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Below are the first 10 and last 10 pages of uncorrected machine-read text (when available) of this chapter, followed by the top 30 algorithmically extracted key phrases from the chapter as a whole.
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OCR for page 75
STATUS OF HYPERSONIC TECHNOLOGIES
53
structural reasons, such schemes as MHD
acceleration of a hypersonic stream
generated by a wind tunnel, are being
investigated as a way to achieve high
Mach numbers. The committee has
reviewed one such proposal for steady
state crossed-field acceleration of air to
25,000 feet per second, and we find the
proposal does not reflect understanding
of the large body of knowledge devel-
oped by the ~~~ I,
last 20 years. The analysis on which
the proposal is based is limited to "one-
dimensional" or channel flow, without
consideration of wall effects. These
effects limit the feasibility of such
devices and have received a great deal
of study. The overall conclusion of
these detailed studies is that the steady
crossed-field accelerator is an ineffec-
tive device for producing large gas
velocities, because too large a fraction
of the input energy goes into heating
the walls, as well as the gas to be
accelerated.
MHD community Over the
2.~.3 Materials and Structures Test
Facilities
The Air Force recently conducted a
study of the high temperature test tech-
nology needed for hypersonic vehicle
applications. They found that heating
capability above 1400° C. will be dif-
ficult to achieve and that instrumen-
tation is not available for use above
800° C. Also, high temperature strain
gauges are not available for temper-
atures over 800° C. When testing must
combine flowing air, and mechanical and
thermal cycling to obtain the necessary
data for structural design, one must
conclude that additional testing capabil-
ities are needed to get the data in an
expeditious manner.
2.8.3.1 Structures Testing
Hypersonic vehicles will require
structures that are ultra-light, temper
ature resistant, inspectable, durable, and
safe. The structural design concepts
depend on vehicle configurations chosen
for the flight profiles that will be flown,
and from the material choices available.
Structural optimization can be done once
all the loads are known, the stresses
determined, heat transfer known, the
aeroelastic behavior of the vehicle
determined, and once the strength, stiff-
ness, and fracture toughness of the
materials selected are known in consid-
erable detail for all conditions of vehicle
operation.
The design criteria for a hypersonic
vehicle will determine the amount and
type of structural development testing
required. This should include the speci-
fication of time, temperature, load
synthesis, cumulative creep criteria,
oxidations criteria which can then
influence coating criteria, fracture
mechanics and fatigue, sonic fatigue and
panel flutter. A typical structural
component is shown in Figure 2-12.
2.X.3.2 Facilities
The Air Force study mentioned
above indicated that there are existing
facilities adaptable for major component
testing. However, a major structural
component test facility was estimated to
cost $90 million dollars. A full-scale
test facility that would be required for
structural certification was estimated to
cost $462 million dollars. In the
National Space Transportation and
Support Study 1995-2010, prepared by
the loins DOD/NASA Transportation
Technology Team, total structures/-
materials funding included facilities.
The facilities funding totaled $554
million dollars that included a structural
certification facility. In fact, these
figures add up to about the same amount
as estimated by the Air Force for
hypersonic vehicle structural testing and
certification.
OCR for page 76
~4
HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION
There has been some discussion of
activating the NASA Plumbrook high
temperature test facility in Sandusky,
Ohio, and the NASA Dryden facility at
Edwards Air Force Base, but to date
there seems to be no funding to activate
these facilities. In 1985, the Aerospace
Industries Associations Aerospace
Technical Council established a High
Temperature Test Facility (HTTF) Col-
laborative R&D Ad Hoc Group to deter-
mine whether the Members of A.I.A.
should form a partnership to develop
such an HTTF. In February of 1987, the
Ad Hoc Group, after visiting the NASP
program office, reported that DOD and
NASA were doing a "good job assessing
and developing the necessary test facil-
ities" that would cover most of the
needs identified by the HTTF ad hoc
group and that the group should be dis-
banded with no further action. As of
March 198S, there seems to be no posi-
tive action with funding to proceed to
define these facility needs. The major
test facility, now being refurbished to
carry out testing in aerothermal loads
and high temperature structures, is the
Langley S-Foot high temperature tunnel,
a Mach number 7 blowdown type of
facility in which methane is burned in
air under pressure and the resulting
combustion products are used as the test
medium with a maximum stagnation tem-
perature near 3800° R to reach the
required energy level of flight simula-
tion. This facility will, however, not be
ready for testing until the late fall of
1988. There is an urgent requirement
for the development of major high
temperature materials and structural
component test facilities. The develop-
ment of such facilities is mandatory to
insure that an adequate data base of
material properties is developed and that
structural design concepts can be
evaluated to support hypersonic vehicle
design.
