ESTUDIO DE LA LOCALIZACIÓN DE LOS ELEMENTOS MINEROS

Evaluation of the General Electric F110 Slimline engine was a significant aspect of the F-16XL-2 flight-test program. Airframe, instrumentation, and control modifications were accomplished to enable the aircraft to accommodate the GE F110-GE-100 Slimline engine. These modifications were completed at the GD factory in Fort Worth in mid-July 1984. After two functional check flights from Carswell AFB, F-16XL-2 was flown to Edwards on July 20 to begin the F110 Slimline engine evaluation. Initial testing primarily involved An unguided AGM-65 Maverick missile was launched from F-16XL-1 store station 3 on March 29, 1985. Weapons carried on this flight included four dummy AIM-120s, two AIM-9Ls, and six AGM-65 missiles. (USAF)

the evaluation of engine transient and air-start capabilities. These objectives were accomplished during F-16XL-2 flights 276 through 281. During flight number 282 on August 17, 1984, telemetered data received in the flight-test control room indicated that an engine compressor stall was imminent. The aircraft was at Mach 0.95 at an altitude of 30,000 feet. GD Project Pilot John Fergione had just switched from the F110 primary engine control mode to the secondary (SEC) control mode. The secondary engine control mode was intended to provide the pilot with a limited “get home” capability in the event of problems with the digital engine control system. When operating in the sec-ondary mode, F110 engine thrust was limited to 70 percent of its full military thrust and afterburner operation was precluded.⁵⁰ When Fergione switched over to the secondary mode, the F110 fan duct pressure sensor (known as Pressure Sensor 14, or PS14) had registered an extremely low reading. The engine experts in the test control room believed that the low fan duct pressure reading signaled that an engine stall had occurred.⁵¹ Discussions between the control room and the pilot resulted in a decision to abort the mission. Fergione landed without incident on Lakebed Runway 15 at Edwards AFB. Subsequent investigation determined that the fan duct pressure sensor instrumentation had registered an erroneous value and actual engine operation was in fact normal.

Flight testing resumed on August 20, 1984, with an uneventful flight.

However, the next day, during XL-2 flight number 284, the control room received a telemetered engine stall warning indication. This was similar to the one previously encountered and was from the same pressure sensor. The pilot, again GD’s John Fergione, switched over to the secondary engine control mode and made a precautionary simulated flameout approach to a landing on Runway 22 at Edwards AFB.⁵² This time, a detailed investigation of the F110 engine revealed a crack in a pressure line. Engine operation was being affected even though no abnormal indications were seen by the pilot. In an attempt to resolve the problem, General Electric field representatives reinforced the PS14 sensor manifold fitting locations with an epoxy and fiberglass mixture. This reinforcement added additional mass to the manifold. Perhaps not surpris-ingly, ground testing of the unsophisticated quick fix showed that the natural vibration frequency of the manifold had been significantly changed. Increased manifold vibration amplitudes were now being encountered, causing the epoxy reinforcement to separate from the fan frame. The next attempt to resolve the issue used bolt-on manifold fittings along with a new “flex-line” manifold.

This approach appeared to work during ground tests. The necessary safety of flight paperwork that allowed F-16XL-2 flight testing to resume was signed by the AFFTC commander, Maj. Gen. Peter W. Odgers, on October 8, 1984.

The following day, a functional check flight was successful and F110 flight envelope expansion resumed.⁵³

Inflight Afterburner Spray Bar Failure

After the F110 test effort resumed, a program of successful air starts, engine/

inlet compatibility, and inverted flight demonstrations continued without any problems from F-16XL-2 flight 286 through flight 305. However, on November 20, 1984, during flight 306, the F-16 chase pilot reported seeing a large plume of flame extending about three airplane lengths behind the nozzle.

This happened during afterburner operation following a go-around from a simulated flameout landing. John Fergione immediately throttled back to idle and performed an uneventful emergency landing on the lakebed at Edwards.

Postflight inspection of the engine revealed that a broken afterburner spray bar had caused the long plume. The resulting abnormal flame pattern had burned through the engine nozzle in the seven to ten o’clock region, when viewed from behind the aircraft. There was also significant fire damage to the left speed brake. A subsequent investigation by General Electric determined that a high fuel-to-air ratio had caused excessive stress levels in the engine, especially in the afterburner section. The fuel-to-air ratio was reduced, and successful ground testing was followed by a spray bar evaluation flight on January 11, 1985.

