Document
AIAA-92-4106-CP
INTEGRATED FLIGHT PROPULSION CONTROL RESEARCH
RESULTS USING THE NASA F-15 HIDEC FLIGHT
RESEARCH FACILITY
James F. Stewart* NASA Dryden Flight Research Facility Edwards, California, USA Abstract FTIT fan turbine inlet temperature HIDEC highly integrated digital electronic Over the last two decades, NASA has conducted control several flight research experiments in integrated flight- propulsion control. Benefits have included increased heads-up display HUD thrust, range, and survivability; reduced fuel consump- INS inertial navigation system tion; and reduced maintenance. These flight programs IPCS integrated propulsion control system were flown at NASA Dryden Flight Research Facility.
This paper presents the basic concepts for control in- NASA National Aeronautics and Space tegration, examples of implementation, and benefits of Administration integrated flight propulsion control systems.
NPR nozzle pressure ratio The F-15 research involved integration of the en- ambient pressure, lb/in Pa.mb gine, flight, and inlet control systems. Further exten- PROTECT propulsive techniques for emergency sion of the integration included real-time, onboard op- control timization of engine, inlet, and flight control variables; PSC performance seeking control a self-repairing flight control system; and an engines- only control concept for emergency control. The flight {j dynamic pressure research programs and the resulting benefits are de- SIDC system impairment detection and scribed for the F-15 research.
classification Nomenclature SRFCS self-repairing flight control system thrust specific fuel consumption TSFC ADECS adaptive digital engine control system USAF United States Air Force CAS control augmentation system calibrated airspeed Vcas DEEC digital electronic engine control angle of attack, deg a DEFCS digital electronic flight control system angle of sideslip, deg /3 DFCC digital flight control computer EEL extended engine life Introduction EGE effective gain estimator The integration of propulsion control systems and EMD engine model derivative propulsion-flight control systems has been shown to EPR engine pressure ratio significantly improve airplane performance parameters FDIE such as thrust, range, and rate of climb. \Vhen sys- fault detection isolation and estimation tems are not integrated, each system must be able to operate in a worst-case combination with the other sys- 'F-15 HIDEC Project Manager Copyright @1992 by the American Institute of Aeronau- tems, and large operating margins are required. Inte- tics and Astronautics, Inc. No copyright is asserted in the gration allows these margins to be reduced when the United States under Title 17, U.S. Code. The U.S. Govern- full margins are not required, resulting in higher thrust, ment has a royalty-free license to exercise all rights under lower fuel flow or range, and better safety and reli- the copyright claimed herein for Governmental purposes. All other rights are reserved by the copyright owner. ability. Integration control laws are developed in an off-line process and stored in an onboard computer for high-performance fighter with excellent transonic ma- implementation. System performance was further im- neuverability and a maximum Mach capability of 2.5.
Two afterburning turbofan engines power the F-15, proved by real-time optimization used in place of the and it has a high-mounted swept-back wing, twin ver- a priori or preprogrammed optimization. The real- time approach is much more challenging to develop tical stabilizers, and large horiwntal stabilizers. The and implement. Because it can adapt to changing engine inlets are the two-dimensional external compres- flight conditions, however, the real-time approach may sion type with three ramps and feature variable capture achieve higher levels of performance. area. Figure 4 shows a three-view drawing of the F-15 aircraft. The airplane is almost 64 ft long and has a To study the problems of integration and to deter- wingspan of nearly 43 ft.
mine the benefits of integration in the actual flight envi- ronment, NASA Dryden has conducted flight research The configuration of the NASA F-15 enhances its over the past two decades. In the mid-1970's, propul- flexibility for research, since it does not have most of sion system digital control and control integration were the weapons systems equipment that is part of the stan- developed and demonstrated in the Integrated Propul- dard F-15 aircraft. Absent are the radar, gun, missiles, sion Control System (IPCS) Program (Ref. 1), a joint and weapons systems avionics. This provides a large United States Air Force (USAF) and NASA program volume of space for experiments and instrumentation.
