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Langley Contributions
to the F-111
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The Variable-Sweep
Wing Concept |
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The F-111 was the
first production military aircraft to capitalize
on years of Langley research to develop a variable-sweep
wing aircraft. Variable-sweep wings provide significant
benefits to the aerodynamic performance of aircraft
that operate over a wide range of altitudes and
airspeeds. An excellent summary of the history
of Langley’s role in the development of this
breakthrough concept was given by Edward C. Polhamus
(contributor to variable-sweep research and the
F-111 program) in his 1983 American Institute of
Aeronautics and Astronautics (AIAA) Wright Brothers
Lecture (ref. 4).
As the advantages of wing sweep for enhanced
aerodynamic performance became known near the end
of World War II, designers considered several approaches
to providing variable sweep for efficient characteristics
at both low and high speeds. The German Messerschmitt
P-1101 research configuration had provisions for
three ground-adjustable sweep angles; however,
the aircraft was only carried to the prototype
stage and flight tests were never conducted. Although
it never flew, the P-1101 was captured and brought
to the United States, where it was later evaluated
by Bell Aircraft Corporation in a study that strongly
influenced the design of the Bell X-5 research
aircraft.
The first wind-tunnel tests of variable-sweep
concepts were conducted at Langley in the mid-1940’s.
These tests included the now familiar symmetric
variable-sweep wing, as well as the variable oblique-wing
concept (free-flight model tests by John P. Campbell).
Researcher Charles J. Donlan, who would later become
Deputy Director of Langley, initiated tests of
an existing model of the famous Bell X-1 (first
aircraft to exceed the sound barrier) in the Langley
300-MPH 7- by 10-Foot Tunnel in 1947 to explore
the challenges of variable sweep. The results of
the study, which used a single centerline pivot
for the wing, clearly showed that such an arrangement
resulted in excessive longitudinal stability and
marginal maneuverability with the wings swept aft.
To be successful, the variable-sweep concept would
have to minimize the dramatic increase in stability.
Several methods were explored, and it was concluded
that some type of variable longitudinal translation
of the pivot point was required—although
this solution was undesirable from a weight and
complexity perspective.
Time-lapse photograph of the
Bell X-5 shows the wing at various sweep angles.
The radical British Swallow configuration
in the Langley 16-Foot Transonic Tunnel
with variable sweep arrow wing and over and under
wing-mounted engines.
In the late 1940’s, under somewhat reluctant
Air Force sponsorship and based on the German and
U.S. studies, the Bell X-5 became the Air Force
and the National Advisory Committee for Aeronautics
(NACA) workhorse for variable-sweep studies. Wing
sweep on the X-5 could be continuously varied from
20 deg to 60 deg through the use of a translating
wing-pivot mechanism. The X-5 configuration was
tested in the Langley 300-MPH 7- by 10-Foot Tunnel,
the 8-Foot Transonic Pressure Tunnel, the 4- by
4-Foot Supersonic Pressure Tunnel, and the 20-Foot
Vertical Spin Tunnel. The first flight of the X-5
was made on June 20, 1951. At about the same time,
interest in variable sweep was growing in the United
Kingdom, and discussions between Langley Assistant
Director John Stack and the British resulted in
exchanges of data and proposed cooperation for
research efforts.
In 1952, the Grumman XF-10F variable-sweep aircraft
began flight tests. (The XF-10F had also been tested
in the 7- by 10-Foot High-Speed Tunnel at Langley.)
Although the aircraft was underpowered and subsonic,
it demonstrated the aerodynamic performance advantages
of variable sweep. However, the wing-pivot translation
feature added considerable weight to the configuration,
and the performance of the XF-10F could not compete
with the rapid increase in supersonic performance
that was being demonstrated on most military aircraft
of its size.
At this point, military interest in the variable-sweep
concept rapidly decreased. It appeared that supersonic
flight could not be sustained with state of the
art engines at that time. Also the heavy, complicated
translating wing-pivot mechanism was not an acceptable
penalty when compared with the performance of a
moderately swept fixed wing. However, as has been
the case throughout Langley’s history, a
few visionaries correctly anticipated the future
military aircraft requirements for sustained supersonic
flight and successfully advocated for fundamental
research to continue to develop and optimize the
variable-sweep concept. Led by John Stack and Charles
Donlan, Langley researchers joined with the British
in a North Atlantic Treaty Organization (NATO)
sponsored cooperative study of variable-sweep concepts
that included extensive tests in the Langley 7-
by 10-Foot High-Speed Tunnel and the 16-Foot Transonic
Tunnel.
The British and Langley cooperative tests in
1958 included work on the British Swallow configuration,
which was a radical tailless slender arrow-wing
configuration with variable-sweep wings and pivoting
wing-mounted engines. The Swallow exhibited numerous
stability and control problems. Langley brought
three variable-sweep configurations to the study,
including two that used pivot locations in the
fuselage near the trailing edge of the inner wing
and folding tail control surfaces to maintain stability
levels. Thomas A. Toll, responsible for the stability
portion of the variable-sweep program, assigned
William J. Alford, Jr. and Edward Polhamus to the
task. In 1959 they arrived at the breakthrough
solution, which was to locate the pivots of the
movable wing panels to positions outboard of the
fuselage. With this outboard wing-pivot arrangement,
sharing of lift between the fixed inner wing and
the movable outer wing panels minimized the movement
of the aerodynamic center of lift. A fourth configuration,
known as configuration IV, was added to the research
program to confirm this design concept. The configuration
was tested across the speed range in Langley tunnels
with great success.
