AAWSAP DIRD Space Access Where Weve Been and Where We Could Go March 8 2010
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8 March 2010
! COD: 1 December 2009
DIA-08-1001-006
Space Access: Where We've
Been ... and Where We Could
Go
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Space Access: Where We've Been . . . and Where We Could
Go
Prepared by:
Acquisition Support Division (DW0-3)
Defense Warning Office
D irectorate for Analysis
Defense Intelligence Agency
Author:
AAP Person 67
Administrative Note
COPYRIGHT WARNING : Further disseminat ion of the photographs in t his publicat ion is not aut horized.
This product is one in a series of advanced technology reports produced in FY 2009
under the Defense Intelligence Agency, Defense Warning Office's Advanced Aerospace
Weapon System Applications (AAWSA) Program. Comments or questions pertaining to
this document should be addressed tolAAP Person 1 I, AAWSA Program
Manager, Defense Intelligence Agency, ATTN: CLAR/DW0- 3, Bldg 6000, Washington,
DC 20340-5100 .
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Contents
Introduction ............................................................................................................v
Propulsion Perspective ........................................................................................... 1
Hypersonic Configuration Concepts ........................................................................ 2
Thermodynamics and Materials ............................................................................ 16
The Qu Tube ......................................................................................................... 23
Rocket Propulsion ................................................................................................ 25
Up-and-Down Operations ..................................................................................... 30
Launch Options .................................................................................................... 33
Atmospheric Variations ........................................................................................ 35
Conclusion ............................................................................................................ 37
Appendix A: Historical Perspective ....................................................................... 43
Appendix B: Aeropropulsion Integrated Vehicle ................................................... 46
Appendix C: TAV Operational Costs ............. ......................................................... 47
Appendix D: Landing Ellipses for Hypersonic Gliders ........ .. .................................. 48
Figures
Figure 1. Hardware Flow ........... .. ...........................................................................vi
Figure 2. Impact of Air-Breathing Rocket ................................................................1
Figure 3. HSVS Hypersonic Cruise Aircraft Showing True Skin Temperature ........... 3
Figure 4. Hypersonic Rocket- Powered Glider Hypersonic Air- Breathing Cruiser ......4
Figure 5. Delta-Lifting Body Designs .......................................................................5
Figure 6. Martin Marietta X-24 A & B Research Gliders ............................................7
Figure 7. Detailed Design Analyses Show the Weight Trends are as Much a
Function of Configuration Family as Lift-to-Drag Ratio ............................9
Figure 8. High-Performance Hypersonic Glide Aircraft .......................................... 10
Figure 9. NASA Langley Wing-Body Configuration WB-004 with Critical Areas for
Wing Bodies Identified ...........................................................................11
Figure 10. FDL-7C/D and FDL-7MC Lifting-Body Configuration .............................. 12
Figure 11. FDL-7C/D with a DuPont Retractable Inward-Turning Inlet ................. 12
Figure 12. Comparison of FDL-7C/D and Model 176 .............................................. 13
Figure 13. Sufficient Cross Range (L/D) Means There is No Waiting to Return ..... 14
Figure 14. Hypersonic Glider Characteristics .........................................................15
Figure 15. Both Delta Planform Lifting Body (Dynasoar) and Model 176 Offer
Superior Landing Performance .............................................................16
Figure 16. McDonnell Aircraft Company Roll-Bonded Titanium Structure .............. 17
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Figure 17. Model 176 in the McDonnell Douglas Hypervelocity Impulse Tunnel .... 18
Figure 18. FDL-7C/D, Model 176 Entry Temperature Distribution ......................... 19
Figure 19. Even At Mach 12, Embedded Vortices in the Boundary Layer Alter the
Local Heat Transfer ..............................................................................19
Figure 20. Thermographic Phosphor Image of Model 176 at Near-Maximum
Angle of Attack .....................................................................................20
Figure 21. From L/D Maximum to Maximum Angle of Attack, There Is Always a
Cool Sub-layer Adjacent to the Wall .....................................................20
Figure 22. This 1988 SEP Bordeaux SiC/SiC Panel Could Sustain Temperatures of
up to 3,000°F ........................................................................................21
Figure 23. UBE Corporation's Tyranno Cloth ..........................................................21
Figure 24. A Porous Nickel Tip Oozing Water ........................................................21
Figure 25. FDL ASSET Flight-Tested From Orbital Speeds to Evaluate 1960s
Materials ..............................................................................................22
Figure 26. Heat Pipe Shuttle Leading Edge Designed and Built by McDonnell
Douglas Astronautics ...........................................................................22
Figure 27. Boost-Glide Strategic Vehicle with Pratt & Whitney XLR-129 Rocket
Engine Installed ...................................................................................25
Figure 28. XLR-129 ...............................................................................................25
Figure 29. Two Rocket Air-Breathing Rocket Cycles to Mach 5.5 ........................... 27
Figure 30. HOTOL Evolution: From Aerodynamic Optimum Configuration to
Practical Launcher Configuration .........................................................27
Figure 31. LACE Air-Breathing Rocket ...................................................................29
Figure 32. The FDL-7 Class of Vehicles ..................................................................29
Figure 33. Takeoff and Landing Speeds of Minimum-Sized Launchers ...................30
Figure 34. Horizontal launch Not Practical Unless Weight Ratio Less Than Four ... 31
Figure 35. Propellant Tanks That Are Not Reentry Vehicles Greatly Reduce
System Weight .....................................................................................33
Figure 36. Simple Horizontal Integration and Vertical Launch Provides Rapid
Launch Capability .................................................................................34
Figure 37. A Vertical Launch Complex Provides Vertical Toss Back Booster
Recovery and Horizontal Landing Facilities for the Hypersonic Gliders 34
Figure 38. A 1964 MDC Astronautics, St. Louis, Briefing ........................................35
Figure 39. Earth's Atmosphere ..............................................................................36
Figure 40. FDL-5 Scale Model of A Stage and One-Half..........................................37
Figure 41. The FDL-7 and Model 176 Class of Hypersonic Gliders ..........................37
Figure 42. Hypersonic Decelerating .......................................................................38
Figure 43. Where We Are Today ............................................................................39
Figure 44. Where We Could Be If We Can Recapture the Engineering Confidence
and Expertise of the Apollo/Saturn V Era .............................................40
Figure 45. TAV Operational Costs ..........................................................................47
Figure 46. Landing Ellipse .....................................................................................48
Tables
Table 1. Characteristics of Selected Flight Dynamics Laboratory Hypersonic
Glider Configurations During the 1958-68 Timeframe ...............................6
Table 2. Elements of the Space Infrastructure Shown in Figure 44 .......................41
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Space Access: Where We've Been . . . and Where We Could
Go
Introduction
Development of commercial access to space by our budding space
faring civilization is a straightforward effort dominated by
propulsion and reliability. The initial focus should be on schedulable,
dependable access to and from low Earth orbit (LEO). For years we
have known the means to accomplish such a task but have lacked a
dedicated organized effort. The key requirement is to develop a
robust and not necessarily a low-cost infrastructure, without which
commercial exploitation of LEO and the moon will not be possible.
This is a matter of skill; operational hardware based on durable,
reliable, and demonstrated components; and operational systems. It
is not necessarily a matter of technology. However, technology
discovery and development are necessary for future space travel
beyond Earth's environs. This paper addresses these issues by
providing a running account of the historical details associated with
the development of the myriad systems proposed and tested to
provide access to space.
Among the many advances in space access that will be possible in
the future,1 the key technology developments will be in the area of
propulsion, because without these we are confined to our solar
system by flight times limited to a project team's functional life. The
Pioneer spacecraft were fortunate to be monitored for 20 years.
However, the issue facing our spaceflight organizations is the lack of
a durable, consistent, schedulable, and frequent hardware system to
and from space assets such as the International Space Station.