flown on an F-lllE airplane. The flight demonstration Flight Control System (Fig. 1) clearly showed the benefits of digital control and control integration. The standard F-15 airplane is equipped with a me- chanical flight control system that provides control of In the late 1970's, a digital cooperative control sys- the ailerons, rudders, and stabilizers. An analog elec- tem was flown on the NASA YF-12C airplane (Fig. 2).
tronic control augmentation system (CAS) operates in This system integrated the inlet control, autothrottle, all three axes.
airdata, and navigation functions. It dramatically im- proved flightpath control and range, though the inte- For the NASA F-15 airplane, a digital electronic gration was not optimized (Ref. 2). This technology flight control system (DEFCS) augments the standard was transitioned into production when the concept was flight control system. The DEFCS replaces the analog implemented on the SR- 71 fleet. CAS. It is a dual-channel, fail-safe system programmed in Pascal. The Military Standard 1553B (Ref. 8) data In the early 1980's, NASA transitioned integrated bus input-output capability and the unused capacity in controls research to the F-15 airplane. First, the digi- the DEFCS computers may be used for other functions.
tal electronic engine control (DEEC) was flight-tested (Ref. 3). Later, the engine control was integrated Engine Control System with the flight control system in the Highly Integrated The standard FlO0-PW-100 engines (Pratt & Whit- Digital Electronic Control (HIDEC) Program (Ref. 4).
ney, West Palm Beach, FL) have a hydromechani- This program demonstrated significant improvements cal control and a supervisory electronic engine con- in thrust, fuel consumption, and engine life. Further trol. The FlO0 engine model derivative (EMD) engines extension of the integration to include real-time, on- have DEEC systems. The DEEC is a full-authority, board optimization of engine, inlet, and flight control single-channel control with a simple hydromechanical variables (performance seeking control (PSC)) was also secondary control. A universal asynchronous receiver- accomplished (Ref. 5). Integration also made it pos- transmitter data bus provides input-output capabil- sible to develop a self-repairing flight control system ity. These engine controls may communicate with the (SRFCS) on the F-15 (Ref. 6), which has been success- DEFCS for integrated control research.
fully tested. A propulsion-only flight control system, which uses the engines for emergency flight control, Inlet Control System was also developed and tested (Ref. 7).
A digital control system positions the three inlet This paper presents an overview of the integration variables. These inlet controllers were modified to ac- research programs conducted on the F-15 HIDEC air- cept bias signals for the inlet cowl and ramps from plane. Figure 3 depicts the chronological order of each the DEFCS, making integrated inlet control research integrated control flight research program. Descri p- possible.
t ions and benefits of the F-15 research are presented.
Avionics Airplane Description The F-15 avionics system has evolved over the years as a result of integrated controls research programs.
The NASA F-15 HIDEC Flight Research Aircraft Figure 5 shows a recent system architecture, and is a national facility for conducting integrated flight- Fig. 6 depicts the aircraft configuration. Three data propulsion control research. It is a single-scat, busses are used to communicate between the various (FTIT) limit is approached. Thrust is increased at the components, and a data bus interface and control unit expense of reduced engine stall margin.
ties these busses together.
Substantial stall margin is built into engine control On the MIL-STD-1553B bus are the DEFCS, the schedules to accommodate the distortion produced at NASA data system, an uplink telemetry system that high angles of attack or high sideslip angles. In the receives information transmitted from a ground-based ADECS at EPR mode, some of the stall margin re- computer, and a general-purpose digital computer. served for extreme inlet distortion is used to increase This general-purpose computer uses 32-bit words and thrust in regions of low distortion. As flight conditions has a throughput of approximately 2.5 million instruc- produce high inlet distortion, the amount of EPR up- tions/sec and a memory of 2 Mbytes. This computer trim is reduced to restore stall margin. Figure 8 shows may be programmed in high-order languages such as a typical stability audit with and without EPR uptrim.
