, the
problems are primarily of a fluid mech-
anical nature, where one must accurately
determine the pressure distribution, skin
friction, heat transfer, flow field details,
and mixing. Above Mach number 10, the
same fluid mechanics problems remain,
with the additional complication of the
rate kinetics of real gases, the special
low density phenomena of high altitude
flight, and the effect of small bluntness
on slender bodies. It is not yet possible
to simulate or compute with any degree
of certainty, these phenomena over the
entire flight range.
3
There is a critical need for aero-
dynamic data, both for input to compu-
tational fluid dynamics (CFD) studies,
for CFD validation and for engineering
design. At low hypersonic speeds no
current facility has a high enough Mach
number (S to 10) with an adequate Rey-
nolds number and size for the necessary
aerodynamic and configuration studies.
At high hypersonic speeds - above Mach
number 10 - it will be necessary to
simulate the aerodynamics (viscous
effects and flow field), while including
real gas effects at conditions ranging
from the continuum low altitude regime
to the free molecular regime at high
altitude. No current facility or facilities
can simulate the full range of flow
parameters with the needed flow quality.
Since complete simulation in ground-
based facilities cannot be carried out, it
will be necessary to check individual
elements of CFD codes. Thus, there will
be a need for flight tests for full valid-
ation of CFD codes particularly at high
Mach numbers.
To provide essential data for hyper-
sonic vehicle designs, we recommend:
· The immediate consideration of a
quiet tunnel in the Mach number 10
range and of a size and Reynolds
number capability to permit testing
of hypersonic configurations at close
to full-scale conditions. The Gas
Dynamics Facility at the Arnold
Engineering Development Center' for
example, might be so modified.
A combined CFD/experimental pro-
gram focused on key hypersonic
aerodynamic elements.
That a high priority be given to a
national integrated program to
determine reaction rates at high
temperatures as a needed input into
CFD computations.
That special attention be paid to the
problems of low density flows be-
cause of the special phenomena
associated with high altitude, high
Mach number flight. We also sug
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4
HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION
gest that the personnel of the NASA
Ames Research Center could provide
a key nucleus for such an effort.
Control, Guidance, and Information
Systems
Hypersonic vehicles present signif-
icant challenges in -control of the
vehicle trajectory and also in the con-
trol of vehicle configuration, aero-
dynamics, 'dynamics, ' propulsion, and
cooling with 'particular emphasis on the
need for their complete integration.
These demands in turn present difficult
challenges in the associated fields of
instrumentation and information systems.
The net result is that information and
control technology joins materials, pro-
pulsion, and aerodynamics as an enabling
technology.
Since the lower surfaces of the pro-
jected hypersonic aircraft are part of
the engines, angle of attack and sideslip
will significantly affect engine perfor-
mance. Precise control of the entire
integrated vehicle, perhaps as accurately
as to within 0.1 degree, is therefore
necessary. ' Further, exacting speed
control may be needed for engines
transition and aerodynamic heatir~g that,
in conjunction with the other demands
on the control system, may prove dif-
ficult to achieve. In the presence of
maneuvers and thrust asymmetries, speed
control presents a challenge in' controls
technology far beyond that posed by any
aircraft yet designed.' Offsetting this
very severe challenge is the potential
offered by advanced control' technology.
But to realize these potentially signif-
icant returns, it will be important to
systematically identify those areas where
such benefits are possible and incor-
porate them into the designs.
Both the performance and safety of
the vehicle will depend fully on onboard
processing. Projections for throughput
requirements for the onboard information
system range up to 25-50 million
instructions per second. This high level
of throughput will cause reliability
reuqirements to be in the range of 10
million to one billion hours between
failures. This level of reliability and
throughput has not yet been achieved in
a flight qualified installation. Other
agencies such as NASA and DARPA have
sponsored high reliability, high through-
put laboratory level programs that may
help assure the availability of an
adequate onboard processing system.