In October 1958, the author's job in the vertical wind tunnel at
Wright-Patterson Air Force Base abruptly changed; hypersonic and
high-temperature flows became a new focus. What was then the
Aircraft Laboratory was to become the Air Force Flight Dynamics
Laboratory (AFFDL), with a focus on space flight. Al Draper of the
AFFDL began working with a select group of aerospace firms on
hypersonic gliders. The initial requirement from the Air Force was to
quickly find operational access to space. Technology application,
hardware design and fabrication with an innovative application, and
extending the industrial capabilities of the time were very much the
issue, as exemplified by the Lockheed A-12/SR-71. When asked
about space access at the time, a group of Aerospace Corporation
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veterans responded, "It was not a technology issue; it was a
hardware issue."
In a keynote address to the Aeronautical Revolutionary Concepts
Workshop sponsored by the Vehicle Applications Panel of the
National Research Council and held at NASA Ames in July 1984,
then-Assistant Secretary of Commerce for Productivity, Technology,
and Innovation Dr. D. Bruce Merrifield identified the problem of
translating ideas into products as preparing technology for product
manufacture. Dr. Merrifield drew an analogy between this step and
Major League Baseball's farm system, which prepares skilled but
untrained players for the major leagues. The United States assigns
projects to accomplish technology tasks so the flow of production
ready hardware is always improving and is not fixed (see Figure 1).
Innovation Focus:
Preparation of Technology for Application
---r- - - - - 7 t o 1 0 y e a r s - - - - - ~ 1
- - - - - - ~ - •"-
-""••"_,
_ ,__
...;;.;.:,
.....:..,----,....-=----_-_-_ , - - -...,..
_.. --"'"'--
".........., .............
.... ·····•.. ,••· ' •, ..
__ __
Prototype
Idea PrOCluct and Pilot lntenm + Commercialization ·.
Generation___., .__
._ __, Develooment Plant I Manufacture
lnve~n Translation ··•....
··•.
••······· ·····•• ..................... ...•• •· • .................. ................• 1:i:t· f A••
ROLP
100/o Cost
Gap in effective 900/o Cost No Gap in
preparation for transformation
application to production
in United States
Or.O. Bruce Merrifield
Oeptment of Commerce
1983 Innovation In Aeronautics Assistant Secretary for Productiv ity,
Wortcshop, NASA AMES Technology and Innovation
Figure 1. Hardware Flow
Saturn I and Saturn V could be readied for a moon flight in such a
short time because most of their hardware was based on a frozen
design, proven production processes, and adaptation of existing
hardware. Using a similar approach, current industrial capabilities
can create the next practical system for accessing space. In the late
1950s and early 1960s, the U.S. Air Force was working toward an
operational capability analogous to its B-52 fleets: flight operations
when required or "on demand." After NASA was assigned
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responsibility for space access, that Air Force's focus switched to
surveillance, communication, and Global Positioning System
satellites.
In the late 1950s, there existed a predisposition-forced by the
military competition between the United States and the former
Soviet Union-to use rockets derived from military ballistic missiles.
That decision curtailed efforts to develop alternatives to chemical
rockets together with practical commercial developments. With the
orbiting of Sputnik, the aircraft path to space, as represented by the
X series of planes, ended with the X-15. With the X-15's demise, all
efforts to fly aircraft to space ended, replaced by the more familiar
(but less practical) strategy of loudly blasting to space with
expendable rockets derived from undertested ballistic missile
hardware, as documented in early failures.
Like their ballistic missile progenitors, current expendable rockets
can be launched only once. With the exception of the experimental
Delta Clipper developed and operated by William Gaubatz and the
late Pete Conrad, no operational launcher has ever successfully
aborted. In this context, a reusable launcher is simply an
expendable with some parts reused a few times. Thus, neither the
United States nor the Soviet Union/Russia has ever realized a truly
commercial approach to space travel, although the Soviets came
close to taking the first step with the since-terminated
Energia/Buran system. Both the United States and the Soviet
Union/Russia historically have generated a large number of concepts
that could fly directly to space and return on a sustained, frequent,
scheduled basis. An all-up air breather such as the NASP was to
solve that problem and fly directly to space and return. Developing
an operational mach 12 to 14 aircraft with air-breathing propulsion
presents a serious design, engineering, and fabrication challenge
analogous to the SR-71 Blackbird.
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Propulsion Perspective
In exiting Earth's atmosphere, the propulsion system and configuration are inexorably
linked. A hypersonic glider exits the atmosphere on either a rocket booster or a first
stage of a two-stage-to-orbit aircraft. As such, it usually exits the atmosphere quickly,
and the key exit design considerations are the high transonic aerodynamic and the
mechanical loads encountered in the exit trajectory. Whether for a new rocket launcher
or the U.S. space shuttle, the phenomenon is the same: the peak mechanical loads
occur during exit. In this case, the exit aerodynamics are important but not vital. The
vital aerodynam ics and thermodynamics (aerothermodynamics) are in the entry glide,
where thermal loads are maximal and must be controlled. The vehicle must always be
controlled in flight so its attitude and direction are within limits set by the
aerothermodynamics. The angle-of-attack limits are very close for high-performance
hypersonic gliders, as their glide angle of attack is 11 to 15 degrees, not the 45 degrees
of the space shuttle. Even the Russian Buran had a lower glide angle of attack than the
shuttle; a TsAGI report given to the author by Vladimir Neyland shows it to have been
about 30 to 35 degrees. 2 Like the Buran, the high-performance glider is best controlled
by an automatic integrated fl ight control system that monitors the thermodynamic state
of the vehicle, as well as its aerodynamic and trajectory states. The sensor array
provides real-time information to the control system that can maintain the correct
attitude in a manner a human controller could not accompl ish. So it is this phase of the
flight that designs the hypersonic glider.
The exception is when powered by an All
Rocket
air-breathing rocket (HOTOL, Skylon, LAC E
and LACE), which must remain lower in Rooket
the atmosphere until reaching the air Aero Spike
breathing rocket transition to Konst anbn FeotkIskov
conventional rocket. The configuration
for the air-breathing rocket is different,
as it must have a retractable air inlet in
the mach O to 5 range but does not
determine the vehicle configuration. ..._Under expanded
·,
The impact is significant, as the carried Over expand'e<I
oxidizer is reduced in the heaviest initial
portion of the flight, as shown in Figure 5. 35 ton payload
2 for a Delta Clipper-type design with 334 ton 101 ton TOGW
an aerospike nozzle tested by 28.6 ton 15.9 ton DEW
678 m2 428 m2
Konstantin Feotkiskov. The example is
from a Senior Capstone Design Study Figure 2. Impact of Air-Breathing Rocket
Team from Parks College, Saint Lou is
University, circa 1992, and is based on the engineering reports the author was
permitted to read from the library of Konstantin Feotkiskov, an aerospace designer and
cosmonaut. The question, as always, is, why bother with air-breathing systems at all if
they are that much of a challenge? The answer is to consider a partial air-breathing
system based on available hydrogen/oxygen rockets that operate to about mach S.S . It
operates in a flight region where the carried oxidizer quantities are the greatest. An
operational system is sought that is capable of a large number of fl ights per year. The
fewer resources required for launch, the greater ease with which the system can
operate and the greater potential to operate from more bases .
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The Russian design bureaus are to thank for arriving at a concept that eliminated the
noisy and hazardous air-breather takeoff and for increasing the operational flexibility of
the British HOTOL concept. Glebe Lozino-Lozinski had a concept for a spacecraft with a
7-metric-ton payload carried atop an Antonov An-225, with a second An-225 carrying
the liquid hydrogen and launch facilities and staff.3 The An-225 was in fact a mobile
launch facility; it could literally launch a satellite for any facility that could
accommodate a B-747 or an MDC-11. With Rolls Royce or General Electric engines, the
An-225 becomes a more easily maintained vehicle with better altitude performance.
The An-225's empennage is modified from the An-124's single vertical and horizontal
empennage to an 'H' configuration. This permits the powered hypersonic glider to
easily lift off the top of the vehicle, as the MBB Sanger wind tunnel test demonstrated.