Ada, FORTRAN, and Pascal. It has been used for the Additional information on stability audits and defini- PSC subsonic flight research.
tion of the amount of stall margin available are found in Ref. 4.
The standard F-15 (H009) data bus communicates with the inertial measuring unit, the attitude and The EPR uptrim control law is implemented in the heading reference set, a horizontal situation indica- digital flight control computer (DFCC). When the pre- tor, an airdata computer, a central computer, and a dicted angle of attack (a) and sideslip angle (/3) are cockpit navigation control indicator. The DEECs in- moderate, the controller issues an EPR command to stalled on both engines constitute the remaining part the engine causing the engine to operate close to the of the avionics system. Their universal asynchronous stall line. The DFCC uses airframe pitch, roll, and receiver-transmitter data bus communicates with the yaw rates and normal, lateral, and axial accelerations aircraft through the data bus interface and control unit. to predict angles of attack and sideslip. As these pre- The NASA F-15 aircraft, configured with this avionics dicted angles become large, the controller decreases the system, provides a uniquely capable and flexible system uptrim signal to ensure stall-free engine operation. De- for controls integration research. tails of the EPR uptrim logic are given in Ref. 4.
The ADECS also provides a constant-thrust or ex- Integrated Flight Propulsion Control tended engine life (EEL) mode that improves the en- Modes gine thrust-specific fuel consumption (TSFC) and in- creases engine life by reducing turbine temperature.
The pilot carried out the only integration of aircraft- This EEL mode increases EPR while reducing engine engine controls in the original F-15 by trying to op- airflow to maintain constant thrust for a given power timize throttle and stick commands for a given mis- setting. Figure 9 illustrates this mode.
sion. Trim control and feedback compensation were carried out by separate flight, inlet, and engine con- Flight Results trollers without benefit of shared information.
The designers were aware of the airflow demands of Figure 10 indicates the improvements in thrust for the engine and designed the F-15 variable inlet geome- intermediate power at various altitudes. These im- try schedules accordingly. The designers of the DEEC provements range from approximately 8 percent at knew what pressure distortion levels were encountered 10,000 ft to 10.5 percent at 30,000 ft. If the engine is behind the F-15 inlet; engine control laws were pro- uptrimmed using the excess stall margin while thrust duced with sufficient stability margin to ensure stall- is held constant, TSFC can be reduced as shown in free engine operation at the worst levels of distortion.
Fig. 11. At 30,000 ft, Mach 0.6 and maximum power, But because the subsystems were not designed to com- a 16-percent reduction in TSFC was obtained. This municate in flight, performance compromises were un- compares well with the predicted value of 17 percent.
avoidable.
More details of predicted versus actual results can be found in Ref. 9.
Adaptive Digital Engine Control System The EEL mode was shown to reduce engine turbine Figure 7 depicts the integration of the engine control temperature up to 80 °F. Figure 12 shows this reduc- system to the flight control system. In using ADECS, tion in temperature. This has been predicted by the additional thrust was obtained at intermediate and engine manufacturer to be equivalent to reducing high- above intermediate power settings by decreasing the pressure turbine wear rate by 50 percent at high-power nozzle throat area to increase the engine pressure ratio settings. Over a typical mission profile, this results in (EPR). This occurs at near constant airflow. The EPR 10- to 15-percent increased turbine life.
is increased until the fan turbine inlet temperature The ADECS test results proved that substantial optimization. The outputs of the optimization are two gains in excess thrust (thrust minus drag) for increased inlet variables (the cowl position and the third ramp performance, or reduction in FTIT for extended engine position), the nozzle area, engine fan and compressor life, can be realized through integrated controls. variable vane positions, core and afterburner fuel flow, fan airflow, and fan speed as illustrated in Fig. 13.