The thermal environment, even
inside a hypersonic vehicle, is likely to
be quite hostile to electronic, sensor,
and hydraulic systems. Active cooling
undoubtedly will be necessary. Advances
in high temperature electronics, cabling,
and hydraulics also will be highly desir-
able to offer more design flexibility and
reduce the weight requirements for
active cooling. In addition, the combin-
ation of speed, low density, and high
temperature impose conditions for
sensors beyond anything required to
date. The problem becomes even more
complex when the need for redundancy
is added. A program to identify and
develop the necessary sensors is
imperative.
A control system designed to accom-
modate all the uncertainties that can be
anticipated at this stage of the design
would have to be very-robust and com-
plex. Considerable savings in controls
expenditures are likely if a phased flight
development approach is adopted to
reduce uncertainty as the flight envelope
is expanded from subsonic to hypersonic
speeds. To make this possible the infor-
mation processing system installed in the
prototype aircraft must be exceptionally
reliable and be able to expand gracefully
to the necessary size and throughpu t
required for the final configuration,
without need for redesign or recon
~. .
~ ~gurat~on.
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EXECUTIVE SUMMARY
Materials, Structures, and Cooling
The system challenge posed by a
hypersonic vehicle to the materials and
structural community is reflected in the
structural weight fraction, which is the
weight of all the structure divided by
the take-off gross weight (TOGW). With
presently envisioned air-breathing
propulsion systems using cryogenic fuels,
a fuel fraction of approximately 0.75 is
required for SSTO performance. Design
studies show that the consequent struc-
tural weight fraction of an SSTO vehicle
must be approximately 0. 15. Present
military aircraft that operate mainly in
the slightly supersonic regime have
structural weight fractions of approx-
imately 0.30. An increase in either the
fuel fraction or the structural weight
fraction means either a multi-fold
increase in TOGW (if the payload is not
to decrease) or a vanishingly small
payload (if TOGW is not to increase).
Thus, from the viewpoint of structural
weight fraction, we consider the SSTO
objective very demanding. It can be
met only with new materials and new
structural concepts.
The structures will face severe
design conditions arising from aerody-
namic heating. The aircraft will require
new materials with elevated temperature
properties comparable to those of the
high strength aluminum alloys at room
temperature. A necessary consort to the
"strength" materials will be coatings to
prevent oxidation at the elevated tem-
peratures at which they will have to
operate. The aircraft will also require
materials for insulation and for both
passive and active cooling.
Experimental material cannot be
readily used in a flight test vehicle.
The scale-up of a material system that
has been proven in the laboratory to the
production quantities demanded by even
a research vehicle requires the develop-
ment of a materials fabrication technol-
ogy. Also, the properties of these new
5
materials are very process-dependent, so
that extensive processing experience
must be accumulated before they can be
reliably employed. Even after the
materials are available with nominally
adequate properties, the structural
designer must confront the effects of
the high temperatures, including:
stresses induced by the differential
expansion of the structure (the so-called
thermal stresses), degradation of the
material properties, and distortion of the
shape of the structure, which will lead
to interactions with the flight loads.
Designing for thermal stresses requires
temperatures to be predicted with great
precision. Also required are an accurate
knowledge of the thermal conductivity,
thermal coefficient of expansion, heat
capacity as a function of temperature,
and the thermal impedances of joints
and connections. Design considerations
are further complicated by the additional
stresses caused by loads due to maneu-
vers, gust, and landing.
In sum, vehicles with orbital or
near-orbital capability in a single stage
will require new materials and novel
structural concepts to meet the struc-
tural integrity requirements of strength,
rigidity, longevity, damage tolerance, and
efficiency; all of these at a cost com-
mensurate with the importance of the
mission. An aggressive and sustained
materials and structures program will be
required to develop these technologies to
meet the presently envisioned objectives
of the aerospace plane. Considerable
progress toward such a program has
been made in the last year, but we still
regard materials and structures as a
probable limiting factor in meeting the
objective of SSTO performance.