Instrumented spray bar data was collected during flights 307 and 308. These data confirmed that spray bar stress levels were now within the acceptable toler-ance range established by GE. One additional F110 envelope expansion flight was conducted following the spray bar stress evaluation before the aircraft was ferried back to Fort Worth on January 28, 1985, for installation of the Large Normal Shock Inlet.⁵⁴

LNSI Testing

F-16XL-2 would be absent from the flight-test program from January 29 until June 25, 1985. During this extended period at the factory, an extensive set of modifications and structural inspections were completed, requiring a total of 105 work days. The Large Normal Shock Inlet was installed, along with additional test instrumentation. Other airframe structural modifications were required based on the results of the inspections. The F110 Slimline test engine that had been in the aircraft was returned to the General Electric factory at Evandale, OH, where it was upgraded to the latest configuration. A functional check flight with the LNSI installation and the upgraded F110 engine was conducted at Fort Worth on June 26, 1985. A second functional check flight on June 28 was followed by a ferry flight to Edwards AFB that afternoon. Once at Edwards, the new spray bar configuration was tested. This was followed by a series of combined structural loads and stability and control evaluations of the aircraft. These included both high-g and elevated-AoA maneuvers to assess the effects of the larger inlet on aircraft performance and handling qualities.

Instrumentation verified that structural loads on the airframe with the LNSI

installed were not significantly increased. Based on these results, the aircraft was cleared to its original maneuver enve-lope. During subsequent flight testing of F-16XL-2 with the LNSI, some airframe vibration was encountered. There was insufficient time remaining in the F-16XL flight-test pro-gram to address the airframe vibration issue associated with the LNSI. However, it was one of the recommendations for follow-on investigation.⁵⁵

On flight number 324, the F110 engine’s fuel-to-air ratio was increased to maximize engine performance with the

higher mass airflow rate available from the LNSI. The LNSI provided a mass flow rate of 270 pounds of air per second as compared to the 254 pounds per second for the standard inlet. The change in fuel-to-air ratio required another round of spray bar evaluation flights, which were successfully completed. The final F110 test effort consisted of completing a comparative evaluation of engine performance using both JP-4 and JP-8 jet fuels. Additional stability and control and performance testing included data gathered at low altitude on flight 358 over the Pacific Ocean off Vandenberg AFB, CA. After this flight, the engine was removed from the aircraft and replaced with the backup engine that had been held in reserve. Following a successful functional check flight at Edwards, XL-2 was ferried to General Dynamics in Fort Worth on October 1, 1985. Once there, it was placed in flyable storage along with XL-1. This was the final flight of the Air Force F-16XL flight-test program. It marked the 361st test flight for F-16XL-2. The aircraft had logged a total of 407.1 flight hours.⁵⁶ Pitch Gallop Investigation

Preliminary observation of what came to be called pitch gallop had been noted during the very first flight of F-16XL-1 on July 3, 1982. The Air Force flight-test evaluation had confirmed that pitch axis oscillations were generally encoun-tered in the Mach 0.90–0.95 region, becoming more pronounced as altitude decreased and g-level and angle of attack increased. This phenomenon had not been seen when flying the early version of the General Dynamics F-16XL F-16XL-2 fitted with the extended air data boom used to measure angle of attack, yaw angle, and static and dynamic pressures. (USAF Photograph courtesy of SMSgt. Norman Taylor, USAF [ret.])

flight simulator. Aircraft characteristics that were modeled in the simulator had been based on aerodynamic data derived from early wind tunnel testing and an analytical model of the F-16XL flight control computer (FCC). Actual flight testing revealed that there was a pronounced increase in F-16XL pitch surface control effectiveness in the high subsonic Mach number range. The magnitude of this pitch effectiveness increase was much greater than had been predicted from wind tunnel–derived data. GD modified the flight simulator using actual flight-test-derived aerodynamic data. However, they still had to incorporate an additional 30-degree phase lag to induce the simulator to create the gallop effect that was observed during actual flight testing. When aerodynamic data derived from actual flight testing was used in conjunction with an actual flight control computer and system hardware instead of the earlier analytical model, the simula-tor produced a longitudinal pitch oscillation. General characteristics of the pitch oscillation in the modified computer were similar to the pitch gallop encountered in flight testing. Subsequent investigation confirmed that the analytical model of the F-16XL used in the original development of the flight control computer did not accurately represent actual aircraft hardware. Significantly, the aircraft hardware models used in the computer simulation had undergone many inde-pendent verification and validation tests prior to and during the test program without the problem having been detected.⁵⁷