Performance Seeking Control These optimized commands are sent to the inlet con- Personnel at NASA anticipated that additional troller and to the engine controller. A detailed descrip- benefits could be realized by replacing the ADECS tion of the model, update logic, and the optimization schedules, which are based on a normal engine, process can be found in Ref. 11.
with a model-based control algorithm that adapts to engine variations. The PSC was designed Flight Results to develop such an adaptive, integrated flight- propulsion control algorithm and to demonstrate this The PSC algorithm was flight-tested throughout the control technique in flight.
subsonic envelope for both degraded and refurbished (overhauled) engines. Supersonic flight tests will be The PSC onboard adaptive real-time optimization initiated in late 1992. Results have shown that PSC algorithm has three modes: the maximum thrust mode produces significant thrust increases at key flight con- which maximizes excess thrust (thrust minus drag) ditions (Fig. 14). Thrust increases of up to 15 percent during accelerations, climbs, and dashes; the minimum were obtained on a refurbished engine. A 9-percent fuel mode, which minimizes fuel consumption during improvement was obtained in the degraded engine.
aircraft cruise; and the minimum FTIT mode, which extends engine life by reducing FTIT.
The PSC extended life mode shows a turbine tem- perature decrease of more than 160 °F at 0.9 Mach, The standard engine sensors provide input informa- 15,000 ft, military power, while holding constant thrust tion to a Kalman filter, which estimates engine com- (Fig. 15). Data with and without the engine model ponent deviations to account for other than nominal update logic (Kalman filter) show that using the com- engine performance. These component deviations rep- ponent deviation parameters improves the optimiza- resent changes in fan low-pressure turbine efficiency, tion process over optimization with a standard engine fan airflow, compressor high-pressure turbine efficiency, model. This can be seen from the additional decrease core airflow, and core turbine area. The deviations are in engine turbine temperature of more than 60 °F.
estimated within the accuracy of the Kalman filter and its inputs (Ref. 10).
The engine manufacturer estimates that at high- power settings, the engine high-pressure turbine wear The component deviation estimates are used to rate is reduced by 50 percent for a 70 to 80 °F tem- match the onboard engine model to the operating char- perature reduction. Therefore, significant engine life acteristics of the actual engine. The engine model, up- extension can be obtained using PSC. In addition to dated with the current engine component deviations, the reduced operating turbine temperature, fuel flow is combined with an engine exhaust nozzle model that was reduced 2 percent while holding constant thrust.
calculates the internal nozzle performance and external boattail drag as a function of engine and flight condi- Self-Repairing Flight Control System Program tion. An inlet-trim drag model represents the perfor- The F-15 HIDEC program, sponsored by the mance of the inlet first ramp on inlet pressure recovery, USAF, has developed, implemented, and flight-tested drag, and pitching moment, and the associated change a SRFCS. This program includes control reconfigu- in the horizontal tail position and its associated trim ration, a heads-up display (HUD) positive pilot alert drag. The inlet third ramp effects on inlet drag and system, and knowledge-based maintenance diagnostics.
recovery are also modeled. This model is assumed to be time invariant.
The SRFCS program approach exploits the inherent control redundancies of advanced aircraft by fully us- These models are simpler than the off-line simulation ing its multiple control effectors and their secondary used in ADECS. The PSC approach has the advan- aerodynamic characteristics. This is accomplished by tage, however, of tuning these models in flight. During reconfiguration, after control effector failures, to allow the current PSC research, only the engine model is re- control substitution by the remaining effectors. Instead quired to change with time to match the actual system of using massive redundant hardware on each effector operating condition.
to achieve fault tolerance and reliability, the redundant The PSC uses a linear-programming algorithm to elements become the aerodynamic forces and moments optimize the performance objectives. The PSC ap- produced by the other control effectors. The necessary proach performs a series of constrained local linear- forces and moments are generated by the alternate con- programming optimizations to converge to a global trol surfaces to provide the required aircraft motion.