Most commercial transport aircraft larger than ERJ 170 are potential mobile launch
platforms for space tourism, point-to-point cargo, or orbital facilities support. Most of
the commercial passenger equipment can be removed, with just enough equipment
remaining for a launch crew. The fuselage is strengthened and fitted with external
mountings for the hypersonic glider. The landing gear need not be modified, as the
same maximum weight as the commercial transport will be ma intained. The flight
control system would be adapted to automatically maintain the correct launch
trajectory until separation. A second modified transport would be modified to carry the
liquid hydrogen and liquid air to fuel the hypersonic vehicle, along with maintenance
and support crew. The intent is to use the automatic launch checkout the author
witnessed at Baikanour in 1988, wherein a Soyuz that arrived on its train carrier at
0500 hours launched carrying a Progress capsule at 1715 hours the same day. That
should make a local launch possible with in hours of arriving at the specified airport
launch departure site. These two elements can provide a commercial space launch
facility that requires no special or dedicated operational base.
Hypersonic Configuration Concepts
The configuration and the propulsion system are linked through aerothermopropulsion
integration. This approach is not new, as a wide spectrum of configurations and
concepts existed in the 1960s. One such McDonnell Aircraft Company concept is shown
in Figure 3. This potential operational mach 12 cruise vehicle was developed for the
U.S. government as a strike reconnaissance vehicle taking off from a U.S. Air Force
base. The concept was to provide on-demand reconnaissance in force operations.
However, as was the case with all such efforts in the 1960s, none of the aircraft derived
from the "flight-to-space" efforts reached a hardware stage. Individuals working on
these projects were convinced that the industrial capability existed to design and
fabricate these vehicles, and that such vehicles were technically feasible. The concepts
varied widely among different nations, but all had as their goal a transportation system
to space that had commercial potential. This discussion is provided to discriminate
between rocket-powered hypersonic gliders and hypersonic cruisers with an air
breathing propulsion system .
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--
McDonnell Aircraft
Advanced Design Dept.
1958 to 1967
Mr. H. D.Altis , Director
Figure 3. HSVS Hypersonic Cruise Aircraft Showing True Skin Temperature
A wide variety of configurations for recoverable spacecraft are possible. But if the
requirements for a transportation system capable of traveling to and returning from
space are to be met, the configurations spectrum is significantly narrowed. Two basic
configuration types emerge. One configuration is for a hypersonic glider powered by
either rocket or air-breathing rocket cycle propulsion that can operate as air-breath ing
propulsion to mach 5.5 or less. A versatile variable-capture, inward-turning inlet4 can
be integrated with the vehicle configuration derived from the FDL series of hypersonic
gliders developed by the U.S . Air Force Flight Dynamics Laboratory (AFFDL) 5 and the
work of the McDonnell Douglas Astronautics Company. Because of the mass ratio to
orbit, these configurations are vertical takeoff and horizontal landing vehicles,
exemplified by the upper-left vehicle in Figure 4. This vehicle is usually an upper stage
in a two-stage-to-orbit rather than a single-stage-to-orbit vehicle.
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Rocket Powered
Hypersonic Glider
Combined Cycle
Airbreather Powered
Hypersonic Cruiser
Figure 4. Hypersonic Rocket-Powered Glider and Hypersonic Air-Breathing Cruiser
The second configuration is for air-breathing propulsion systems operating at between
mach 6 and mach 14 that require a propulsion-configured vehicle, where the underside
of the vehicle is an integral part of the propulsion system (forming most of the air
capturing inlet). This is typified by the lower-right vehicle in Figure 4. The thermally
integrated, air-breathing, combined-cycle configuration concept is derived from the
McDonnell Douglas (St. Louis) Advanced Design organization. The vehicle concept
initially conceived in the late 1950s and early 1960s was an air-breathing propulsion
configured vehicle accelerated by a main rocket in the aft end of the body, as shown in
Figure 3. The vehicle's underside is the propulsion system; the engine is in the engine
module.
Both basic shapes are functions of tau-that is, for a given planform area, the cross
sectional distribution is determined by the volume required. Tau was reported in
D. Kuchemann's book on supersonic aerodynamics 6 as:
'C =
Vtotal
splan
1.5 (1)
The only configuration discussed in the book in any detail is the rocket-powered
hypersonic glider. The hypersonic glider has greater near-term potential to become an
operational system, considering the failure of the National Aerospace Plane (NASP) to
reach a functional hardware stage.
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Whatever goes into orbit must enter the atmosphere many t imes if it is to be a
sustained-use vehicle . If it is to be a commercial vehicle, then the flex ibility to land
wherever the commercial customers are is essential. Consider how successful FedEx,
UPS, or DHL would be if there were only two pickup and delivery sites in the United
States and a few more elsewhere in the world. A ballistic capsule has even fewer
landing options, and a saltwater landing and recovery is too costly to be commercially
feasible. What is needed is a hypersonic glider with the flexibility to enter when
necessary, without waiting, and to land at different operational bases, just as a
transport might. There were three serious competitors in the United States with
respect to hypersonic glider configurations: the AFFDL at Wright-Patterson Air Force
Base, the McDonnell Douglas Corporation (MDC), and the Lockheed Corporation.
NASA Ames and NASA Langley were also generating hypersonic configurations, but
NASA's views on hypersonic gliders (fundamentally research and development projects)
and their glide range requirements differed from those of the three organizations listed
above. That difference is clearly exemplified by the difference between the operational
requirements of an experimental aircraft (such as X-1, X-2, X-1 0, X-15, or X-20) that
flies infrequently and at the convenience of the research organization and those of an
operational Air Force or Navy aircraft that must be able to fly on any day in almost any
weather when needed (also a Russian spacecraft operational rule). From the middle of
the 1960s to the early 1970s, the U.S . Air Force and NASA had disagreements over the
operational capability of these aircraft and their requirements. As a result, each went
its own development direction, and much of the originality and practicality of the AFFDL
concepts has not been reflected in the space access configurations developed by NASA.
There was a final attempt to apply the AFFDL's philosophy of a high lift-to-drag (L/D)
ratio delta planform configuration to the NASA space shuttle, as detailed in the article
"A Delta Shuttle Orbiter" in the January 1971 issue of Astronautics and Aeronautics. 7
Figure 5 shows the array of delta planform configurations the AFFDL considered during
the 1958-68 timeframe.
Figures. Delta-Lifting Body Designs. Array of delta -lifting body designs shows configuration is not limited to
high hypersonic lift-to-drag ratios, high cross range, and large size. 8
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The AFFDL's approach was to design a hypersonic performance configuration that would
minimize the waiting t ime in orbit to return to the continental United States (CONUS).
This resulted in configurations with sharper leading edges and smaller nose radii than
found in NASA and Russian configurations. All of the material, structural, and
thermodynamic details related to the sharper configurations were tested and verified in
ground test facilities and flight tests (BGRV and ASSET). Characteristics of selected
AFFDL hypersonic glider configurations are identified in Table 1.
Table 1. Characteristics of Selected Flight Dynamics Laboratory Hypersonic
Glider Configurations During the 1958-68 Timeframe
# Model Observation
Flat bottom, sharp leading edges, conventional tails, as
l FDL-24B
designed
All Body Glider, similar
2 Upnirned spatular nose, conventional tail s
to Russian BOR vehicles
Test vehicle to evaluate aerodynamics,
3 ASSET thermodynamics and materi als, based on nose of
DynaSoar
Flat bottom , sharp leading edges, variable geometry
4 FDL-7MC
wing, ex perimentally developed tail X configuration
5 Blunt nose, wing-body DynaSoar type configuration
Spaturlar Nose Version First integration of 2-dirnensional nose (less drag) on a
6
of DynaSoar type hypersonic glider (R.D. Newmann)
7 FDL-8 Flat bottom, sharp leading edges, outboard tails
ASA Ames flat up-swept with bottom, ro und upper
8 HL-10
body, high dihedral angle tai ls
9 X-24A NASA Langley round body, high dihedral angle tails
LO Star Body based on Russian Star Body type configuration
Configuration 2 was a higher wing-loading, relatively blunt all-body with an upswept
spatular nose that is not unlike Russia's Bor series of Lozino-Lozinski hypersonic gliders.
When the author was at Wright-Patterson, interest in this waned quickly because of the
limited cross range available. Because of the longitudinal extent of the former Soviet
Union compared with the United States, the minimum L/D ratio to ensure a landing on
the continental land mass was less for the former Soviet Union than it was for the
United States-1. 7 for the Soviet Union versus 2. 7 for the United States .