Encouraged by their success, the Langley team
conducted in-depth analyses of variable-sweep aircraft
representative of configurations for Navy and Air
Force missions. Tests of all types—force
and moment, dynamic, and free flight—were
conducted in virtually all of the major tunnels
at Langley. Variable-sweep configurations applicable
to the Navy combat air patrol mission were studied
in a program known as CAP, and the Air Force tactical
aircraft mission was the basis for configurations
studied in a program known as TAC. Extensive wind-tunnel
tests and analysis by Langley researchers had matured
the variable-sweep concept.
Alford and Polhamus became internationally recognized
for their research on the outboard wing-pivot concept,
and they mutually hold the U.S. patent on this
revolutionary discovery, which has been successfully
applied to numerous U.S. and foreign military aircraft.
Configuration IV, which was
the breakthrough model for variable-sweep technology.
James L. Hassell, Jr. with a
free-flight CAP model that was flown in the Langley
Full-Scale Tunnel.
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The Tactical Fighter
Experimental (TFX) Program |
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In 1957, the U.S.
Navy requested industry responses for the design
of a low-altitude strike fighter. John Stack briefed
senior Navy managers that a proposed British low-altitude
strike fighter, the NA-39, would be much more advanced
than the Navy aircraft. He also suggested the application
of variable sweep to leapfrog the capabilities
of the NA-39. Following briefings by Langley personnel
to the Navy, the mission specifications for the
new Navy fighter were expanded to include multimission
capability with a requirement that variable-sweep
applications be studied. The request for proposals
went to industry in early December 1959 and set
the stage for what would ultimately become the
Tactical Fighter Experimental (TFX) Program.
Meanwhile, the Air Force Tactical Air Command
(TAC) Requirements Division at Langley Air Force
Base (adjacent to the Langley Research Center)
was attempting to define a replacement for the
F-105 fighter-bomber aircraft. TAC was interested
in an aircraft that could carry nuclear weapons
internally, fly transatlantic routes without refueling,
operate from semiprepared fields in Europe, have
a top speed of Mach number of 2.5 at high altitudes,
and fly at high subsonic speeds at low altitudes.
The aircraft would perform a “low-low-high”
mission, wherein it would cruise into the vicinity
of the target at low altitudes and subsonic speeds,
perform a low-altitude dash to the target at high
subsonic speeds, and perform a high-altitude, long-range
cruise back to base at subsonic speeds. The Mach
number of 2.5 capability would be used for high-altitude
engagements against enemy fighters. Initial analysis
by industry of the request indicated that a fixed-sweep
aircraft capable of meeting the requirements would
weigh in excess of 100,000 lb (too heavy for unprepared
fields) and demand the attributes of low sweep
for transatlantic flight, but high sweep for the
high-speed requirements. TAC was therefore in a
stalemate without a viable design approach to its
requirements.
John Stack approached the TAC planners in 1959
with the benefits of variable sweep to enable an
aircraft to meet the requirements. The extended
ferry range that is provided by variable sweep
was of prime importance to TAC, since estimates
indicated that transatlantic range might be possible.
Together with the commander of TAC, Stack laid
out a realistic set of aircraft performance requirements
that included the desired low-altitude dash capability
at high subsonic speeds. Unfortunately, as the
requirements went through the TAC system for approval,
the final specifications called for a 210-n-mi,
sea level dash at a speed that had increased from
a Mach number of 0.9 to a Mach number of 1.2. Upon
learning of the supersonic low-altitude speed requirement,
Langley quickly informed the Air Force that this
capability was impossible to meet for the range
specified. Nonetheless, TAC was committed to the
unrealistic specification. (In flight tests of
the F-111A in 1969, the actual low-altitude supersonic
dash performance of the aircraft was only 30 n
mi.)
In 1960, the Air Force and the Navy were both
attempting to develop new fighter aircraft. The
Kennedy administration’s Secretary of Defense,
Robert McNamara, ordered the development of a single
aircraft for both the Air Force and the Navy (to
be led by the Air Force), called the Tactical Fighter
Experimental (TFX). Mr. McNamara defined the basic
mission requirements when the Air Force and Navy
could not agree, and in October 1961, a request
for proposals (RFP) was issued to industry. Boeing
won all four stages of the competition that followed,
but McNamara overruled the source selection board
and decreed on November 24, 1962, that the General
Dynamics and Grumman Team would build the TFX.
Langley’s variable-sweep wing concept was
nationally recognized as the technical key that
would unlock the capabilities of the TFX. McNamara
said (ref. 27)
New developments in engine performance
and in aerodynamics, particularly the variable-geometry
wing concept evolved by NASA, now make it possible
to develop a tactical fighter that can operate
from aircraft carriers as well as from much shorter
and cruder runways, and yet can carry the heavy
conventional ordnance loads needed in limited war.
As an example of the plaudits given Langley for
the variable-sweep contribution to the evolving
TFX program, the following comments of editor Robert
Hotz of the internationally acclaimed Aviation
Week magazine appeared in the magazine on December
3, 1962:
Underlying the whole TFX concept
is one of the solid, basic technical explorations
of the old National Advisory Committee for Aeronautics
(NACA) that did so much to keep this country the
international leader in supersonic aircraft development.
Without the fundamental research into the variable-sweep
wing and the detailed development of this principle
by the Langley research laboratory group headed
by John Stack, the current TFX concepts of both
final competitors would have been impossible...
When Congress convenes again and begins carping
over the Fiscal 1964 NASA budget for aeronautical
research, the full story of the Langley contributions
to the TFX program should be hammered home as an
example of how these research and development investments
eventually pay substantial benefits.
After the TFX contract was awarded, Langley,
Ames, Glenn, and Dryden all supported the F-111
development program. Because of the strong interest
in this aircraft, and the large magnitude of NASA
support, the program was rigorously managed and
documented, beginning in November 1962. Mark R.