In the end, flight experiments may
be necessary as an adjunct to evaluate
the structural concepts being considered
because of the inability to adequately
simulate the combined environments of
temperature and flow. Experimental data
can be obtained in several ways.
Rocket or free-flight tests lack
adequate communication between the
flying vehicle and the test engineer.
Only a small amount of data are
obtained from each flight, and each
flight represents a very large expendi-
ture of money and engineering time. To
get a maximum return from flight tests,
simulated flight tests first should be
performed in the laboratory.
2.X.3.3 High-Speed Wind Tunnel Tests
To obtain high temperatures in a
wind tunnel the air must be heated to
the desired temperatures, and this poses
new problems in wind tunnel design.
Such a venture entails its own host of
difficulties and results undoubtedly will
not be forthcoming for some time.
2.~.3.4 Laboratory Tests with Heating
Devices
Laboratory heating devices that
produce thermal energy can be devel-
oped. Such devices should be inves-
tigated, and the possibilities are many.
Radiant devices, such as those in
ordinary household cooking ranges or in
refractory ovens, can be arranged in a
dense pattern over a broad area to
obtain a distributed source of high
radiant energy. The transient heating
phase could be controlled by changing
the distance between model and heater.
It could also be done by a system of
shutters on the heater. Chordwise and
spanwise variation of temperatures could
be obtained by painting or otherwise
preparing the exposed surface of the
model to achieve different adsorptions.
The design of the heater could also be
arranged so that its thermal output
varies across its face. But the scarcity
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OCR for page 79
STATUS OF HYPERSONIC TECHNOLOGIES
55
. . .
Of experimental data and the complexity
of the problem would indicate that much
can be learned and perhaps should be
first learned from experiments on simple
models using simple experimental appara-
tus.
1. Test section 1 ft. in diameter or more.
2. In contrast, there is evidence that the Soviet Union continued to build hypersonic
facilities through the 1970s and 1980s.
3. Chord length defined as the square root of the test section area.
OCR for page 80
56
Below are the principal findings we
have adduced from our review of hyper-
sonic technology for military application,
and the recommendations we offer to
further these technologies and their
applications.
3.1
Potential Military Hypersonic
Applications
( 1 ~ Hypersonic aircraft technology, in
association with air-breathing
propulsion, offers potentially large
increases in speed, height, and
range of military aircraft, and may
enable or extend important Air
_ · .
force missions.
(2) Operational hypersonic aircraft will
necessarily have very large turning
radii, and useful missions therefore
require global or near-global range.
(3) Cryogenic fuels are necessary, and
any studies of hypersonic aircraft
missions should therefore include a
careful examination of the base
support requirements which they
imply.
(4) The simplest class of hypersonic
cruise vehicle would fly up to Mach
number 8. This class can signif-
icantly advance the reconnaissance
and strike missions now done by
the SR-71.
(5) The most attractive potential Air
Force missions involve flight to
orbital or near-orbital speeds above
the sensible atmosphere. In con-
trast to ballistic missiles and
satellites, these offer flexible
recall, en route redirection, and
return to base.
HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION
3.0 FINDINGS AND RECOMMENDATIONS
(6) Sustained hypersonic flight in the
atmosphere between the two ex-
tremes of (4) and (5) above pre-
sents major technical difficulties.
Problems of surface heating, thrust,
vehicle stability and control,
infrared signature, aiming, and
weapon release could make any
potential military advantage in this
speed range unlikely.
3.2 Propulsion-Airframe Integration
Engine-airframe integration is a
key aspect of configuration def-
inition for hypersonic vehicles-
increasingly so as the maximum air-
breathing Mach number increases.
The combination of long forebody
and low Reynolds number produce a
thick entropy layer that must be
ingested by the engine or diverted.
Its thickness is sensitive to Mach
number and Reynolds number, and
will vary significantly over the
flight corridor.