The Air Force investigated many possible causes for the pitch gallop phe-nomenon. Exterior aerodynamics and the pitch axis flight control system were found to be within design specifications. However, flight testing did determine that the autopilot had some effect on the severity of the pitch oscillation, espe-cially at higher altitudes. Changes to the pitch axis gain in the flight control computer were found to have the greatest effect. A notch-type filter was inserted in the pitch control path, and a multiplier of 0.50 or 0.75 was applied after the basic pitch gain was scheduled by the flight control computer. Both multipliers reduced the magnitude of the pitch gallop problem. A pitch gain reduction of 25 percent in the flight control computer was subsequently used as an interim fix for the remainder of the F-16XL flight-test program. However, the Air Force did not consider this interim approach to be a final solution to the pitch oscillation issue as the actual root cause of the problem had yet to be determined. Even the modified flight simulator did not fully replicate the observed pitch oscillation.

Therefore, the Flight Test Center decided that a dedicated flight-test effort was needed. This would be focused on determining the frequency response of the total F-16XL system. For this purpose, 11 flight-test missions were dedicated to the pitch gallop evaluation. Actual aircraft hardware and aerodynamics would be used with the pitch gain and phase angle in the flight control computer varied over a range of in-flight conditions. Data obtained using this approach would be used to refine the flight control system to eliminate the gallop problem.

Another goal was to improve the fidelity of the flight simulator at high subsonic Mach numbers.

To determine the system gain and phase angle of the total aircraft, an in-flight excitation system was developed and installed in F-16XL-1. The objective of the in-flight excitation tests was to obtain the system gain and phase angle at specific excitation frequencies and to identify any nonlinearities in gain or phase angle as a function of control surface position based on excitation amplitude. In the test approach developed for the F-16XL, the excitation signals were inserted into the pitch axis control loop, and the system response to pitch inputs from the pilot was measured. The excitation installation was intended to determine the actual FCS pitch channel gain and phase margins with various aircraft configurations and flight conditions. This in-flight excitation approach had been used during flutter and flight control testing with the F-16. In those earlier tests, excitation signals had been sent to the flight control surfaces in order to stimulate the structural modes of the aircraft in various regions of the flight envelope. The excitation system was pilot-activated through a modification to the autopilot panel. When activated, increased levels of aileron/elevon movement would be induced, along with various frequencies of control surface movement. Either a random or one of three sinusoidal control surface movement frequencies could be selected.⁵⁸ Maximum signal amplitude was limited to ±1.8-g commands to the pitch control surfaces.

Pitch oscillation investigations conducted during 1985 involved excitation testing throughout the F-16XL’s flight envelope. The major focus was on the high subsonic region, known to be the most troublesome area for pitch oscilla-tions. All excitation tests were used the 25 percent gain reduction in the flight control computer pitch path that had been adopted earlier as an interim fix to the pitch oscillation problem. Test results determined that the flight control system installed in the F-16XL prototypes had a high pitch gain with very close to a –180 degrees of phase angle in the high subsonic Mach number region. This was true at all altitudes flown during the excitation tests. A nonlinear effect on system gain as a function of control surface position was apparent from the test-ing. Ground excitation testing of the system was subsequently used to investigate the nonlinearly of the system in the absence of aerodynamic effects. These ground tests also confirmed the same nonlinearity in gain with neutral pitch trim. This nonlinearity in system gain was found to be a characteristic of the integrated servo-actuators used on the F-16XL prototype pitch control surfaces. The actua-tor nonlinearity was of sufficient magnitude to drive the aircraft system, consist-ing of aircraft and aerodynamics, into a limit-cycle oscillation. This was defined as a gain of 1.0 and a phase angle of –180 degrees in the 0.90 to 0.95 Mach range.

However, this limit cycle resulted in a stable oscillation that had no tendency to diverge or become unstable. This was due to the fact that as pitch control surface

deflection increased in magnitude, system gain got smaller, resulting in a stable system. The pitch oscillation test excitation procedure used on the F-16XL was successful in producing data that correctly defined the aircraft frequency response throughout the flight envelope. These data would have been used to refine the flight control system had the aircraft completed full-scale development.⁵⁹

Refining the F-16XL’s Flight Control