Implementation In today's fighter and transport-commercial aircraft, the control systems have the power and surface dis- placement to maneuver the aircraft in a very large flight The SRFCS tested was capable of emulating an im- envelope, with surplus control capacity available from pairment and reconfiguration after detection of the im- each control surface. If failure or loss of a control sur- pairment. The SRFCS impairment failure modes could face occurs, the SRFCS uses this surplus capacity by re- be selected by the pilot and flown to assess the perfor- configuring control commands to the remaining control mance of the F-15 aircraft with and without the im- surfaces, and thus preserve the maneuvering response. pairment. Figure 17 is a block diagram of the F-15 SRFCS implementation, which includes both standard Reconfiguration is one of the few technologies that mechanical and electronic CAS. The F-15 HIDEC holds promise to meet the availability and survivabil- CAS serves to provide stability augmentation and com- ity requirements for aircraft in a hostile environment, mand response enhancement through control laws im- while minimizing the complexity and costs of the sys- plemented in a dual-channel DFCC.
tem. Knowledge-based diagnostics can provide timely and accurate fault isolation for maintenance and reduce The baseline mode was unchanged until an impair- the unnecessary removal of nonfailed equipment. ment was introduced. Two SRFCS commands, shown in Fig. 17, were added to the F-15 HID EC CAS servo The technologies demonstrated in this joint NASA controller commands. The first command forced the and USAF flight program include control reconfigura- control system to represent failure conditions. (This tion, fault detection and isolation, positive pilot alert, software was for flight test only.) The second command and maintenance diagnostics. Figure 16 illustrates added a reconfiguration correction to each control sur- how the technologies were integrated with the F-15 face servo controller. Additional details of the imple- HIDEC aircraft. Details of each technology area and mentation in the F-15 HIDEC aircraft can be found in the SRFCS process can be found in Ref. 6. Each tech- Ref. 12.
nology area is highlighted in the following list: The flight test aircraft was configured with three im- 1. Control Mixer Reconfiguration Strategy.
pairments that were selectable by the pilot. All im- The core element of the reconfiguration strat- pairments affected the right horizontal stabilator. The egy was the control mixer. The mixer accepted impairments were activated with software commands the outputs of a preexisting set of control laws to the stabilator servo actuator to accurately represent designed for an unimpaired airplane and reallo- the desired failure (Fig. 17). The commands negated cated these outputs to the surviving effectors of the mechanical system inputs and set the stabilator for an impaired airplane.
the desired impairment. Once the failure type was se- lected and activated by the pilot, it remained active 2. Fault Detection Isolation and Estimation throughout the fault detection sequence and pilot eval- (FDIE). Fault isolation was accomplished by hy- uation of the reconfigured airplane. Both the failure pothesis testing through sequential probability ra- and the correction commands disappeared upon pilot tio tests, a scheme successfully used on the NASA deactivation of the reconfiguration test mode through F-8 digital fly-by-wire analytic redundancy man- a switch on the control stick. Three types of failure agement experiment.
modes were mechanized and fight-tested: 3. Positive Pilot Alert. An integral part of the reconfiguration philosophy was the presentation in 1. Locked at trim - representing hydraulic or mechan- the HUD of the surviving flight control system sta- ical failure.
tus information to the pilot, including a situation 2. Locked at an offset position - representing a failure assessment of the existing performance limits of caused by hydraulic or mechanical jam. Values up the damaged aircraft.
.
to 6° offset locked position could be flown .
4. Maintenance Diagnostics. In addition to the reconfiguration, the SRFCS had an expert sys- 3. Partial surface loss - representing a portion of the tem capability that could detect and isolate sys- right stabilator missing because of midair collision tem component failures occurring in routine air- or battle damage of 50-, 80-, and 100-percent sur- craft use. These onboard diagnostics were adept face loss.
at finding intermittent faults that happened only in flight and relating them to casual events such Flight Test Process as maneuver action, cooling temperature, pilot in- put sequence, or other fact relationships that may Figure 18 shows the flight envelope used for be impossible to reconstruct in postflight mainte- SRFCS development. The system was developed nance troubleshooting.
for the design envelope, but it was also tested in close to the undamaged F-15 response. Additional re- the pilot maneuver envelope. The pilot could se- lect various impairments and SRFCS test modes. sults of the SRFCS flight test program can be found in The following table displays the conditions flown: Refs. 6 and 12.