Configuration 3 was a subscale research vehicle to evaluate the thermodynamic and
materials for hypersonic gliders. The nose and leading edge radii were full-scale size.
ASSET was successfully flown on a Thor intermediate-range ballistic missile (IRBM)
booster. One that was recovered after an ocean landing is on display in the U.S. Air
Force Museum in Dayton, Ohio . Configuration 6 was the first two-dimensiona l nose
applied to a conventional winged-body (configuration 5) in the United States.
Configuration 4 was a product of cooperation between the AFFDL (Alfred Draper) and
McDonnell Douglas Astronautics Company (Robert Masek) to develop a vehicle to
support the Manned Orbiting Laboratory (MOL). This concept was briefed to the U.S.
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Air Force in 1964, and elements of that configuration will be shown later. The intent
was a 9- to 12-person vehicle for crew rotation that could alternatively carry supplies to
the orbital station on a regular, frequent schedule (about one flight per week per
vehicle) . The variable geometry switchblade win g permitted landing with heavy loads
returning from space and eventually horizontal takeoff. The experimentally determined
configuration feature was the tail configuration. This configuration was wind tunnel
tested and demonstrated inherent stability and control at speeds ranging from mach 22
to landing speed.
Configuration 6 was a product of cooperation between the AFFDL (Richard D. Neumann)
and McDonnell Douglas Astronautics Company (Robert Krieger) to reduce t he drag of
hypersonic gliders. Based on the physics that a two-dimensional wedge has less drag
than a right circular cone of the same volume, these engineers devised the "spatular
leading edge." The wind-body configuration formed the basis of the X-20 and DynaSoar
configurations that had a limited hypersonic L/D ratio, primarily because of drag. With
the spatular nose, the nose wave drag could be reduced by 35 to 40 percent, thus
increasing the hypersonic L/D ratio. Configuration 6 was derived from the conventional
wing body, configuration 5.
Configuration 10 is an adaptation of the Russian "Star Body" concept that can enter in
one of three orientations and need not always have one side facing the flow
(compression side). The theory was that in a damaged situation, one of the three sides
would be available for a safe entry . The limitation of this configuration concept is a
small internal volume and a high ratio of wetted (surface) area per planform area that
reduces the hypersonic L/D ratio.
Martin Mar•tt ■
The X-24B was based on the FDL-8 X-23A
configuration. The different approaches
to hypersonic glider configuration are
best exemp lified by Figure 6. The X-
l' 0[
24A, built by Martin Marietta at its
Denver, Colorado, facilities, is a round USA f Mod1lied
fuselage configuration with outboard Th• X·24 I
high-dihedral-angle vertical tails. All
the configurations of this type have
serious lateral-directional stability
problems at low speeds and tend to roll
about the horizontal axis through the
fuselage . One designer, the Russian
Glebe Lozino-Lozinski, solved the
problem by employing variable dihedral Figure 6. Martin Marietta X-24 A & BResearch
tails. The AFFDL solved the problem by Gliders. X-24A based on USAF PRIME config uration.
using nonround configurations; that is,
the quest for high hypersonic L/D ratios led to the solution of the low speed problem.
Under an AFFDL program, Martin Marietta modified the X-24A into a flat-bottomed
configuration with trailing edge elevens called the X-24B, shown in Figure 6.
Comments by Bill Dana, the NASA pilot who flew the X-15 and the X-24A/B, about the
change in the slow speed performance of the X-24B confirmed the advantage of the
AFFDL approach. 9
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The design parameters that largely determine a spacecraft's weight are its configuration
and the amount of wetted or surface area relative to the planform area. The
hypersonic gliders shown in Figure 5 have differing values of wetted area to planform
area. Another important factor is the presence of wings, such as for configurations 5
and 6, which are wing bodies with a relatively thin wing or no wing, such as the lifting
body FDL-class hypersonic glider (configurations 2, 4, or 7). In this case the lifting
bodies have a shape advantage that reduces the amount of surface area that is thin or
subject to high heating. In the 1960s, when the U.S. Air Force's high-performance
lifting body was competing with NASA's modest-performance wing body, there was
much debate regarding the weight of these lifting concepts compared with that of a
ballistic capsule (see Appendix A). At that time, with the large sea-recovery fleets,
ballistic capsules were the only entry vehicles in either the United States or the former
Soviet Union. A number of studies in the early-to-mid-1960s attempted to rectify and
quantify the weight of a lifting entry vehicle compared with a ballistic capsule. In all
the discussion in the Mercury, Gemini, and Apollo programs, the cost of the sea
recovery was almost taken for granted, so the focus was on the cost of the vehicle
itself, not the entire vehicle system. The government assembled a chart representing
the relative weight of hypersonic entry systems-from ballistic to high-performance
(high L/D ratio) gliders-collected from contractor and government reports. The
relative weight was the system weight compared with that of a ballistic capsule with the
same payload capacity. The result was a correlation curve that showed the high
performance wing-body gliders could weigh as much as twice what a comparable
payload ballistic capsule weighed. This correlation was based on the L/D ratio of the
vehicle. Apollo has an L/D ratio of about 0.5, but the system was still a ballistic vehicle
with a very limited cross range. One correlation of the data is:
W/W0 = l + 0.1259aL/D)-0.1029aL/D)2 + 0.0621aL/D)3
(2)
W0 = the weight of a ballistjc capsule with the same payload
This correlation yields a high hypersonic L/D ratio glider with a weight almost twice that
of the ballistic capsule. In this correlation, different configuration concepts were mixed
and correlated as a single data set. A report cited in Appendix A (Stephens, 1965)
concluded, "Weight factor for lifting spacecraft results primarily from larger surface area
and only secondarily from the associated spacecraft environment and may be as large
as a factor of two greater than ballistic spacecraft."
Engineers at the McDonnell Douglas Astronautics Company examined the database and
concluded the large weight impact for a lifting spacecraft was as much a function of the
configuration as the L/D ratio. The engineers set out to separate the database into
families of like configurations. Where gaps existed, they established a configuration
that provided the L/D ratio sought that was based on the configuration rules for that
family. Three fam ilies were identified. The SV family configurations were based on
circular/ell iptical cross-section configurations that were characteristic of the HL-10 and
X-24A NASA configuration concepts. The FDL family configurations were based on the
trapezoidal delta planform configuration. And the MRS family configurations were based
on a McDonnell Douglas modified version of the FDL family, with an emphasis on
creating metal-radiative thermal-protection shingles that were flat, thereby reducing
the cost of the shingle and perhaps introducing an element of hardware
interchangeability. Altogether, 10 configurations from among the 3 families were
8
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designed, weighed, and performed using the same industrial fabrication capability. The
result was a curve representing each configuration family, as shown in Figure 7. The
three families are represented by the three parallel straight lines.
SV FAMILY
1.5
10
• ~-----
- - -
w
0.5
MRS FAMILY
W -- W0 • (1 +11
L/DJ
.1
0 1 2 3
Hypersonic LID
Figure 7. Detailed Design Analyses Show the Weight Trends are as Much a Function of Configuration
Family as Lift-to-Drag Ratio
W/W0 = K{l+0.090l!(L/D)]
K =1.000 for the SV Family
(3)
K = 0.9297 for the FDL Family
K = 0.7622 for the MRS Family
This illustrates that the configuration and its individual wetted area to planform area
can vary as much in their spacecraft weight as they can in their L/D ratio. The high
L/D-ratio configurations had weights comparable to same-payload ballistic capsules of:
sv = 1.9 FOL= 1.7 MRS =1.4
So the penalty for having a lifting-body configuration is less than expected if the
configuration characteristics are taken into consideration in the design and weighing of
the spacecraft. In addition, the reason the lifting spacecraft with high performance was
considered was to eliminate the need for sea recovery and therefore the cost of a
recovery fleet and the damage incurred by the spacecraft in a saltwater landing. The
9
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goal was to be able to recover the spacecraft at any airport in CONUS, to eliminate the
need for an overseas recovery site, and to eliminate the waiting required until a lower
L/D ratio could land in CONUS (up to 14 orbits for the Apollo capsule, or 21 hours). In
an emergency, that may be too long. The AFFDL's goal for the spacecraft to support
the Manned Orbiting Laboratory was no waiting but to be able to reach CONUS from
any arbitrary MOL position in its orbit. This was considered possible in the 1964-65
briefs to the government with respect to MOL, specifically the Model 176 configuration
the MDC proposed for the MOL support in 1964.