Nichols, Laurence K. Loftin, Jr., Edward Polhamus,
Jack F. Runckel, Theodore G. Ayers, and M. Leroy
Spearman were the leaders and spokesmen for the
F-111 support activities at Langley. In addition,
Polhamus served as the Langley focal point for
overall F-111 activities and spent considerable
personal time in coordination of tunnel requests
and NASA, industry, and DOD joint meetings.
In 1963, political turmoil surfaced as a special
Senate subcommittee chaired by Senator McClellan
of Arkansas held hearings on the award of the TFX
Program. This committee gathered a considerable
amount of Langley data, briefings, and testimony
during various phases of the hearings, which took
place over several years.
For the first 3 years of the F-111 development
program (1963 to 1965), a total of nearly 20,000
hr of tests were conducted in NASA wind tunnels
(about 15,000 hr at Langley). Over 15 wind tunnels
were utilized, making the F-111 program the most
extensive wind-tunnel support effort ever provided
for one aircraft by NASA or the NACA. By 1968,
over 22,000 hr of tunnel tests had taken place
at Langley. (In contrast, Langley expended 5,000
hr for development of the F-105.) The large number
of test hours was the result of the multiple versions
of the aircraft (F-111A, F-111B, RF-111, and FB-111),
the addition of wing sweep as a test variable,
a vast number of external store configurations,
and concentrated technical assaults on a multitude
of problems—especially the transonic drag
issue. During the development effort, the Langley
staff also participated in numerous F-111 advisory
and assessment teams and briefings for DOD, Congress,
and industry. For example, Edward Polhamus presented
the scope and results of the NASA effort to the
McClellan Committee during its second session in
1970.
The first F-111A flew in December 1964, and the
first F-111B flew in May 1965. The most positive
result from early flight evaluations was the very
satisfactory behavior of the variable-sweep wing
system. However, the aircraft were judged to be
sluggish and underpowered. Furthermore, the engines
exhibited violent stalling and surging characteristics.
An outstanding, in-depth discussion of the details
of the initiation and early years of the F-111
development program is given in the book Illusions
of Choice by Robert F. Coulam and Robert S.
McNamara (ref. 27).
|
Aerodynamic Performance |
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On December 19,
1962, representatives of General Dynamics and Grumman
visited Langley for discussions of the supersonic
performance of the F-111. The manufacturers were
informed that the supersonic trim drag of the aircraft
could be significantly reduced and maneuverability
increased by selecting a more favorable outboard
wing-pivot location. Unfortunately, the manufacturers
did not act on this recommendation, and it was
subsequently widely recognized that the F-111 wing
pivots were too far inboard. (It should be noted
that the F-14 designers, aware of this shortcoming,
designed the F-14 with a more outboard pivot location.)
The F-111 subsequently exhibited very high levels
of trim drag at supersonic speeds during its operational
lifetime.
In March 1963, the initial supersonic tests of
the F-111 in the Langley Unitary Plan Wind Tunnel
by David S. Shaw confirmed the Langley expectations
of high trim drag at supersonic speeds. A month
later, tests conducted by Theodore Ayers of a 1/24-scale
F-111 model at transonic speeds in the Langley
8-Foot Transonic Pressure Tunnel indicated that
the transonic drag was considerably higher than
General Dynamics predictions. Therefore, a large
drag reduction had to be accomplished to meet mission
requirements. Several discussions between Langley
and General Dynamics were also held to define approaches
to improve supersonic maneuverability. Langley
continued to emphasize the importance of wing-pivot
location and recommended a change in pivot location
and a forward shift of the wing as a solution to
the problem. It was decided that the modified wing
suggested by Langley would be built and tested.
Supersonic tests of the Langley wing modification
indicated a large increase in maneuverability that
would allow the F-111 to approach the proposed
maneuverability levels. However, because of changes
in maximum cross-sectional area, some transonic
drag penalty would be expected with the modification.
At a later meeting in June between Langley, the
F-111 Systems Program Office (SPO), and General
Dynamics, the positive results of the modified
wing were discussed, but the F-111 SPO expressed
concerns over any possible transonic drag penalty,
the engineering effort required for the change,
and potential schedule slippage. The Air Force
cancelled further studies of the wing modification.
Supersonic tests of the early
F-111 design in the Langley Unitary Plan Wind Tunnel
in 1963.
Meanwhile, a controversy erupted over the discrepancy
of transonic drag estimates between the data obtained
by Theodore Ayers in the Langley 8-Foot Transonic
Pressure Tunnel and data obtained in the Cornell
Aeronautical Laboratory 8-Foot Transonic Pressure
Tunnel. The Langley staff investigated these differences
and concluded that the lower drag measured in the
Cornell facility was probably due to interference
effects caused by an oversized model and the large,
blunt support sting. Langley offered to investigate
the problem by testing a mock-up of the Cornell
support system. In October 1963, special tests
were conducted in the Langley 8-Foot Transonic
Pressure Tunnel with a mock-up of the Cornell support
system. The results of the test indicated a very
large buoyancy effect, which accounted for the
erroneous low transonic drag measurements in the
Cornell tunnel. As Langley’s support for
the F-111 continued, the staffs of the 16-Foot
Transonic Tunnel and the 8-Foot Transonic Pressure
Tunnel remained sensitive to wind-tunnel accuracies,
and numerous tests were conducted with the same
model in these two tunnels to establish confidence
in the Langley projections of transonic drag.
With a much higher level of drag established,
emphasis was placed on reducing the transonic drag.
The drag problem included large afterbody and nozzle
drag components caused by high closure slopes (rapid
variations in external aircraft contours), high
drag components of the cockpit and inlets, and
very high boundary-layer spillage drag. Dr. Richard
T. Whitcomb and his staff conducted exhaustive
tests of engine alignment, wing and tail twist,
and even antishock bodies to reduce drag. Studies
by Ayers concluded that the area ruling for the
F-111 was unsatisfactory.