(a) impairments of the right stabilator, (b) the maneu- ver sequence, and (c) the SRFCS subroutine or test Planned Research Using Propulsion-Only mode that could be selected by the pilot. Tests were Controls Technology also conducted on the maintenance diagnostics system using maneuver sequences designed to trigger the fault The SRFCS is flight-proven technology for practi- scenarios shown in Fig. 19.
cal application of new flight control systems which will greatly increase the survivability of combat aircraft Test conditions.
and enhance survivability of combat aircraft and com- (a) Right-stabilator impairment.
mercial aircraft. Part of the NASA Dryden research Locked at trim investigation was undertaken to develop methods for Locked at +2° emergency control for multisurface failures and for the Locked at +4° extreme case when most or all of the flight control sys- Locked at +6° tem became inoperative. For multiengined aircraft the 80-percent missing span research led to techniques that use the throttles for 100-percent missing span emergency controls. This research has shown that to 50-percent missing span some degree, most multiengined aircraft can be con- (b) Test maneuvers.
trolled by a closed-loop, propulsion-only flight control Pitch and roll stick doublets system. This breakthrough technology for emergency Pushover and pullup control will be flight-demonstrated on the NASA F-15 3-g windup turn HIDEC aircraft in late 1992.
3-g bank-to-bank roll The augmented control system has been imple- (c) Configuration.
mented on the NASA Dryden F-15 simulator. The No impairment propulsion-only control technique has two important Impairment features: Impairment with fault detection Impairment with effector estimator Impairment with reconfiguration mixer 1. Flight controllers such as a stick or autopilot type Impairment with complete reconfiguration sequence pitch and bank angle control knobs can be used to control the aircraft.
2. The system uses feedback of key pitch and roll Flight Results parameters to stabilize and accurately control the flightpath.
Figure 20 shows the summary results of the FDIE.
The FDIE performance was directly related to the on- board simulation model fidelity.
Figure 23 shows a block diagram of the augmented con- trol system. In the pitch axis, flightpath angle and The flight performance of the reconfiguration mixer was judged satisfactory by the NASA evaluation pilots, pitch rate feedback provide phugoid damping to sta- with the largest effects occurring for the 6° locked- bilize the system. In the roll axis, the roll rate, bank stabilator impairments. This impairment required angle, and sideslip parameters are used as required to large stick offsets just to maintain level flight, while obtain the satisfactory bank angle control. Details of the reconfigured system permitted the pilot to control the F-15 propulsion-only flight controls can be found with normal stick position. Figure 21 is an example in Ref. 7.
of the stick position change. The indication was that Results of simulations using propulsive techniques after reconfiguration, no offset was required to control for emergency control (PROTECT), Fig. 24, have the reconfigured aircraft.
shown that precise control capability was greatly en- Figure 22 shows an example of the F-15 test aircraft hanced using the closed-loop (augmented) control sys- SRFCS software performing the reconfiguration for a tem. Simulation results indicate that even inexperi- enced pilots were able to make acceptable emergency battle-damaged right stabilator missing 80 percent of its span. The fault was detected as the pilot initi- landings on the first try. Details of the simulation re- ated a bank maneuver, and the reconfiguration engaged sults for the F-15 and commercial aircraft can be found 0.35 sec later. The bank response was maintained very in Ref. 13.
