Highlights of Research by Langley for the F-111

1. conducted studies to enable the variable-sweep wing that was employed by the F-111 and subsequent variable-sweep aircraft such as the F-14 and B-1.
2. staff conducted wind-tunnel tests, identified major problems, and recommended critical improvements to the aerodynamic performance of the F-111.
3. personnel presented key technical results in congressional and high level Department of Defense investigations of the technical merit of the F-111.
4. led studies to provide solutions to major propulsion integration problems exhibited by the F-111, which were used on all subsequent high-performance fighters, such as the F-14, F-15, and F-22.
5. identified critical deficiencies in the high-angle-of-attack characteristics of the F-111.
6. detected the presence of a dangerous flat spin for the F-111 and precipitated the development of a stall inhibitor system to protect against inadvertent spins.
7. Following a catastrophic wing fatigue failure of an operational F-111, Langley conducted critical materials tests, and Langley personnel served on advisory boards to recommend approaches to ensure safety of the fleet.
8. conducted tests to improve the robustness and landing loads of the F-111 crew module.
9. wind-tunnel tests to determine the flutter characteristics of the F-111 identified a potential tail flutter problem that resulted in modifications to the aircraft.
10. supercritical wing technology was applied to the F-111 and demonstrated during the TACT and MAW Programs.

The General Dynamics F-111 was a multipurpose supersonic tactical fighter-bomber aircraft. The F-111 was one of the more controversial aircraft in the U.S. military inventory, yet it achieved one of the safest operational records in Air Force history. It was regarded as a highly effective, all-weather interdiction aircraft. As the result of a high level government directive in 1961, both the Navy and Air Force became committed to a common Tactical Fighter Experimental (TFX) Program. The TFX Program called for developing a single aircraft to fulfill a Navy fleet defense interceptor requirement and an Air Force supersonic strike aircraft requirement. The mission requirements were impossible to achieve, especially since planners placed priority upon the Air Force requirement, and then tried to tailor a heavy land-based aircraft to the demands of carrier-based naval aircraft. The naval version, the F-111B, was never placed into production. The Air Force aircraft was produced in a variety of models, including the F-111A, F-111D, F-111E, and F-111F fighter-bombers; the FB-111A strategic bomber; the F-111C for the Australian Air Force; and an EF-111 electronic warfare version. The U.S. Air Force versions were retired in 1996, but the Australians plan to operate their fleet until well into the twenty-first century.

Arguably, the political and technical issues associated with the F-111 program resulted in more research activities at the Langley Research Center than any other production military aircraft in Langley’s history. The staff participated in top level assessments of the aircraft’s capabilities (including briefings at the highest levels of the Department of Defense (DOD) and government), identified several critical problems, and provided recommendations for solutions. Significant Langley technical contributions included the variable-sweep wing concept, and test activities in aerodynamic performance enhancements, high-angle-of-attack characteristics, spin recovery, flutter, propulsion integration, wing structures and materials, and robustness of the crew escape module. Finally, Langley was a major participant in the joint NASA and Air Force F-111 Transonic Aircraft Technology (TACT) Program and the Mission Adaptive Wing (MAW) Program that applied supercritical wing technology, which was developed at Langley, through flight tests of a modified F-111 research aircraft at NASA Dryden Research Center.

Langley facilities that contributed to the F-111 program included the 7- by 10-Foot High-Speed Tunnel, the 4- by 4-Foot Supersonic Pressure Tunnel, the Unitary Plan Wind Tunnel, the 16-Foot Transonic Tunnel, the 8-Foot Transonic Pressure Tunnel, the 26-Inch Transonic Blowdown Tunnel, the 30- by 60-Foot (Full-Scale) Tunnel, the 20-Foot Vertical Spin Tunnel, the 16-Foot Transonic Dynamics Tunnel, the Fatigue and Fracture Laboratory, the Impact Dynamics Research Facility, the 12-Foot Low-Speed Tunnel, and radio-controlled drop models.

Langley Contributions to the F-111The Variable-Sweep Wing Concept
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.

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.

The Tactical Fighter Experimental (TFX) Program
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). - see books page

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.

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
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.

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.

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.

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.

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.

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.

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.

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