Also in 1963, an F-111 aerodynamic consulting
group consisting of Air Force, Navy, and NASA members
met and concluded that the transonic aerodynamic
performance of the F-111 would be considerably
below the requirements for the projected missions.
Calculations based on Ayers’ drag measurements
predicted that the aircraft would only have a range
of about 20 to 30 n mi for flight at low altitudes
and speeds of a Mach number of 1.2, in contrast
to the 220 n mi capability predicted by General
Dynamics. Actual flight tests later verified Ayers’
projection. The group, which included Polhamus
and Spearman, recommended that aft-end modifications
suggested by Langley should be studied for transonic
drag reduction.
As concern over the aerodynamic performance of
the F-111 increased, Charles Donlan and Edward
Polhamus briefed the Assistant Secretary of the
Air Force in April 1964 on the situation. They
recommended that the staff of the Langley 16-Foot
Transonic Tunnel define the benefits of Langley
conceived aft-end modifications. It was also suggested
that the wing with the longer span of the Navy
aircraft be used on the Air Force aircraft. Polhamus
and Spearman also briefed Air Force General Schreiver
and Navy Admiral Schoech a month later on the transonic
drag problem.
During May 1965, representatives of Grumman visited
Langley several times to discuss methods of improving
the acceleration and maneuverability of the Navy
F-111B. Modifications considered by Grumman included
several of the early Langley suggestions, such
as a modified wing and pivot location, a straightened
tailpipe, and an improved interengine fairing.
In addition, Grumman examined a modified horizontal
tail, alternate missile arrangements, and an aft-fuselage
modification. Although these modifications never
came to fruition for the F-111B, the discussions
had a large impact on the later design of the F-14
by Grumman, which became an outstanding Navy aircraft.
Langley also assessed the effect of the bomb
bay cavity and doors on directional stability in
the Unitary Plan Wind Tunnel; the impact of missile
carriage at supersonic speeds in the Unitary Plan
Wind Tunnel; the aerodynamic damping during subsonic,
transonic, and supersonic flight in the 30- by
60-Foot (Full-Scale) Tunnel, the Unitary Plan Wind
Tunnel, and the 8-Foot Transonic Pressure Tunnel;
the transonic and supersonic flight of the strategic
bomber (FB-111) and reconnaissance (RF-111) versions
of the F-111 in these same wind tunnels; and the
development of the F-111B by the Navy, until it
was cancelled on July 10, 1968. Over 25 test entries
in the Unitary Plan Wind Tunnel were made for the
F-111 variants.
|
Propulsion Integration |
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Hot-jet tests of
a 1/9-scale ejector nozzle at transonic speeds
were conducted in April 1963 in the Langley 16-Foot
Transonic Tunnel to begin what would become an
extensive series of F-111 propulsion integration
studies. (The staff ultimately con-ducted 17 entries
of F-111 models or components during the program.)
Later that year, a 1/6-scale inlet model was tested
to determine the effects of aircraft nose shape
(Air Force and Navy) and engine inlet spike configuration.
Following this test entry, the inlet, cowl, and
spike geometry were revised.
In 1964, General Dynamics, in consultation with
the 16-Foot Transonic Tunnel staff, completed fabrication
of a 1/12-scale model designed to investigate propulsion-airframe
integration characteristics. This model represented
the most realistic and complex model of a military
fighter ever tested in the 16-Foot Transonic Tunnel.
The model had multiple strain-gage balances and
balance arrangements to independently measure thrust,
drag, and thrust minus drag. It also contained
three independently controlled internal flows to
simulate the F-111 blow-in-door ejector exhaust
system: a hot hydrogen peroxide primary-jet flow
system, a high-pressure air secondary-flow system,
and a low-pressure air boundary-layer bleed system.
In addition, the model contained fully variable,
aerodynamically actuated blow-in doors and ejector
shroud flaps on the nozzles. Tests in 1964 on this
model in the 16-Foot Transonic Tunnel indicated
a significant nozzle-thrust deficiency that was
associated with an adverse fuselage afterbody flow
field. At the end of the year, additional tests
of the hot-jet model were directed at a wide range
of ejector nozzle geometries in an attempt to solve
nozzle-thrust and flutter problems. Langley also
initiated studies directed toward reducing the
large base drag associated with the short interengine
fairing (interfairing). Relatively large drag improvements
were obtained with a long interfairing design conceived
by Langley.
Unfortunately, the naval F-111B configuration
was too long to met the requirements for aircraft
carrier elevator spotting (compatibility of the
aircraft dimensions with the elevator on the aircraft
carrier that transports aircraft to and from the
flight deck and the lower hangar area). Follow-on
tests of the 1/6-scale inlet model during 1964
included studies of the effects of an extended
nose (for the RF-111), a weapons bay pod, and bleed
doors.
Jack Runckel and Edward Polhamus briefed Air
Force management in October 1964 on the ejector
nozzle problems of the F-111. Since the nozzle
problem was associated with the aircraft aft-end
flow field and since the aft-end drag was relatively
high, the Langley representatives recommended that
improvements to the back end of the aircraft be
investigated as a high priority item. Runckel subsequently
became a key figure in a joint F-111 nozzle committee
that included the F-111 SPO, the Bureau of Naval
Weapons, Pratt and Whitney, General Dynamics, and
NASA. In a meeting of this committee in January
1965, Runckel proposed that a truncated, concave
base interfairing be investigated as a means of
reducing drag at transonic speeds, while still
meeting the length restriction imposed by the F-111B
naval version.
Inlet tests of an F-111 model
in the Langley 16-Foot Transonic Tunnel.