Concluding Remarks Urnes, J.M., Stewart, J., and Eslinger, R., "Flight
l
Demonstration of a Self Repairing Flight Control Sys- The use of digital control systems and their ability tem in a NASA F-15 Fighter Aircraft," AGARD Guid- to share information and act on that shared informa- ance and Control Panel 49th Symposium, Toulouse, tion in an intelligent manner allow for better control France, Oct. 1989.
of the individual systems and the overall aircraft. This Burcham, Frank W., Jr. and Fullerton, C. Gordon, has resulted in significant performance benefits as high- Controlling Crippled Aircraft-With Throttles, NASA lighted in this paper and the referenced research. The TM-104238, Sept. 1991.
potential payoff for integrated technologies has barely begun. The implications of integrated technologies on USAF, MIL-STD-1553B, Digital Time Division future aircraft design are only now starting to be un- Command/Response Multiplex Data Bus, Sept. 8, derstood. These and other integrated control systems 1986. Available from Aeronautical Space Division, and the synergistic effect of integrated technologies in AFSC, Wright-Patterson AFB, Ohio.
new designs will improve the performance, reliability, Myers, Lawrence P. and Burcham, Frank W., Jr., and survivability of future aircraft.
"Preliminary Flight Test Results of the FlO0 EMD En- References gine in an F-15 Airplane," AIAA-84-1332, June 1984.
Orme, John S. and Gilyard, Glenn B., "Subsonic Burcham, Frank W., Jr. and Batterton, Peter G., Flight Test Evaluation of a Propulsion System Param- "Flight Experience with a Digital Integrated Propul- eter Estimation Process for the FlO0 Engine," AIAA- sion Control System on an F-lllE Airplane," AIAA- 92-3745, July 1992.
76-653, July 1976.
R.H. Smith, Chisholm, J.D., and J.F. Stew- Burcham, F., Gilyard, G., and Myers, L., "Propul- art, "Optimizing Aircraft Performance with Adap- sion System - Flight Control Integration - Flight Eval- tive, Integrated Flight/Propulsion Control," ASME uation and Technology Transition," AIAA-90-2280, Gas Turbine and Aeroengine Congress and Exposition, July 1990.
June 11-14, 1990, Brussels, Belgium.
Burcham, F.W., Jr., Myers, L.P., and Walsh, K.R., Stewart, James F. and Shuck, Thomas L., "Flight- "Flight Evaluation of a Digital Electronic Engine Con- Testing of the Self-Repairing Flight Control System trol in an F-15 Airplane," J. Aircraft, vol. 22, no. 12, Using the F-15 Highly Integrated Digital Electronic Dec. 1985, pp. 1072-1078.
Control Flight Research Facility," AIAA-90-1321, Myers, L.P. and Walsh, K.R., "Performance Im- May 1990.
provements of an F-15 Airplane with an Integrated Gilyard, Glenn B., Conley, Joseph L., Le, Jeanette, Engine-Flight Control System," J. Aircraft, vol. 28, and Burcham, Frank W., Jr., A Simulation Evaluation no. 12, Dec. 1991, pp. 812-817.
of a Four-Engine Jet Transport Using Engine Thrust Lambert, H.H., Gilyard, G.B., Chisholm, J.D., and Modulation for Flightpath Control, NASA TM-4324, Kerr, L.J., Preliminary Flight Evaluation of an En- Sept. 1991. (Available also as AIAA-91-2223.} gine Performance Optimization Algorithm, NASA TM- 4328, June 1991. (Available also as AIAA-91-1998.)
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Tests Stall-free •Sea level engine tests throttle •Altitude engine tests at NASA Lewis 16% increase •27 flights at NASA Dryden in supersonic dash range Features •Digital engine control Altitude • Digital inlet control operation •Advanced engine control logic 7%increase • Engine-inlet integration in thrust Faster throttle response Payoff: Established feasibility of lower idle thrust integrated propulsion controls Mach number 920239 Figure 1. Results of the F-lllE integrated propulsion control system.
ECN 2704 Figure 2. YF-12C research aircraft.
-, 1976 1994 1978 1980 1982 1984 1986 1988 1990 1992
IF-1117
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YF-12 co-op control
I F-15DEEC I
F-15HIDEC I
,---F• 1-5 --,I
_ SRFCS ... , --F--1-S_P_S_C ___ F-15 PROTECT F-18 HARV 920238 Figure 3. Integrated control flight research programs.