The hypersonic glider based on the FD L- 7C and the hypersonic air-breathing aircraft in
Figure 8 both have hypersonic L/D ratios in excess of 2. 7. In very practical terms, that
means unpowered cross ranges in excess of 4,500 nautical miles and down ranges on
the order of the Earth's circumference. So these two craft can depart from any location
of a low-altitude orbit and land in CONUS or in continental Europe. Both are
dynamically stable over the entire glide regime.
FDL-7 CID
Hypersonic
Glider
Bl ended Body
Hypersonic
Cruiser
Figure 8. High-Performance Hypersonic Glide Aircraft. Rocket boost-gl ide and air-breather cruiser.
The wing-body, cylindrical fuselage advocates have strongly criticized the lifting bodies,
contending that they are poorer configurations and much more complicated than the
conventional-wisdom wing-body configurations (see Figure 9). However, that is far
from the truth . The structural specialist sees this configuration as a lightweight
propellant tank and assumes it is t his consideration that drives the design. Rather, that
observation introduces problems for all other technical disciplines that are far more
difficult to rectify than a noncylindrical tank or a cylindrical tank in a nonsymmetrical
cross section. The lone lifting surface with t railing edge controls introd uces control
issues just as it did for the space shuttle.
10
UNCLASSIFIED/ ;'POlt err1e1J1tt U.!I!! er~t I
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Glide ang le of attack
Complex very high ~ 45° compared
curved TPS to 15° for high UO design
system
A ALang l
Wing- yl inder
nfigu rati n n ep
WB-004 Unsymmetrical
Inadequate vortices can
lateral-directional produce super
stability and sonic control
control problems
Cylinder heating
extends up to
55° to 60°
Wing training Poor hypersonic
edge control s lift-to-drag ratio
provide limited Thin wings limits glide range
control capability encounter
for a military severe heating
maneuver vehicle from both sides
i.e. "hot structure"
Figure 9. NASA Langley Wing-Body Configuration WB-004 With Generally Critical Areas for Wing Bodies
Identified
With a high entry angle of attack, the cross flow over the cylinder produces high
heating rates beyond the mid-cylinder line . With a lower L/D ratio, the down and cross
ranges are limited as to what might be achieved but more in line with NASA one
missed-orbit criterion. The thin wings are heated on both sides to create added thermal
problems, as well as added surface area to increase drag.
Al Draper and his team, together with Bob Masek's t eam at McDonnell Douglas
Astronautics, worked long and diligently to arrive at the FDL- 7/Model 176
configurations shown in Figu re 10 . The insert photo is Dale Reed 's model of the FDL-
7MC radio-controlled model at NASA Dryden. The AFFDL and MDC configu rations were
inherently stable at all operational ang les of att ack from at least mach 22 to landing
speed. The remainder of this report will focus on the characteristics of t his class of
lifting body . The statements in Figu re 10 were all based on wind tunnel data. A real
advantage of t he trapezoidal shape was not only flat metallic shingles but heating on
the sides and upper surface that was at least t hree-fifths that of the conventional
shapes. The glide range was such that th is configuration could land in CONUS from any
location on any inclination orbit from its current orbit with no wa iting .
11
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Flight tested nose
Lifting body configuration AFFDL-7C/D Lifting Body
eliminates thin wings with
tip design assures wing-like lift. High lift-to-drag Configuration Concept
100 °C temperature ratio and global glide range
All-flying control
surfaces provide
stability and control
from Mach 02 to 22
Heat pipe leading
edges assures
high LID leading
edge radii. Flat pane
TPS system
Trapezoidal cross
section increases
lift-to-drag ratio
and reduces side
and upper surface Anhedral fixed fins
heating by 1/3 with trailing edge
flaps assure subsonic
lateral-directional
stability and control
Figure 10. FDL-7C/D and FDL-7MC Lifting-Body Configuration. Offers inherent stability and control with
sufficient volume and high lift-to-drag ratio.
The switchblade wing version of the FDL-7MC was the preferred version for 1983
studies that were part of the McDonnell Douglas TAV (transatmospheric vehicle) effort;
that vehicle was powered by either an Aerojet Sacramento air turboramjet or an air
breathing rocket propulsion system. The inward-turning, variable-capture area inlet 10
provides the correct engine airflow from landi ng speeds to mach 5.5, as illustrated in
Figure 11.
Figure 11. FDL-7C/D with a DuPont Retractable Inward- Turning Inlet
The propellant tanks were cylindrical-segment, multilobe structures with bulkheads and
stringers to support the flat, metal-radiative thermal-protection shingles (very similar to
those fabricated by Goodrich Aerospace for the now-defunct X-33). The nose was
12
UNCLASSIFIED/ lFOR. OFFI&I.t.k Wliliii 8Plk¥
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transpiration cooled with a low-rate water-porous spherical nose. The sharp leading
edges (the same leading edge radius was used for the nose tip) were liqu id-metal heat
pipes. This approach was tested successfully during the 1964-68 timeframe and was
found to be equal in weight and far more durable than a comparable ceram ic
tile/carbon-carbon system. The AFFDL's experience with carbon-carbon leading edges
on the ASSET test vehicle convinced the Air Force it needed a more durable solution.
Figure 12 compares the FDL-7C/D and the McDonnell Douglas Model 176. The Model
176 had a power law nose that was essentially a curved spatular nose and a higher
sweep angle, resulting in the same usable volume but with a higher hypersonic L/D.
Sacrificed were the flat-panel thermal-protection shingles over part of the fore body.
Both configurations retained the X-tail configuration developed by Gil Gaumer of
McDonnell Douglas. At the time, the launch vehicle would have been a Martin Titan
IIIC. Had an engine with the performance of the Pratt & Whitney XLR-129 been
available, there would have been lateral recoverable, fuel/oxidizer tanks on either side
of the vehicle, with all of the engines installed in the hypersonic glider. This was similar
to the Lockheed Star Clipper (see Figure 38).
FDL-7 CID
ea
&a
5 met ers
e&a
circa 1962 Modified
Model 176 ..,., Gffl Wd1
Modified circa 1964
5 mete rs
Figure 12. Comparison of FDL-7C/D (top) and Model 176 {bottom)
13
UNCLASSIFIED/ ,«FOil OPFISl.t.L Y&liii 0PIL¥
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The author was aware of three people-James S. McDonnell, the AFFDL's Albert Draper,
and Russia's Glebe Lozino-Lozinski-who clearly understood the need for a long cross
range and down-range capability, not just for one missed orbit. Critics will observe that
Lozino-Lozinski had limited his BOR vehicles to an L/D ratio of 1. 7 to 1.8 and not the
2.7 to 3.0 required for Earth circumferential glide range. First, the longitudinal extent
of the former Soviet Union was twice that of CONUS, and an Earth circumferential glide
range was not necessary to ensure recovery within the continental Soviet Union;
therefore, a lesser L/D ratio was acceptable. Second, in personal conversations with
the author, Lozino-Lozinski indicated a Russian government agency forced him to limit
the glide range to ensure recovery in continental Russia and prevent escape to the
United States. In a further step to prevent escape, when the vehicle was in range of
CONUS, ground control disabled its deorbit system .
The need for a long cross-range and down-range capability so there is no waiting in
orbit in the case of an emergency or military need is presented graphically in Figure 13.
Interestingly, the greatest lateral-range (cross-range) requirement for no waiting is for
55° orbital inclination, the usual Russian orbital inclination. The nominal U.S. orbital
inclination is 28.5 degrees, with a waiting time of 8 orbits (approximately 12 hours) for
a space shuttle-class glider. At the International Space Station orbital inclination, the
orbital waiting time for a shuttle-class glider is 6 orbits. In comparison, Apollo's orbital
waiting time was about 14 orbits, provided the return trajectory included an Earth
parking orbit before entry into the Earth's atmosphere. The FDL-7 and Model 176 class
of gliders could immediately enter a return glide from their orbits. This provides a
significant advantage for the International Space Station operators and vehicle crew,
who need only enter hypersonic gliders attached to an orbital station and initiate
deorbit procedures to be on the ground in less than 90 minutes in an emergency.