In February 1965, tests of the 1/12-scale hot-jet
model indicated large improvements in transonic
performance because of airframe changes in the
region between the nozzles. This model configuration
had the best aerodynamic characteristics to date
in the program. A few months later, a meeting was
held at Langley with representatives from the F-111
SPO, the Bureau of Naval Weapons, General Dynamics,
Grumman, Pratt and Whitney, and Langley. At that
meeting, it was agreed that emphasis should quickly
move to reducing transonic drag by developing an
optimum engine interfairing and speed bumps (additional
area added to shape the aircraft to comply with
Whitcomb’s area rule). In December 1965,
the hot-jet model was tested to develop the interfairing
and define a configuration that would be flight
tested in early 1966. The aft-end modifications
from this effort were ultimately adopted and resulted
in a significant improvement in transonic drag.
The early F-111A exhibited numerous engine problems,
including compressor surge and stalls. NASA was
a participant in finding solutions to these problems,
as its pilots and engineers flew test flights of
the aircraft to determine inlet pressure fluctuations
(dynamics) that led to these events. Eventually,
as a result of NASA, Air Force, and General Dynamics
studies, the engine problems were solved by a major
inlet redesign.
The F-111 program brought many difficult challenges
in propulsion-airframe integration to the staff
of the Langley 16-Foot Transonic Tunnel. These
problems were extremely complex and demanded timely
solutions in a highly visible, controversial national
program. However, the staff responded with outstanding
technical expertise and innovation. By the end
of 1967, 19 F-111 related model entries had been
completed in the 16-Foot Transonic Tunnel, and
over 283 configurations were investigated during
12.5 months of tunnel occupancy. This effort ultimately
required participation of almost every 16-Foot
Transonic Tunnel staff member. Staff members who
made significant contributions to the F-111 program
included Richard J. Re, Odis C. Pendergraft, Jr.,
Richard G. Wilmoth, Charles E. Mercer, Francis
J. Capone, and Bobby L. Berrier. As a result of
the F-111 program, the research capability on aft-end
transonic drag problems greatly increased. This
research capability contributed to all subsequent
high-performance military aircraft and placed Langley
in a position of world leadership in this critical
technical area.
|
High-Angle-of-Attack
Characteristics |
|
Initial free-flight
model tests of the F-111 configuration were led
by Peter C. Boisseau in the Langley 30- by 60-Foot
(Full-Scale) Tunnel in October 1964. During the
flight tests, the wing sweep was varied from 16
deg to 72.5 deg. In December, the model was flown
to determine the effects of stability augmentation
in roll and pitch for the clean and landing configurations.
The flight tests were extended to high angles of
attack, including the stall.
When the model was flown to high angles of attack
with the wings at the 50-deg and 72.5-deg sweep
conditions, the model exhibited a sudden, uncontrollable
yaw divergence prior to maximum lift. General Dynamics
personnel, including the test pilot who was scheduled
to make the first high-angle-of-attack flights
with the aircraft, witnessed the tests.
The model free-flight tests also indicated an
unusual unsteadiness in lateral behavior at moderate
angles of attack for the landing configuration.
The unsteadiness was apparently caused by an unsteady
flow off the wing root glove (the fixed, highly
swept inner wing). In early 1965, extensive flow
visualization tests were made in the Langley 12-Foot
Low-Speed Tunnel in an effort to change the vortex-flow
field set up by the glove and to delay separation
on the inner wing for the landing configuration.
This work included an investigation of a rotating
glove vane, which was then evaluated during the
free-flight model test in the Full-Scale Tunnel.
The glove vane cured the roll unsteadiness previously
noted and was subsequently incorporated into the
F-111 landing configuration.
Time-lapse photograph of the
F-111A free-flight model at several
wings sweep angles during tests in the Langley
Full-Scale Tunnel.
As F-111 operations expanded within the Air Force
in the late 1960’s, a rash of incidents involving
unexpected departures from controlled flight during
maneuvers at high angles of attack occurred. The
Air Force requested industry and NASA assistance
in analyzing and solving the problem, which was
viewed as a significant flight safety issue. Langley
researchers Joseph R. Chambers and James S. Bowman,
Jr. served on an Air Force, industry, and NASA
committee that identified a shortcoming in the
F-111 flight control system that promoted the unintentional
departures. The F-111 had been designed with a
g-command flight control system that provided
g-forces in direct proportion to the deflection
of the pilot control stick. However, in providing
the pilot with the level of g-force, the
system would increase the angle of attack of the
aircraft. Unless the pilot was monitoring the angle
of attack, the aircraft could enter a range of
high angles of attack where a loss of directional
stability resulted in an unintentional yaw departure
and spin entry. These findings led to an Air Force
program in 1973 to develop a stall inhibitor system
(SIS) for the F-111. Langley participated in the
design and analysis of this system. The SIS was
designed to automatically monitor and limit the
angle of attack of the aircraft during flight maneuvers
and was incorporated into the F-111 fleet.
|
Spin Recovery |
|
Another area of
controversy within the F-111 development program
arose in the early 1960’s—in the area
of spin and spin recovery technology. The conventional
approach to assessing and improving spin characteristics
of new military aircraft was to conduct tests of
dynamically scaled models in the Langley 20-Foot
Vertical Spin Tunnel to determine spin and spin
recovery characteristics, as well as the size of
the emergency parachute required for spin test
aircraft. In conjunction with these tests, radio-controlled
drop-model tests were conducted to assess spin-entry
tendencies and to assess the effectiveness of out-of-control
recovery procedures.
For the F-111 program, General Dynamics proposed
to assess spin recovery characteristics with analytical
methods, in lieu of the traditional dynamic scale
model tests. In addition, General Dynamics requested
over 1,000 hr in Langley tunnels to provide the
data required for the proposed study. Langley’s
reaction to this proposal was extremely negative
because analytical procedures for spin studies
had not been validated, and in the opinion of James
Bowman and the Spin Tunnel staff, could not be
trusted for such an important aircraft program.
