The F-16XL zero-dihedral cranked-arrow wing had an overall span of 32 feet 5 inches. When the wingtip-mounted AIM-9 missiles and their launchers were installed, wingspan was increased to 34 feet 3 inches. Wing area was a little over 663 square feet compared to 300 square feet for the F-16’s trapezoidal wing. The aspect ratio of the wing was a very low 1.6 as compared to the F-16 aspect ratio of 3.0. Leading-edge sweep angles were 70 degrees on the inboard segments of the wing and 50 degrees on the outboard segments. The long-chord wing had a thickness-to-chord ratio of 4.5 percent. The mean aerodynamic chord of the cranked-arrow wing was 24.7 feet.3
A modified National Advisory Committee for Aeronautics (NACA) 64A airfoil was used on the inboard portion of the cranked-arrow wing. This por-tion of the wing had its camber conically tailored from the root outward. This feature was designed to reduce drag at transonic and supersonic speeds and improve the acceleration characteristics of the aircraft. The outer wing panels used a modified biconvex airfoil. The angle of incidence of the wing was –0.65 degrees over the inboard wing segment and then varied as the wingtip was approached, reaching a negative incidence of 4.1 degrees at the tip.4 This coni-cal camber resulted in a distinctive twisted effect when the wing was viewed from the front. This can be seen in the three-view drawing of the aircraft.
Elevons were mounted on the aft inner wing segments. Ailerons and leading-edge flaps were installed on the outboard wing panels. Hydraulically operated aileron actuators were located in large pods on the aft portion of the wings between the ailerons and the elevons. These pods were also intended to house electronic warfare equipment including chaff and flare dispensers. The hydraulically oper-ated elevon actuators were contained inside the wing structure close to the root on each side. The leading-edge-flap rotary actuators—electrically commanded but hydraulically operated—were mounted on the forward spars of the outer wing panels.5 Further details concerning the F-16XL’s elevon, flap, and aileron design and their functional modes of operation are contained in the flight control section.
Computer-aided design (CAD) and computer-aided manufacturing (CAM) systems were extensively used in both the design and manufacture of the F-16XL. The use of computer-based structural analysis, design, and manufacturing systems in aerospace applications was greatly influenced by a major NASA initiative known as the NASA Structural Analysis System (or NASTRAN for short). NASTRAN had been developed under NASA Langley Research Center management leadership. NASTRAN led to the development and use of many commercially developed CAD/CAM systems, with various tailored capabilities, that were subsequently used in the aerospace industry and many other design and manufacturing industries.6
General Dynamics purchased a CAD/CAM system that had been devel-oped by Lockheed called CADAM. It used an IBM mainframe computer that served interconnected CADAM terminals. During engineering design of the F-16XL, it was quickly found that these terminals had to be carefully prioritized and scheduled because there were not enough of them to go around for every member of the engineering team to have full-time access. In fact, the fuselage parts were drawn mostly on paper for this reason. The CADAM system played a major role in the design, manufacture, and smooth integration of the complex cranked-arrow wing into the prototypes. The two segments of the F-16XL’s cranked-arrow wing were designed with new wing attachment fittings. These fittings mated with the stretched fuselage at the existing F-16 wing attachment points and at additional attachment points located on the aft fuselage exten-sion. The greater physical depth of this wing over the standard F-16 wing led to the need to “invert” the wing attachment fittings by introducing an angle into the inboard portion of the wing skins.7 Later, during stability and control flight tests involving very high roll rate maneuvers, structural loads instrumentation indicated that the stress in the wing-fuselage shear tie at Fuselage Station (FS) 463.1 could exceed the flight-test limit.8 This led to the design and installation of a reinforced shear tie in both aircraft.
The robust internal structure of the F-16XL wing was designed to handle the stresses associated with aerodynamic loads including loads created during operation of the independently actuated elevons, ailerons, and leading-edge flaps. Loads on the wing hard points produced when external stores were carried and forcefully ejected from the aircraft also had to be accommo-dated in the wing structural design.