The Low Reynolds number is dic-
tated at high Mach number by the
need to reduce heat transfer rates
and pressure loadings, to transition
to rocket propulsion for orbital
insertion, or both.
4) A very large ratio of capture area
to frontal area results from low
Reynolds number (high altitude) and
small fractional energy addition due
to combustion.
5) Efficient operation at very high
Mach numbers require configur-
ations that pose serious integration
problems at off-design Mach num
OCR for page 81
FINDINGS AND RECOMMENDATIONS
her. The large nozzle expansion
leads to very large base drag at
transonic speeds. The interaction
of the nozzle expansion plume with
the slipstream, and with the
reaction control system, will
influence both the net thrust and
moments at near orbital speeds.
6) Integration of the low speed pro-
pulsion system with the hypersonic
propulsion system, in a way that
does not degrade the performance
at hypersonic speeds, is a major
concern.
7) The variation of engine-inlet
boundary layer conditions with
flight conditions (Mach number,
Reynolds number, and altitude)
must be quantitatively predictable,
or an engine concept must be
devised that is insensitive to the
boundary layer thickness.
8) Items 5, 6 and 7 above are un-
solved problems. Engine-airframe
integration should receive more
emphasis, by teams drawn from
both engine and airframe contrac-
tors.
3.3 Propulsion Systems
I ~ Injection of hydrogen fuel and
rapid mixing with air with minimum
loss is the most influential factor
affecting the engine length and
heat load.
(2) The heat release pattern in the
engine is determined by the rate of
molecular mixing, and the super-
sonic flow in the engine is
extremely sensitive to the heat
release pattern.
The ingestion of ramp boundary
layer and bow shock layer by the
engine poses difficult problems of
engine design and penalizes engine
57
performance.
(4) The stability of the scramjet flow
with hydrogen reaction is not
understood; instability poses the
possibility of developing strong
shock waves and catastrophic loss
of engine.
Short term design studies and long
term research studies of hydrogen
injection and mixing should be
increased as soon as possible to
assure that this issue does not
become an obstacle to high-speed
engine development. Measurement
of molecular mixing should be
emphasized and the exploration of
novel techniques of mixing augmen-
tation must be encouraged.
Because the combustor heat release
pattern is mixing controlled and,
further, because the state of the
air entering the combustor may be
extremely non-uniform, the mixing
process must be understood to the
extent that it can be controlled" as
well as accelerated.
(6) One-dimensional or quasi one-
dimensional computation of reacting
flow in the combustor is inadequate
and often misleading.
(7) The H-OH reaction must be com-
pleted for the scramjet to perform
well. Much of this reaction will
happen during expansion in the
nozzle and this reaction may
"freeze out" early in the expansion
process.
(8) Under the most severe conditions
of operation the molar flow rate of
hydrogen in the cooling passages is
more than double the total molar
flow rate of air through the
engines. Effective use of coolant
and minimization of pumping losses
is imperative and an unusual degree
of integration with structural
design is required.
OCR for page 82
58
HYPERSONIC TECHNOLOGY FOR MILITARY APPLICATION
(9) The hydrogen requirement to cool
the engine exerts an unusually high
leverage on the airplane size and
weight. It is essential to refine
the accuracy of and confidence in
estimates of cooling requirements
before final selection of airplane
size.
( 10) High priority and additional
emphasis must be given to the
research and design studies con-
cerned with the utilization and
management of hydrogen coolant
flow. This is of particular
importance in the portions of the
engine that experience geometric
changes during the acceleration.
(1 1)
Film cooling and sweat cooling with
hydrogen have very attractive fea-
tures and both technological and
research efforts must be augmented.
The gas dynamic peculiarities of
using hydrogen as the coolant
should be emphasized in these
studies. This work must be accel-
erated because coolant requirements
have such a powerful impact upon
the airframe design.
(12) The scramjet must operate at peak
performance throughout its entire
Mach number range during accel-
eration. The configuration and
geometric changes required over
this range are very extensive and
must be done with the minimum
introduction of shocks and other
losses.
(13) The geometric changes required of
the scramjet over its Mach number
range place demands upon design of
cooling passages, coolant flow
management and seals that are of
unprecedented difficulty.