However, the Air Force accepted Langley’s
recommendation to terminate the General Dynamics
plans for the analytical approach and elected to
continue with the traditional Spin Tunnel and drop-model
tests.
At the request of Langley, a 1/24-scale model
of the F-111 was tested in the Ames 12-Foot Pressure
Tunnel in early 1964 to examine Reynolds number
effects at high angles of attack and sideslip.
The results of these tests necessitated forebody
modifications for the spin tunnel model to simulate,
at the low Reynolds numbers of spin tunnel investigations,
the cross-flow characteristics approximating the
Reynolds number conditions of full-scale flight.
During the period from October 1964 to May 1966,
extensive spin tunnel tests were made by James
Bowman and Louis White on 1/40-scale models of
the F-111A and F-111B aircraft to determine spin
and recovery characteristics. Tests were conducted
for wing-sweep angles of 20 deg, 26 deg, 50 deg,
and 72.5 deg. The results of the tunnel tests indicated
that the F-111 would exhibit several spin modes,
including steep oscillatory spins from which recovery
could be accomplished and a fast flat spin from
which recovery was marginal or impossible using
aerodynamic controls. The flat spin was especially
stable. A number of radical approaches to breaking
the spin were attempted, including sweeping the
wings forward and rearward during the spin. For
these tests, the model was equipped with a small
electric motor that was remotely actuated to drive
the wing-sweep angle. However, the flat spin could
not be slowed or stopped using this technique.
All of these spins (including the flat spin) were
subsequently encountered in F-111 spin tests at
Edwards Air Force Base and in fleet operations.
Bowman served on several spin accident investigation
committees formed by the Air Force.
Extensive studies were made of the spin entry
and post-stall motions of the F-111A by Charles
E. Libbey with two 1/9-scale helicopter drop models.
Over 50 successful drops were made with wing sweeps
varying from 16 deg to 72.5 deg. Results of these
tests showed that certain pilot inputs following
the yaw departure at high angles of attack (as
previously discussed for the wind-tunnel free-flight
model) would promote the fast flat spin. These
results were discussed with Air Force representatives
for inclusion in the pilot handbook procedures
for avoiding this extremely dangerous condition.
Several F-111 aircraft were lost in spin accidents
during fleet operations; however, the subsequent
implementation of the SIS prevented stalls and
eliminated spins as an operational concern.
During the late 1960’s, Langley researcher
William P. Gilbert conducted fundamental research
on automatic spin prevention systems for fighter
aircraft. Gilbert’s work was stimulated by
the fact that flight control systems were beginning
to use flight parameters (such as angle of attack
and yaw rate) that would permit the mechanization
of spin prevention for routine operations in highly
redundant systems. Previously, the concept of automatic
spin prevention was a highly desirable concept,
but the mechanization would have required a special
system that might be prone to failure and would
only be utilized on very rare occasions. With the
emergence of the new operational control system
components, Gilbert became interested in demonstrating
the effectiveness of spin prevention systems. Following
a series of analytical studies, Gilbert teamed
with Charles Libbey to implement and test the prototype
system on an F-111 drop model. With the system
engaged, the spin-prone model could be maneuvered
to extreme angles of attack without entering a
spin, even with full prospin control inputs from
the pilot.
The F-111A drop model prior to
launch from a helicopter for a spin-entry test.
On one occasion, Gilbert briefed NASA Administrator,
Dr. James Fletcher, on the very positive results
of his study. Fletcher praised the work as a forerunner
of future systems that would enable carefree maneuvering
of military aircraft. Gilbert’s work with
the F-111 model represented one of the first efforts
to develop the highly sophisticated control systems
that are now used in virtually every domestic and
foreign fighter aircraft to prevent spins during
strenuous air combat maneuvers.
|
Flutter Tests |
|
In early 1963, flutter
trend models were tested in the Langley 26-Inch
Transonic Blowdown Tunnel to determine the general
flutter boundary for the isolated wing and tail
surfaces of the F-111. Several tests were conducted
in this facility, before a dummy (stability) model
of the complete F-111 configuration was tested
in the Langley 16-Foot Transonic Dynamics Tunnel
(TDT) in August 1963. The initial TDT tests were
to check out the wind-tunnel suspension system
for future flutter tests. Subsequent tests of the
isolated wing and horizontal tail in the Blowdown
Tunnel revealed that the horizontal-tail design
had an inadequate margin of safety for flutter
at low supersonic speeds, and the geometry of the
F-111 tail configuration was changed.
Flutter clearance tests of the F-111 empennage
model and a 1/8-scale complete flutter model of
the F-111 were made in the TDT during February
1965. Flutter tests continued through 1965 and
subsequent years to examine the effects of a number
of external store configurations on flutter boundary.
External stores for the F-111 included combinations
of bombs, missiles, and fuel tanks. The wing pylons
pivoted as the wings swept back, keeping the ordnance
parallel to the fuselage. F-111 tests led by Charles
L. Ruhlin, Maynard Sandford, and Irving Abel ultimately
included 13 test entries in the TDT of versions
such as the F-111A, F-111B, and FB-111.
F-111 model in the Langley 16-Foot
Transonic Dynamics Tunnel in 1963.
(Tail surfaces were later changed to avoid flutter.)
FB-111 flutter model illustrating
the wide scope of external stores that were tested
in the program.
|
Crew Escape Module |
|
The two crew members
in the F-111 sat side by side in an air conditioned,
pressurized cockpit module that served as an emergency
escape vehicle and a survival shelter on land or
in water. In emergencies, crew members remained
in the cockpit, an explosive cutting cord separated
the cockpit module from the aircraft, small rocket
engines ejected the module from the aircraft, and
the module descended by parachute. The ejected
module included a small portion of the inner wing
glove to stabilize it during aircraft separation.