Time (orbits) Required to Land in Cal ifornia or Fl orida
Orbital Altitude is 200 nautical miles, 230 .1 statute miles, 370.4 km
100 - - - - - - - - - - - - - - - - - - - - - - - ~
Minim um later I Range
For Once A Da Landing
,.....
.,.,"'
6b
~
'--' 60 1- + - - - - - > d - - - - - - ' r i - - - - - - + - r - - - -- - ~ -- - - - - i
<l
0
'.d
.5"'
u
..£1 40 l-l------',.,__---l-nao.~ ~---+------+---+--i'------ - . 4 - -------l
]
'.S
...
0 Mo•el 176
Operadq
20 1---""<,-"""'d-- - -- - + - - + - - - + - - - , < - - - l o " ' - - - Eanlope - - - - 1
Apollo
0'---------'-- -- - = : : . . L - - - - - - L - - -- - - ' - -- - - '
0 2000 3000 4000 5000
Spacecraft Lateral Range (nautical miles)
Figure 13. Sufficient Cross Range {L/D) Means There is No Waiting to Return
14
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UNCLASSIFIED/ /POil OFFIEiIAk Ulilii QPIL¥
Figure 14 compares the glide ranges and L/D ratios of selected hypersonic gliders. The
shuttle would be in the area of the nominal wing body on the righthand chart. From
Figure 13, the no-waiting cross range (lateral range) is 3,600 to 4,400 nautical miles.
That means the hypersonic L/D ratio needs to be in the 2.7 to 3.2 range. Appendix D
has down and lateral ranges shown as landing ellipses. The key to a successful landing
is a subsonic L/D ratio that is in the 4.5-or-greater range. The NASA round-bottom
configurations were not capable of that subsonic L/D ratio. The X-24B was in that
category and was therefore easily landed compared with the X-24A or Prime vehicles.
These high-performance gliders were unique to the AFFDL. The intent in case of a fire
would be to immediately evacuate to the hypersonic gliders and then depressurize the
station to control any fire (remember that Mercury and Gemini could be depressurized).
A crew could then be launched to recover the operation and repair the station.
j
30000
•· 4.0
T MDC data
·e
<O
~ 20,0CJ0 1----+---+---,..,,...--i1----i1---1 a
.S Nominal ::i
-~
Lifting
(I,)
Body § 2.0 t - --t----t--:::~ -t-'---+---1
~
ij ~
o::: 10,000 1----1■ --------+-----I >
C I
1.0 ~ =--....-.
ei
a
1----1-- - Apollo + Par
0 .....,___..___.___...___..____..____ 0
0 2,000 4,000 6,000 2.0 3.0 4.0 5.0 6.0 7.0
Lateral Ran e (nautical miles) Su sonic UD
Figure 14. Hypersonic Glider Characteristics
There is always the question about landing these blended bodies. Figure 15 shows the
blended body handling qualities to be very good, and therefore a pilot's fear factor is
less than with the X-15. Bill Dana said the X-24A was difficult to handle when landing
but that he could land the X-24B almost with no hands as it flared automatically . To
the author's knowledge, all of the wind tunnel tests showed inherent static and dynamic
stability over the entire speed range.
15
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Minim um Acceptable In itial Speed
Based on Acce lerated Stall M argin
Psychological LimitDue
to Excessive Initial Speed
Limit Due to Margin
Fl are Capabil ity
0
100 120 140 160 180 200 220 240 260
Fl are Initiation Speed (pe rcent of minim um spee d) %V min
Figure 15. Both Delta Planform Lifting Body {Dynasoar) and Model 176 Offer Superior Landing
Performance
Thermodynamics and Materials
The structure of Model 176 was based on diffusion-bonding and super-plastic forming of
flat titanium sheets. Forty years ago, the method was called "roll bonding" and
executed with the titanium sealed within an evacuated steel envelope and processed in
a steel rolling plant. With a lot of effort and chemical leaching, the titanium part was
freed from its steel enclosure. All of that has been completely replaced today by the
current titanium diffusion-bonding and super-plastic forming industrial capabilities. The
photo in Figure 16 is from a Society of Automotive Engineers book titled Advanced
Engine Development at Pratt & Whitney: The Inside Story of Eight Special Projects 146-
1971, by Dick Mulready. Chapter 6, "Boost Glide and the XLR-129 - mach 20 at
200,000 Feet," shows a surviving remnant from the 1960s program.
The super-plastic-forming, diffusion bonding that was so difficult in early 1960 is now
an accepted fabrication procedure. One of the F-1 S's major bulkheads was fabricated
from titanium sheet elements using this procedure instead of machining away more
than 90 percent of a titanium forging. Had the procedure been adopted as a product
manufacturing method, it would have eliminated the almost 2-year manufacturing cycle
in acquiring titanium forging of the wing spars and major bulkheads. Note that only 2.5
percent of the thermal heating enters the primary structure .
16
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•
UNCLASSIFIED//POil 8FFIEiIAk Wlilii QJslL¥
Aero Heating 97.5% Radiated
to Space
metal shingle
vacuum multi
layer insulation
- -
- --
----
----
----
----
-----------------------
----------------------
-
--
,-
--
--
--
--
--
--
--
--
--
--
--
--
--
--
--
--
--
-
--
--
--
-
-
--
-
--
-
-
--
-
-
--
-
-
--
-
-
--
-
-
--
-
-
--
-
-
--
-
-
--
-
-
-
-
, - - - - - - - - - - - - -- -- - - - - - diffusion bonded
super-plastically
formed structure
l-1gure 6. i Mel onnc 11 111m111m1 into 1rt' (('m1rrr1y of .!"Im Rohwn1
Figure 16. McDonnell Aircraft Company Roll-Bonded Titanium Structure (circa 1963). Today this
structure would be super-plastically formed and diffusio n bonded from RSR titaniu m sheets. u
Mulready's book mentions the McDonnell Doug las boost-glide strategic vehicle, as well
as citing key personnel at McDonnell Aircraft Company. Low thermal conductivity
standoffs set off the insulated-metal, therma l-protection -in sulated shingles from this
wall so that there was an air gap between them. The X-33 applied the metal shingle
concept, albeit with significant improvement in the standoff design and thermal
leakage, in the orientation, thickness, and weight of the sh ingles. This is one aspect of
the X-33 that can be appl ied to future spacecraft for a more reliable and repairable
therma l protection system than ceramic tiles. All efforts by the author to obtain
information on the shingles manufactured by Goodrich Aerospace have been met with
the response, "We lost the contract and are investing in more productive products."
The titan ium diffusion-bonded and super-plastically formed wall was both the primary
aircraft structure and the propellant tank wal l. The cryogenic propellants were isolated
from the metal wa ll by a metal foil barrier and sealed insulation on the inside of the
propellant tank. Significant testing of this structural approach confirmed its superior
capabil ities as a hypersonic rad iation-cooled structure.
The Model 176 was proposed for the Manned Orbiting Laboratory. It was a thoroughly
designed and tested configuration with a complet e, all -metal thermal - protection system
that had the same weight as ceramic til e and carbon -carbon concepts used for the U.S.
space shuttle but was sturd ier and could be repaired in a hanger or in orbit. A wind
17
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tunnel model of the McDonnell Douglas Astronautics Company Model 176 installed in
the McDonnell Aircra~ Company Hypersonic Impulse Tunnel for a heat transfer mapping
test is shown in Figu re 17. Note that, conforming to the piloting concepts of the 1960s,
it has a clearly distinct windsh ield . The model accomplished thermal mapping to
determine the heat transfer distributions on the body and upper fins.
Figure 17. Model 176 in the McDonnell Douglas Hypervelocity Impulse Tunnel (circa 1964} for
Thermographic Phosphor Heat Transfer Mapping, Includ ing the Upper Fin.
