The F-16XL’s primary flight control system was a four-channel, analog-com-puter-based fly-by-wire flight control system. It was similar to the FBW flight control system used in the standard F-16, but it was modified to function with the XL’s unique flight control surfaces.20 Changes were also made to the air data system that determined the various parameters needed by the flight control system, such as static and dynamic air pressures and angles of attack and yaw. The flight control computers were modified to incorporate new control laws neces-sitated by the radically different aerodynamic configuration. The air data sensor installation was modified from that used on the standard F-16. A third conical angle-of-attack transmitter was added on the left side of the forward fuselage with an L-shaped probe on the right side providing the third source of total and static air pressure needed by the four independent flight control computers.
The cockpits installed in both F-16XL prototypes were configured with the same flight controls, displays, avionics, and sensors used in early production F-16A and F-16B aircraft. However, a significant new feature was installed in the F-16XL; this was the microprocessor-based self-test function for the flight control system. Following engine start, a button push activated the self-test function that evaluated the functionality of the entire fly-by-wire flight control system. This resulted in the complete flight control system, from controller inputs to rate sensors to actual servoactuator operations, being automatically checked—a process that took about 90 seconds. During the self-testing process, the pilot could detect the chatter of the flight control surface movements over the nose of the idling engine as the system went through its preprogrammed series of very-high-frequency control surface actuations.21
On the F-16XL, aircraft pitch attitude was controlled by symmetric deflec-tion of the two elevons mounted on the trailing edges of each inner wing segment. Elevon area was about 44 square feet. The elevons were capable of deflection over the angular range between 30 degrees up and 30 degrees
down. Additional pitch control was provided by symmetrically deflecting the outboard ailerons. The ailerons had a total surface area of 29.45 square feet.
Each aileron was capable of movement through an arc that ranged from 20 degrees up to 30 degrees down. Roll was controlled by asymmetric deflection of the ailerons assisted by elevon deflection. The roll axis of the flight control system was based on roll rate command and used roll rate feedback. The maximum available roll rate was 308 degrees per second. The actual allow-able roll rate was scheduled as a function of aircraft angle of attack to prevent overstressing the aircraft or encountering potential out of control situations.
The rudder provided yaw control. It was interconnected to the ailerons via the FBW flight control system. This ensured smooth aileron-rudder coordi-nation during rolling maneuvers, especially at higher angles of attack. The yaw axis of the flight control system used rudder position commands based on yaw rate and lateral acceleration feedback along with inputs from the aileron-rudder interconnect.22
Secondary flight control surfaces included the leading-edge flaps, installed on the outer wing panels, and standard F-16 speed brakes (though recon-toured on their outer mold-line for use on the XL) were mounted one on each side of the aft fuselage adjacent to the engine exhaust nozzle. Total surface area of the leading-edge flaps was about 18.2 square feet. The LEFs were capable of being deflected through an arc ranging from 6.4 degrees up to 36.5 degrees down. The leading-edge flaps were automatically scheduled by the flight con-trol system to optimize maneuvering performance and assist the elevons in providing adequate pitch control at higher speeds and dynamic pressures.
The leading-edge flaps were automatically deflected upward at higher Mach numbers to minimize aerodynamic drag. In addition, asymmetric leading-edge-flap deflection was used to assist the rudder and ailerons in ensuring
lateral-directional stability at higher angles of attack.
Aircraft spin recovery was automatically assisted by the FBW flight control system, which differen-tially deflected the leading-edge flaps when the angle of attack was greater than 35 degrees. The differen-tial deflection capability of the LEFs ranged up to an angle of 40 degrees.
Two clamshell-type speed A combination of flight control surfaces provided pitch,
roll, and yaw control across the entire F-16XL flight envelope. (NASA)
brakes were located on each side of the engine nozzle. They functioned simul-taneously when actuated by the pilot and opened symmetrically, with each segment deflecting 60 degrees.23
During flight operations, the pilot commanded normal load factor (air-craft g-level) throughout most of the flight envelope using the standard F-16 sidestick controller. The pitch axis of the FCS was essentially a g-command system that utilized g-level, AoA, and pitch rate as feedback. During the powered landing approach
configura-tion, with landing gear and flaps down, and at angles of attack above 19 degrees in the cruise configuration, the FBW flight control system automatically pro-vided blended g-level (load factor) and AoA commands to ensure speed stability.