Air bags cushioned the landing impact and helped
to keep the module afloat in water. After separation,
the air bags were inflated with nitrogen (stored
behind the pilot’s seat).
The module could be released at any speed or
altitude—even under water. For underwater
escape, the air bags raised the module to the surface
after it had been severed from the plane. Initial
concerns for the module design centered on potential
windblast; however, this threat was properly addressed
in the design. Unfortunately, impact of the level
of acceleration on the crew during landing proved
to be a problem. The nominal descent rate of the
module on the parachute was about 32 ft/sec.
As a result of a number of back injuries to crew
members during use of the recovery system and undesirable
postimpact overturning, the Air Force requested
that the NASA Crash Dynamics Group conduct drop
tests of the F-111 crew escape module at the unique
Langley Impact Dynamics Research Facility (IDRF).
The facility had initially been used to train astronauts
for moon walks as part of the Apollo Program, when
it was known as the Lunar Landing Training Facility.
After the successful Apollo Program, the Langley
staff recognized the value of the tall, large gantry
structure for simulating ground impact of large
aircraft structures and full-scale general aviation
aircraft and rotorcraft. The Air Force provided
F-111 crew escape modules and air bags for the
tests that were conducted at the facility.
Huey D. Carden and Lisa E. Jones led the F-111
module tests at Langley. The tests (between 60
to 70 drops) spanned from the early 1980’s
to 1995. The objectives of the tests were to assess
- Impact loads that were generated
by the air bag attenuation system
- Structural loads that were
transmitted to the seats and occupants
- Module stability under various
impact attitudes which could exist due to drift
from winds
- Design changes to the air
bag attenuation system or seats.
Tests with controlled pitch, yaw, and roll orientations
of the module relative to the forward velocity
vector were conducted to account for various attitude
envelopes of the module during descent. Additional
vertical tests were also performed. Impact velocities,
structural impact loads, air bag pressures, and
loads transmitted to the seats and dummies representing
the crew were measured and provided to the Air
Force for assessment.
Additionally, impact and postimpact behavior
of the module was provided via extensive onboard
and ground-based cameras, which also provided module
stability information. The last series of drop
tests in 1995 assessed the performance of and qualified
an entirely new air bag design that was required
for a new parachute design.
The results of the Langley tests were analyzed
and provided to the Air Force and the industry
contractor for continual refinement of the system.
As a result of the data provided to the Air Force,
load attenuating crew seats were included in the
F-111 and air bag and blowout plug design changes
were made in the original air bags. Various changes
to the module led to the design of a new air bag
system, which was tested for qualification on the
F-111 in the final series of tests prior to the
retirement of the U.S. F-111 fleet.
F-111 crew escape module during
tests at the Langley Impact Dynamics Research Facility. |
Wing Box Problems |
|
The F-111 airframe
utilized a significant amount of high-strength
D6ac steel in the wing carry-through structure.
This component was heat treated to a tensile strength
of 220,000 psi and designed for -3g to 7.33g
with design flight life goals of 4,000 hr and 10
years of service. However, a full-scale static
test program that was conducted over a 6-year period
encountered several failures, including a failure
at the wing-pivot fitting. Various modifications,
including the first use of an advanced boron-reinforced
composite doubler to reduce stress levels, coupled
with an extension of the structural tests to 40,000
hr, were believed to have provided for 10,000 hr
of safe operations.
In December 1969, an F-111 experienced a catastrophic
wing failure during a pull-up from a simulated
bombing run at Nellis Air Force Base. This aircraft
only had about 100 hr of flight time when the wing
failed. The failure originated from a fatigue crack,
which had emanated from a sharp-edged forging defect
in the wing-pivot fitting. As a result of the accident,
the Air Force convened several special committees
to investigate the failure and recommend a recovery
program. James C. Newman, Jr. and Herbert F. Hardrath
represented Langley on the recovery team deliberations,
and along with Charles M. Hudson and Wolf Elber,
they conducted fatigue crack growth and fracture
tests on specimens made from the D6ac steel used
in the aircraft. These tests were conducted in
the Langley Fatigue and Fracture Laboratory under
conditions that simulated aircraft operations.
The original material had low fracture toughness
due to the heat-treatment process. The committee
recommended that every F-111 be subjected to a
low-temperature proof test. This proof-test concept
had been developed and successfully used in the
Apollo program, as well as other missile and space
efforts. To screen out the smallest possible flaw
size, the F-111 full-scale proof tests were conducted
at temperatures of about -40¾ F, where the fracture
toughness of the D6ac steel was lower than the
fracture toughness at room temperature. The heat-treatment
process was also corrected to provide improved
toughness for the D6ac material in newer aircraft.
A decade later, the same material with improved
toughness was also successfully used in the Space
Shuttle solid rocket boosters. As a result of the
revised proof-test approach and the improved toughness
material, there were no F-111 aircraft lost due
to structural failure in almost 30 years of operations
before the aircraft was retired from service in
1996.
The F-111 failure was most responsible for the
U.S. Air Force developing the damage-tolerant design
concept, where flaws, such as a 0.05-in. crack,
are assumed to exist in critical aircraft components.
The structural components must then be tolerant
of these defects during flight conditions. This
concept relies on fatigue crack growth and fracture
criteria to establish an inspection interval to
insure the safety and reliability of the aircraft.
|
The Transonic Aircraft
Technology (TACT) Program |
|
Richard Whitcomb’s
pioneering research and development efforts on
supercritical airfoils for enhanced transonic performance,
which began at Langley in 1964 and continued until
the 1980’s, included extensive wind-tunnel
and flight evaluations for potential military applications.