Among the important determinations that resulted from these heat transfer tests was
that the sharp-leading-edge, flat- bottomed, trapezoidal cross section reduced the
heating to the sides and upper surfaces, as shown in Figure 18. In the range of angles
of attack corresponding to maximum hypersonic L/D ratio, the sharp leading-edge
corner separates and reduces the upper surface heating. Because of this separation,
the isotherms are parallel to the lower surface and are 2,100 to 2,400 °F (1,149 to
1,316 °C) cooler than on the compression surface. The upper control fins are hot, but
there are approaches and materials applicable to control surfaces. The temperatures
shown are radiation equilibrium temperatures. With nose water transpiration cooling
(demonstrated in a flight test in 1966) and heat pipe leading edges (demonstrated at
NASA Langley in 1967-68), the temperatures of the nose and leading edges are 212 °F
and 1,300 °F (100 °C and 704 °C ), respectively. The thermal mapping enabled
identification of primary flow characteristics in the boundary layer of the vehicle. In
18
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Figure 18, radiation equilibr ium skin temperature is the skin temperature that resu lts
when the radiated t hermal energy stemmi ng from the skin temperature equals t he
input aerodynamic heating minus any conduction into the airframe.
900 ° F
Upper
Surface
---
!
•4,Joo°F
l
Lower
ALE= 0.5 in. Surface
• 4,200°F
1,ooo·F " 3,800°F
800 °F
1,200°F
=
RL E 1.6 in.
1,500 ° F
2,200 ° F
Figure 18. FDL-7C/D, Model 176 Entry Temperature Distribution. Upper- su rface heating is minim ized by
cross-section geometry.
Figure 19 is a thermographic phosphor
image of the model in Figure 17 at a 12-
degree angle of attack (maximum L/D
ratio) at mach 12. Even at mach 12,
there are vortices embedded in the
boundary layer. The three black dots are
heat transfer gauges that were used to
establish the value for qrefin Figure 19 .
The stagnation heat transfer was 25
times the reference value. Th is
technique, when calibrated with the ,, - no
reference heat transfer gauges, provided
a rap id and accu rate means to determine
heat transfer distributions with a
minimum of installed gauges. The
technique was adopted by other wind
tunnel facil ities, including the Arnold
Eng ineering Development Center at
Tu ll ahoma, Tennessee.
Figure 19. Even at Mach 12, Embedded Vortices in
the Boundary Layer Alter the Local Heat Transfer
19
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On other McDonnell hypersonic configurations with all-movable control surfaces, the
interface between the fin and the body became a critical heating issue for the rotating
shaft attach ing the fin to the body. This was an area of concern on this vehicle, and
specially instrumented fins were installed to measure the local heating. Again, the
thermographic phosphors were used to map the heating. Figure 20 shows the model in
Figure 17 at a maximum 48-degree angle of attack. Fin heating distributions were
made at 16-, 24-, 34-, and 48-degree angles of attack 12 and are shown in Figure 21.
The brighter the phosphor is, the lower its temperature is (the phosphor darkens as the
surface temperature increases). So the area adjacent to the body is at a lower
temperature than on the fin. In fact, examining Figure 21 shows that for all angles of
attack tested, there was always the cool layer adjacent to the body. So the fin
attachment journal/shaft would not be a thermal problem. At angles of attack lower
than 16 degrees, the heating became less intense. This tail configuration of a fixed
anhedral lower fin with trailing edge controls and an all-movable upper fin provided the
control authority over the entire mach range required for stability and control and did
not have a thermodynamic issue with fin attachment heating.
o = 48°
FIG. 6. ISODENSITRACER maps for lifting body tail fin for four
angles ot attack; contours show qr.ocw,/qun1u,,m,
Figure 20. Thermographic Phosphor Image of Figure 21. From L/D Maximum to Maximum Angle
Model 176 at Near- Maximum Angle of Attack of Attack, There is Always a Cool Sublayer
Adjacent to the Wall
With 1960 materials and manufacturing methods, about 95 percent of the aerodynamic
heating was radiated to space, about 2.5 percent was retained in the shingles, and
about 2.5 percent was transferred into the titanium tank/primary structure. Using
Goodrich Aerospace's standoff/attachment techniques developed for the X-33, today
around 0.5 to 1 percent of the aerodynamic heating would be transferred into the
titan ium tank/primary structure . The shingle material would also be better today.
Figure 22 shows a silicon carbide matrix re inforced with silicon carbide fibers that was
shown at the 1988 Paris Air Show. A combustor of this material was operated at 3,000
° F continuously for a number of days at SEP's Bordeaux plant, as witnessed by the
author. Unfortunately, SEP was subsequently taken over by another company and
promptly closed. The parts manufacturing at Bordeaux was truly impressive to
someone in space systems but too costly to a subsonic round engine manufacturer.
20
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One of the difficulties with silicon carbide (SiC) is that it is ridged fibers, like very-small
diameter rods. When the NASP team visited Japan in 1988, one of the very interesting
products of the UBE Corporation was Tyranno® cloth (see Figure 23). Composed of
strong, flexible fibers from natural Feldspar, Tyranno cloth handled and felt like tweed
cloth. The cloth could be wetted by liquid aluminum or titanium, and the NASP team
saw examples of both aluminum and titanium metal matrix composite (MMC) products .
For example, an aluminum MMC piston and connecting rod was being used in Kawasaki
racing engines. And a powder form of aluminum MMC was being used as a dry pigment
that was fused onto the surface of Kawasaki motorcycle mufflers. The MDC NASP team
foresaw many applications for Tyranno cloth for its NASP aircraft.
Figure 22. This 1988 SEP Bordeaux SiC/SiC Panel Figure 23. UBE Corporation's Tyranno Cloth. The
Could Sustain Temperatures of up to 3,000 °F. cloth was used industrially in Japan.
With SEP having subsequently closed, little of this
capability remains.
The FDL-7 and Model 176 configurations
always elicit comments that their sharpness
and the associated high heating rates make
them nonviable concepts. However, that is
not the case. Figure 24 shows a 1-inch
diameter, sintered-nickel nose tip attached
to a 1,000-psi water tank that sweats
water. The result is a functional sharp,
low-drag nose with a minimum-thickness
entropy boundary layer. 12 This tip was
flight-tested on a hypersonic glider (BGRV
flight in 1966) beginning at about 22,000
feet/second . Today, Aerojet Sacramento's
platelet diffusion-bonding technique would
make this a much easier task. Some Figure 24. A Porous Nickel Tip Oozing Water.
This test lasted 4,300 seconds before the arc
experimental evidence from the BGRV flight heater cathode-not the nose tip-failed.
indicates the water vapor film in the
boundary did act to reduce the heat transfer to the body aft of the nose.
21
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Al Draper assembled the ASSET (Aerothermodynamic Structural and System
Environmental Test) experimental flight-test program to evaluate current U.S. Air Force
and NASA materials for hypersonic entry veh icles . The intent was to launch a test
vehicle from an Air Force Thor IRBM in the 18,000 to 20,000 feet/second range and
recover the vehicle. The ASSET glider was approximately the forward portion of the X-
20 Dynasoar vehicle and is shown in Figure 25 after recovery. Except for the carbon
leading edges, the materials generally performed as required. As a result, a small team
of McDonnell Douglas Astronautics Company engineers and model builders began
working on an alternative approach for the leading edges. This team's efforts resulted
in the heat pipe leading edge shown in Figure 26, a series of formed stainless steel
tubes brazed together to form a leading edge based on NASA space shuttle
requirements. The tubes contained a stainless steel mesh wick and were filled with
metallic sodium.
figure 25. FOL ASSET flight-Tested From Orbital Figure 26. Heat Pipe Shuttle Leading Edge
Speeds To Evaluate 1960s' Materials. Carbon Designed and Built by McDonnell Douglas
leading edges proved the least durable. Astronautics. Tested extensively at McDonnell
Douglas Corporation, St. Louis, and NASA Langley, it
never failed .