Not long after the July 1982 first flight of the F-16XL-1, data from the inflight structural loads instru-mentation system led structural engineers to conclude that the aft wing spar could fail at 86 percent of its design load limit. To deal with this issue, GD engineers developed a structural fix that resulted in the replacement of a 22-inch-long sec-tion of the aluminum rear spar with a strengthened part made of high-strength steel. The new steel part also incor-porated a flange with increased depth for even further strength. The wing on F-16XL-1 was modified with the new component at Fort Worth. Inflight stress The robust internal structure of the F-16XL’s
cranked-arrow wing viewed looking out-board from the wing root toward the wingtip.
(Lockheed Martin)
measurements validated the adequacy of the aft wing spar structural fix. The strengthened aft wing spar modification was incorporated into the wing on F-16XL-2 prior to its first flight at Fort Worth in October 1982.
Specially tailored upper and lower composite wing skins were developed for use on the F-16XL. These composite skins were mechanically attached to the aluminum understructure of each wing using special fasteners. The composite wing skins covered most of the surface area of each wing segment. The skin design was based on the use of aeroelastically tailored graphite-bismaleimide laminates (composed of T300 fiber with a V378-A matrix) that were intended to save structural weight. The F-16 had used a graphite-epoxy composite mate-rial (T300 fiber/3501-6 matrix), but GD moved to the polyimide-based family (of which bismaleimide is a subset) for their superior high-temperature prop-erties. The skins were tailored
to deform favorably under load while meeting the strength requirements of a minimum weight, damage-tolerant struc-ture. They ranged in thickness from 0.25 to 0.75 inches. By the time of the F-16XL devel-opment, graphite-polyimide composites were being increas-ingly used in aerospace appli-cations, including on the Space Shuttle. The development and widespread use of graphite-polyimide composite struc-tures in the aerospace industry
had been heavily influenced by NASA-sponsored research conducted with many companies including General Dynamics. In addition, GD had more than a decade of development experience with composite materials, sponsored in large part by the USAF Materials Laboratory at Wright-Patterson AFB.
The graphite-bismaleimide composite material had excellent mechanical strength, being able to handle stresses of up to 50,000 pounds per square inch (psi), as well as good heat-resistance properties. For the F-16XL application, the composite wing skins were built up of laminated layers of graphite-bisma-leimide material in which its thickness consisted of anywhere from 30 to 166 plies. The individual layers used to build up the laminates had been laid-up over a mold by hand as automatic tape-laying equipment was still in develop-ment. The specific number of plies used in an individual laminated layer was dependent on the highest computed local structural load and stress levels likely
Composite wing skins increased structural strength, saved weight, reduced manufacturing cost, and improved durability. (Lockheed Martin)
to be encountered across the F-16XL’s aerodynamic flight envelope. During assembly of the skins, the laminates were carefully laid down in a precisely planned geometric orientation intended to optimize the distribution of stresses across the wing skins to reduce structural weight. The process had to be done in a time-critical way, as the bismaleimide resin in the graphite tape would slowly begin to cure as it warmed up after removal from refrigerated storage.
Following the room temperature lay-down process, the skins were placed on a mold that replicated the required curvature of the outer wing surfaces. The skins were then cured under high pressures and high temperatures in a special autoclave. This resulted in the formation of permanently hardened graphite-bismaleimide composite wing skins with the proper compound curvature.
Following final manufacturing details, such as insuring an exact fit and drilling the fastener mounting holes, the skins were ready for installation onto the basic metallic wing substructure.9
The graphite-bismaleimide composite wing skins not only saved weight (a savings GD claimed was 595 pounds), but they also reduced manufacturing Graphite-bismaleimide laminates were used in various combinations and orientations to build up the composite wing skins. (Lockheed Martin)
and production costs. This was due to the fact that the composite materials used to build up the wing skins were more adaptable to being formed into the com-pound curvature surfaces that were features of the F-16XL arrow-wing design.
The compound curvature of the wing skin surface was due to its complex combination of wing camber and twist. Composite wing skins did not require machine milling or chemical etching as metallic skins did. Other benefits were increased stiffness and better rigidity compared to metallic counterparts and an anticipated 2.5 times improvement in durability in operational service.10