After flight tests of a modified F-8 Crusader validated
the benefits of supercritical wing technology that
had been predicted by theory and wind-tunnel experiments
for potential civil applications, NASA and the
Air Force became interested in assessing the benefits
of supercritical wing applications to high-performance
aircraft during transonic maneuvers. Significant
increases in the drag-divergence Mach number, the
maximum lift coefficient for buffet onset, and
the Mach number for buffet onset at a given lift
coefficient were demonstrated for the supercritical
airfoil when compared with a NACA 6-series airfoil
of comparable thickness. Theodore Ayers, in cooperation
with General Dynamics, conducted exploratory tests
in the Langley 8-Foot Transonic Pressure Tunnel
in 1966 to test the effects of a slotted supercritical
airfoil on a 1/15-scale model of the F-111 with
the existing flap system. Although the overall
results were not satisfactory because of high subcritical
drag levels, the results encouraged additional
studies of an integral or unslotted supercritical
airfoil. Tests of a 1/24-scale F-111 model showed
significant benefits to drag-divergence Mach number,
maneuver drag, and buffet onset characteristics.
These tests also spurred additional interest by
General Dynamics in potential improvements of the
F-111.
Following these exploratory tests at Langley,
interest in supercritical applications continued
to increase and a joint NASA and Air Force study
that included ground and flight activities was
proposed. Specific objectives of the study included
the effects of supercritical wings on transonic
drag, buffet onset and magnitude, and handling
qualities. In a study known as the Transonic Aircraft
Technology (TACT) Program, several candidate military
aircraft were examined for potential modifications
to provide flight validation of supercritical wing
military applications.
The program was ultimately based on the application
of supercritical technology to the F-111 configuration,
which had outer wing panels that could be relatively
easily replaced with modified supercritical sections.
In addition, the potential retrofit of advanced
wings to enhance performance of the F-111 fleet
was an interest in some areas of the Air Force.
The TACT Program, which started in 1969, included
NASA Langley, Dryden, and Ames Research Centers,
and the Air Force Flight Dynamics Laboratory. Theodore
Ayers served as the Langley focal point for TACT
activities, which included extensive experimental
development work of the modified wing in the 8-Foot
Transonic Pressure Tunnel by Ayers and James B.
Hallissy (with considerable oversight and participation
from Dr. Whitcomb), tests of a propulsion model
in 1973 in the 16-Foot Transonic Tunnel by Charles
Mercer, and flutter tests led by Charles Ruhlin
and Maynard Sandford in 1971 in the 16-Foot Transonic
Dynamics Tunnel.
James B. Hallissy inspects the
F-111 TACT model in the Langley 8-Foot Transonic
Pressure Tunnel.
F-111 TACT model in the Langley
Unitary Plan Wind Tunnel for supersonic tests.
F-111 TACT model mounted for
flutter tests in the Langley 16-Foot Transonic
Dynamics Tunnel in 1971.
The NASA and Air Force F-111
TACT aircraft during flight tests at Dryden.
The F-111 TACT aircraft began flight tests at
Dryden in 1972. Flight-test results showed that
the supercritical wing generated up to 50 percent
more lift during maneuvers than the conventional
F-111 wing and significantly delayed the onset
of wing buffet to higher angles of attack. Special
flight tests were also conducted to demonstrate
that the carriage of external stores on the wing
pylons did not significantly degrade the benefits
of the supercritical wing. Although the Air Force
decided not to retrofit the F-111 fleet, the supercritical
wing technology had dramatically demonstrated its
benefits for incorporation into future military
aircraft.
|
The Mission Adaptive
Wing (MAW) Program |
|
In 1976, Theodore
Ayers transferred from Langley to Dryden, where
he accepted a position as Director for Aerodynamics.
Ayers continued his interest in advanced wing configurations.
Working with his technical peers within NASA, DOD,
and industry, he advocated for another important
flight program with the modified F-111 at Dryden.
After the TACT Program ended in the 1980’s,
the Air Force and NASA engaged in a new technology
development program known as Advanced Fighter Technology
Integration (AFTI). One element of this joint program
was the modification and flight tests of the F-111
TACT aircraft with a Mission Adaptive Wing (MAW).
The MAW used flexible wing skin and internal hydraulic
control mechanisms to recontour the wing shape
as a smooth variable-camber wing for varying flight
conditions. The objective was to provide the technical
confidence for significant performance improvements
with a wing system that varied the wing contour
in flight as a function of pilot inputs, flight
conditions, and structural loads. The wing box
of the existing TACT aircraft was equipped with
flexible wing leading and trailing edges. A high-lift
section was used for low-speed landing conditions,
and the wing was recontoured to a supercritical
shape for transonic flight and adjusted to a symmetrical
section for supersonic flight. Boeing, Grumman,
and General Dynamics bid on the request for proposals
for the aircraft modification. Tests of the competing
configurations in the 8-Foot Transonic Pressure
Tunnel were conducted by James Hallissy. Boeing
was awarded the contract in 1979. The MAW Program
was managed by the Flight Dynamics Laboratory of
the Air Force Wright Aeronautical Laboratories
with Dryden as the responsible flight-test organization.
Langley supported this activity with exploratory
tests of smooth variable-camber concepts in the
8-Foot Transonic Pressure Tunnel, which were conducted
by James C. Ferris. Langley also actively participated
in the flight-test program, including conducting
additional tests with a 1/24-scale model. Langley
personnel were also assigned temporarily at Dryden
during the flight program. The F-111 MAW flight
research was conducted from 1985 to 1988 and included
an assessment of automatic camber modes.
The F-111 Mission Adaptive Wing
(MAW) aircraft during flight tests at Dryden.
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