(14) Transition between the three oper-
ating modes of the propulsion
system, subsonic to ramjet, ramjet
to scramjet, and the re-start and
reverse transitions upon re-entry,
present extremely sensitive and
difficult problems. These must be
solved to avoid placing unaccept-
able structural and thermal loads
on the airframe and engine, which
may lead to failure.
( 15) The transition from one engine
mode to another, especially from
the ramjet to scramjet, might
produce large unsteady loads and
unsatisfactory starting. To insure
against these problems, sufficient
rocket propulsion should be incor-
porated into the powerplant com-
plex to suppress any severe
problems during transition.
( 16) Some rocket propulsion must be
incorporated into the final propul-
sion system a) to reach orbit from
the scramjet Mach number limit, b)
to facilitate the gradual introduc-
tion of advanced scramjet technol-
ogy over the life of the airplane,
and c) for de-orbit maneuver.
(17) The high-speed engine development
should be predicated on the prob-
ability that most of the develop-
ment will be done in flight test.
To develop an engine in flight, an
auxiliary rocket propulsion system,
separate from the NASP engine
package, will be needed to augment
thrust and assure airplane trim
during high-speed engine tests.
( 18) Complete scramjet engines will
undergo development and testing
during the flight program, not in
ground-based facilities. Con-
sequently, it is necessary to
incorporate some rocket propulsion
- separate from any that may be
integrated with the scramjet - for
setting desired engine test condi-
tions and extending the flight
envelope of the airplane.
OCR for page 83
FINDINGS AND RECOMMENDATIONS
(19) With boundary layer ingestion, the
scramjet engine is quite sensitive
to angle of attack. This variation
in engine operation will be
reflected in changes of pressure
distribution over the discharge
nozzle, resulting in a large pitching
moment. These difficulties may be
avoided only through unusually
careful integration during airframe
and control system design and
development.
(20) The modular design of the scramjet
engine allows interaction between
the inlets of adjacent modules
during ramjet start-up and transi-
tion from ramjet to scramjet oper-
ation. This interaction can
propagate inlet malfunctions from
one module to adjacent modules.
This behavior appears likely to be
particularly sensitive to yawing
motions of the airplane.
(21) The strong interaction between the
high speed engine, the forebody
ramp, and the external nozzle, and
the very powerful dependence of
this interaction upon pitch and yaw
of the entire airplane has important
implications upon the control
system and the coupling of engine
and airplane. Consideration should
be given to design compromises
that would reduce this potential
problem, even at the expense of
reduced performance of the earliest
NASP configurations.
3.4 Aerodynamics
This committee has identified two
main aerodynamic problem areas. At low
hypersonic Mach numbers (below about
10), the problem is mainly one of fluid
mechanics. The prediction of the boun-
dary layer and flow field characteristics
are required to permit the detailed
determination of the pressure distribu-
tion, skin friction, heat transfer, and
59
the flow field condition around the body
and through the inlet to the combustion
chamber. Above Mach number 10, the
aerodynamic problems involve the factors
identified in the lower Mach range, with
the additional complication of the rate
kinetics of real gas effects and the
special problems of low density flows
and small bluntness dimensions. Neither
the low nor high Mach number areas are
currently amenable to detailed wind
tunnel exploration or validated compu-
tation to provide a well-grounded base
for design, although some results are
available from facilities that partially
simulate the real flows. Therefore,
progress must rely on a fragmented
approach, where limited experiments and
computation will in time provide an
adequate base for design. Validation of
this base will require flight tests that
include many elements simultaneously, a
situation not amenable to full simulation
on the ground or by validated compu-
tation.
3.4.1 Low Hypersonic Speeds (Mach
Numbers 6 to 10)
The prime requirements are to
identify the location and details of
transition initiation, the transition
region, the mixed flowfield and
boundary layer characteristics
around and through complex
geometries with cold walls, and the
mixing phenomena in the combus-
tion chamber. These problems are
recognized and are within reach of
current technology but have not
been solved.
2) A unique Mach number 3.5 "quiet"
research facility at NASA Langley
is beginning to provide results
indicating the necessity for such
flow characteristics. A Mach
number 6 research facility has been
approved, but not built. These
research facilities are inadequate
for the requirements of Mach
Representative terms from entire chapter:
test facility