The leading edge was tested in NASA Langley's 8-foot High-Temperature Structures
Tunnel, the NASA Langley Radiation Thermal Test Facility and the McDonnell Douglas
Graphite Thermal-Altitude Test Facility (graphite radiation heaters within a vacuum
altitude chamber). All of these tests showed the installed leading edge to be durable,
robust, and lightweight (equaled the installed NASA carbon-carbon leading edges).
Starts from cold tubes showed the sodium melts and began the heat pump process
without any difficulties. Because this leading edge was made by the engineers and
mechanics as a one of a kind, the tubes developed thermal shorts and other problems
over the span of the testing, all of which were rectified before the test continued.
Although brittle and difficult to manufacture, this leading edge met the
thermodynamacists' solution of a simple radiation structure, not a heat pump. 13
22
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The Qu Tube
In the 1990s, a colleague, Ying-Ming Lee, who then worked at MSE in Butte, Montana,
showed the author a copper tube about a foot long that was a heat pipe from his
colleague in Taiwan. If the tip was put into a cup of hot water, the other end almost
instantly was too hot to hold. If it was quickly put into a glass of cold water, that t ip
just as quickly became ice cold. As documented in the excerpted page below from the
University of Alabama, Huntsville, Annual Report, the apparent conductivity is greater
than copper and could not be melted, as the thermal energy would be removed so fast
that a significant temperature rise could not be attained. This could have significant
industrial application for the United States. The late Clark Hawk had obtained a 10-
foot-long Qu tube and tested it in his laboratory. The results were published in the
University of Alabama, Huntsville, Annual Report. However, as Clark Hawk discovered,
the Chinese team associated with Professor Qu in mainland China was not about to let
this discovery into American hands. Plus his team was composed of a number of
young, ambitious technocrats who thought they knew how to make their fortune. So
both Ying-Ming's and Clark's attempts to advance beyond a demonstration tube ended
in frustration . Any attempt to open the tube results in failure, as whatever is in the
tube reacts into an inert powder. The author has two smaller tubes in his possession.
With applications including hypersonic vehicles, nuclear power plants, and electronic
cooling, these devices would lead to an economic breakthrough in practical therma l
control.
23
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the u Tube, or uperrube, is somewhat conrrovcrsial. meter. A water calorimeter is used ro measure the h~ I
According co the inventor (Patenl o. 6,132,823) and to conducred along rhe Supertube. Fins, hown in Figu
claims by the company and their quoted rcsul of rescs 19, are necessary ro transfer the high heal Bux ro d1e
conducred by ranfurd Research lnslirute, it reportedly warer; these were fabricaced in such a way that they can
has an effective thermal conductivicy of the order of J 0 be easily auaclled to the uperrub using hos clamps.
co J00 times grearer than cha of conventional liquid-va A1 o the fin design increases turbulence in the flow
por heat pipes, and over 30,000 tim char fan equiva through the heat exchanger. Having the turbulent Aow
lemly sized solid rod of silver. Repons cited also have increases the heat transfer rate inro the warer, and also
indicated puzzling temperature di tribtttions n. o ur di ourages boiling, which could oc ur with rJ1e high
with these rubes unlike conventional thermal condu - heat .Auxes used.
cor and liquid-vapor heat pip . 11,e tube al o app ar CnlA l•111ptealb2: Dt£ribulon Afon;Qu l'11oeaidCClflPII' f1,1blful
to h ve the ability co function at v ry high re.mperatures, 0111• U.t:S.oi ec,..- and 'ID,,2045 tvptl'l:11
,a - - - - - - - - - - - - -
even up ro the melting point of the material used, nd
ro supporrvery high heat .Auxes. ur cesrs to dare sup
port the high tempera.cure capabilities in addition co
the high thermal conductivities. However, the thermal
ondu ivity is so high chac accurately measuring the
value i very difficult. We have ther fore acquired nine
10' long uperrubes, 5/16" in diameter, and have er up
a method for decermining the thermal conductivity us
,- - ~- --r-- - r - -- - r - - r - - - ,
ing high heat flux, a water cooled ca.lorimcte.r, and a rake :( ,c l3 u ll
of over 30 carefully ca.librared thermistor . Thi appara• __,__,_,,
• :.t..1-;a
X ""'
_,_..,.
_____..,,..
--c:i..-:.....mi-:
tus should provide an accurate means of d rermining the ~°!Altl"l
. : . ...... On ~,-c...-.. tC'
~tk.-•e.u...,m
thermal conductivity, and will also allow us 10 1eck the
high heat flux capabiliiy of ch rube and co po • Figure 17. Compari.Jon of&enrially Co,uwu Superrube
ible puzzling rempe.rarure di rributions. Tt:mperarure Dfrtribution with 1heort:ticai Dutribution
for a Standard Copper Tithe and Rod. Showing that the
Figure 17 shows a data sec in which the ccmpe cure 7herma/. Conductivity Must be at Len.st 30,000 limts that
across the length of a 1O' long cube hear.ed rom the end
ofCopper.
and cooled in air is essenrially connanr, wh r s similar
si.ze copper tube would have the tcmpe cure djsuibu
tion shown in the bocrom curve. Increasing the th rmaJ
conductivicy, k, of che copper by fuctors fr m I 000 to
30,000 shows agreement becween the analysis and the
daca ar 30,000 times chat of the copper, akhough chi is
only a lower limit on the actual condu tiviry. Increas
ing the fucror even more do s nor produ ea di mibl
cllange in the curve relative to the data. We have also
tried ocher methods to estimate the high th rmal con
ductivities we have measured, such the lngenhouss
rechniq_ue; these results also indicate very high th rmal
conductivities. However, we needed a more accurate
mean than previous tests with free conv ion o led
rubes with thermocouples, and therefor d ·sed the
rest appararus shown in Figure I .
The 10 fr Superrubes are heaced by three 2kW oil
rs. Power for the heaters comes from a -ph 2 Figure 18. Fxperimmra/. Appararusfor 10 ft Qu Super
output power concroUer with two I gs fu ed at 20 tubes.
:\mps. For saf-ccy reasons the pow r contr0l1cr and fus
placed in an endosure along with a Watmode" power
24
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Rocket Propulsion
The photo in Figure 27 is from the Society of Automotive Engineers book, Advanced
Engine Development at Pratt & Whitney: The Inside Story of Eight Special Projects 146-
1971, by Dick Mulready. Chapter 6 of this book, "Boost Glide and the XLR-129 - mach
20 at 200,000 Feet," mentions the McDonnell Douglas boost-glide strategic veh icle, as
well as citing the key personnel at McDonnell Aircraft Company. Some time ago, a
model showed up on the desk of a now-Boeing employee that was not readily
identifiable. The author has an original model of this boot-glide strategic veh icle, and it
matched the unidentified model shown in Figure 27.
The XLR-129 was a shuttle-class engine that operated with turbopump exit pressures of
3,500 psi. It was brought to full pressure operation for the U.S. Air Force in just over 3
months. 14 In comparison, the space shuttle main engine (shown in Figure 28),
operating with a lesser turbopump exit pressure, required 3 years to reach full pressure
operation and never demonstrated reusability without overhaul.
Figure 27. Boost-Glide Strategic Vehicle With Pratt Figure 28. XLR-129 , a Shuttle-Class Eng ine That Ran
& Whitney XLR-129 Rocket Engine Installed, Circa for 40 Test Cycles With Only Component Maintenance and
1964 No Engine Rebuilding
The final paragraph of chapter 6 in Mulready's book contains the following passage:
"The liquid oxygen turbopump was the next component in line. However, before it was
funded, NASA had started the space shuttle campaign, and the Air Force gave the XLR-
129 program to NASA granting free use of the existing hardware to Pratt & Whitney.
NASA promptly canceled the liquid oxygen turbopump because it would be unfair to our
competitors to fund it." With the demise of the XLR-129, a rocket engine with a run
record of 42 simulated flights (in the test chamber) without any overhau l disappeared.
Th is engine was really the type of hardware a Kelly Johnston would oversee-that is,
the best application of the industrial capabilities available in the skilled mechanics,
engineers, and manufacturers. The only other eng ine of its class is the Russian RD-
0120 engine
Fuente: archivo UAP oficial del gobierno de EE.UU. (dominio público) · war.gov/ufo ↗ · ver en el archivo de Nodriza