The Role of Computational Fluid
Dynamics
Computational fluid dynamics has
become a principal tool for aerodynamic
and propulsive-flow design of hypersonic
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6
HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION
vehicles. Computers can simulate simul-
taneously the hypersonic flight para-
meters of velocity, free stream density,
physical scales, and free stream thermo-
chemical state, while the present
ground-based test facilities cannot. The
role of CFD is especially significant for
high altitude hypersonic flight conditions
involving reaction-rate dependent air
chemistry.
Current supercomputers and numeri-
cal methods are able to simulate three
dimensional flows using the Reynolds-
averaged (time-averaged) Navier-Stokes
equations. However, the lack of capa-
bility for modeling of turbulent stresses,
heat flux, and transition locations, form
the principal current limitations of CFD
techniques. Although the validation of
three-dimensional hypersonic CFD codes
is now relatively limited, we expect this
situation to improve considerably in the
near future as various results from suet
codes and hypersonic experiments
become available.
The most intense local heating rates
on vehicles such as the projected NASP
research vehicle are expected to be on
cowl lips, caused by shock-on-shock
heating. Present CFD simulations of
this phenomenon agree well with experi-
ments for the lower portion of the
hypersonic flight corridor where shock
wave thickness is negligible. However,
neither computations nor experiments
have yet been made for these high
altitude hypersonic flight conditions.
The limitations of current CFD
simulations are not inherent to CFD
itself, but are due largely to the present
state of supercomputer development.
Advanced CFD technology will require
more powerful supercomputers than those
expected to be available in the near
future.
Our recommendations are that:
· an effort be made to significantly
improve the modeling of turbulent
.
mixing flows relevant to hypersonic
flight and of turbulent boundary
layer flows in the presence of heat
transfer and pressure gradient.
high-altitude shock-on-shock cowl
lip heating be investigated by direct
simulation Monte Carlo methods,
appropriate new experiments, and
continuum equations more advanced
than Navier-Stokes.
the direct numerical simulation of
turbulence in compressible flows
should be an integral part of the
long range technology development
program of the Air Force.
Test Requirements
Most U.S. hypersonic facilities were
built when the major national effort was
on ballistic missiles and space vehicles.
A careful examination is required of
their utility for the design and testing
of the new generation of hypersonic
vehicles. The nation must consider not
only current, but also long term needs
for hypersonic testing.
Since ten years or more are required
to design, build, and commission new
facilities, it must be accepted that only
currently available facilities will con-
tribute to the testing needs of the NASP
program during the technology develop-
ment and design phases. It is very
important, therefore, to ensure that all
existing facilities are used as effectively
as possible. We should note that many
of these are 20 to 40 years old and flow
conditions in them are not sufficiently
well known. Research must be done in
such facilities, to identify their flow
characteristics.
Components of hypersonic vehicles
will have to be tested at unpreced-
entedly high temperatures, for which
existing but inactive facilities such as
those at Plum Brook, Ohio, and the
NASA Dryden Test Facility should be
reactivated. Serious consideration should
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EXECUTIVE SUMMARY
also be given to the construction of
facilities for component testing at
temperatures above the 1200° C limit of
these existing facilities.
We do not know the local composi-
tion, temperature, and turbulence of the
upper atmosphere well enough to infer
high altitude hypersonic flight conditions
from ground test data. Measurements
and analyses should be made of these
characteristics of the atmosphere to
provide a firmer base for use of test
data in vehicle design and in flight test
planning.
It will not be possible to test
hypersonic vehicles with orbital or near-
orbital capability over the full range of
7
their flight conditions prior to flight.
Thus, hypersonic aircraft research above
Mach number ~ will require a philosophy
of flight research somewhat more akin
to that practiced for missiles and launch
vehicles. A clear recognition of this
will be essential to success of the
program.
The NASP flight research program
will collect more basic information on
aerodynamics, propulsion, control, and
structural behavior than any previous
flight research program. We recommend
that very careful consideration be given
to the flight test scenario during the
design of the research vehicle, not
afterward.
Representative terms from entire chapter:
hypersonic aircraft