An AoA limiter was incorporated in the F-16XL flight control system to mini-mize opportunities for out-of-control situations during maneuvering flight.
This limiter was similar to that used in the standard F-16A, but the allow-able AoA envelope was expanded to 29 degrees at low speeds (up from the
25-degree limit in the basic F-16). At speeds above Mach 0.9, the allowable AoA was restricted to a maximum of 26 degrees. The allowable angle of attack when the landing gear was extended was automatically limited to 16 degrees.
In the event of a departure from controlled flight, the leading-edge flaps were automatically used to augment the rudder in limiting high yaw rates from developing at angles of attack between 35 and 50 degrees. If a stabilized deep stall condition was encountered, the flight control system featured a manual pitch override (MPO) capability. MPO allowed the pilot to bypass the AoA limiter, providing him or her with full pitch command authority to initiate a pitch-rocking maneuver and break the deep stall. Pitch-rocking was success-fully used on the F-16 to break deep stalls, and it also was fitted with an MPO capability.24 NASA Langley had played a significant role in the development of the F-16 deep stall recovery system. The flight control systems installed in the two F-16XL prototypes were fitted with an automatic pitch override system intended to break any deep stall that might be encountered. This automated system provided pitch-rocking commands to the flight control system in the event of a stabilized deep stall.25 F-16XL flight control system functionality and effectiveness were evaluated during extensive wind tunnel, spin tunnel, and drop model test efforts at NASA Langley.
The F-16XL used the force-displacement sidestick controller adopted from the standard F-16. (NASA)
Avionics
Both F-16XL flight demonstrator aircraft were equipped with standard F-16A avionics suites. If the F-16XL had been selected for full-scale development as the USAF dual-role fighter, GD planned on incorporating the avionics suite developed under the F-16C/D Multi-Stage Improvement Program (MSIP) into production aircraft. These aircraft would have been equipped with an improved APG-66 fire control radar; the AN/ALR-74 threat warning system;
the LANTIRN auto navigation, terrain avoidance, and targeting system; an airborne self-protection jammer; the ALE-40 chaff/flare dispensing system;
both an inertial navigation system (INS) and a global positioning system (GPS); and MIL-STD-1760 data interface capability with advanced weapons, including the AIM-120 AMRAAM air-to-air guided missile. In addition, an Identification, Friend or Foe (IFF) subsystem with air-to-air IFF capability would have been provided as would provision for HAVE QUICK secure ultra-high frequency (UHF) communications capability. When fully implemented, the MSIP would have provided night, under-the-weather, navigation, and The physical and geometric properties of the F-6XL aircraft are illustrated in this NASA drawing. The effects of conical camber and twist on the wing are evident in the front view of the aircraft. (NASA)
air-to-ground weapon delivery and beyond-visual-range (BVR) air-to-air mis-sile engagement capabilities. The front cockpit would have retained the features and capabilities of the F-16C. The LANTIRN wide-angle, improved-video Head-Up Display (HUD) incorporated a forward-looking infrared (FLIR) navigation video display at night. Two 4-inch high-resolution multifunction displays (MFDs) would have been available for both sensor and weapon-video and interactive control of the weapon system. An optional 5-inch color MFD was available for integration into the center console. This display was to provide moving presentations with overlaid navigation information, color-coded elec-tronic flight instruments, and other color-coded aircraft pilot-selectable infor-mation. Integrated upfront Communications, Navigation, and Identification (CNI) controls were to be provided along with a dedicated threat warning system. The rear cockpit in the dual-role fighter version of the F-16XL was to have the controls and displays, including a color-moving map, to accomplish a full range of day/night/all-weather missions.26