ID ( ( { ) llD 920241 Figure 4. NASA F-15 HIDEC flight research aircraft.
data bus 920242 Figure 5. F-15 avionics system architecture.
B Avionics and computers
[m Inlets
EEi Engines
~ Flight control 920243 Figure 6. F-15 avionics system configuration.
Digital flight control Buzz o e7 Airflow
Percent I_/ .?1/
recovery~ Airplane data: Inlet ramp angle Mach, altitude, a., ~. stick, rudder throttle and Active stall margin control loglc surface positions, Airflow, EPA Fan stall line INS data, attitudes, rates all margin EPA rmal op line Airflow 920244 Figure 7. Engine-inlet-flight control integration.
With EPA uptrim Without EPA uptrim Percent 5.5 Remaining
T
Margin available Aug mentor 5.5
r•lm
sequencing Percent Remaining 4.0 Inlet 7.5 distortion Inlet 3.0 distortion 1.5 Reynolds number 1.5 Reynolds number 1.5 1.5 Control tolerance Control tolerance Engine variation Engine variation 2.5 2.5 920245 Figure 8. Stall margin available.
Digital flight control Extended engine life mode I Temperature
Ir Digital
Fan stall Airplane data: decreasing 1 flight ~ Mach, altitude, ex., ~. stick, : control rudder throttle and Normal computer Engine surface positions, operating Airflow, EPA pressure INS data, attitudes, rates line ratio Thrust (EPA) fl EPA, fl Airflow I
.__ _____ _
decreasing I I Airflow _J 920246 Figure 9. Extended engine life mode.
r3o,ooo ft ____________________ _
Thrust 10 /._____________ ....
Increase, ------------ ~20,000 ft - ........
percent L._ ....
___..--·-·-·-·--·-·-·-·-·-·-·-·-~.-;:---
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10,000 ft 40,000 ft 5 L.. ______ .._ ______ ......_ ______ _ .6 .7 .8 .9 Mach 920247 Figure 10. Engine pressure ratio mode.
EEL □ Off OOn 16% (predicted ~17%) Thrust
T
specific fuel 0.1 consumption,
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lbm/hr/lbf ~Olb Corrected net thrust, lb 920248 Figure 11. Percentage reduction in thrust-specific fuel consumption for afterburning power at Mach 0.6 and 30,000 ft, using advanced engine control system (EEL).
Reduction in turbine temperature, OF o--_ ...... __________ _, .5 .6 .7 .8 .9 Mach number 920249 Figure 12. Results of extended-engine-life mode, military power.
Inlet parameters .1 Cowl .1 Ramp .-- __ ..__ __ ..___...., .1 Airflow ___ O __ p_t_im_·_1z_a_ti_o_n __ .1 Fan speed Aircraft and Engine flight control Real-time on-line parameters .1 Fan vane pos'n parameters optimization for thrust, .1 Core vane pos'n fuel flow, engine life .1 A/8 fuel flow .1 Nozzle pos'n .1 EPR Inlet/horizontal tail model Identification Component Real-time parameter Identification Compact Model efficiency Nozzle (Kalman filter) engine model 1-+--1--1 update ------'-----1 model
logic I Dynamic engine model I
Factors Figure 13. Performance seeking control for onboard, adaptive real-time optimization of performance.
Refurbished Thrust
Increase, 1 o
percent 0 .__ __ 920251 Figure 14. Maximum thrust mode results for performance seeking control (0.9 Mach, 15,000 ft, military power).
PSC off PSC on without Kalman filter PSCwlth Kalman filter 2000 2050 2100 2150 2200 Temperature, °F 920252 Figure 15. Extended engine life mode results for performance seeking control (0.9 Mach, 15,000 ft, military power, constant thrust).
-------------------------1
Airplane data, Positive sensors, Reconfiguration pilot surface strategy alert Onboard positions, Ground maintenance pilot maintenance diagnostics
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Inputs diagnostics expert system Failure Input detection signal and Isolation management I••••••••••••••• I Onboard Hawk ' : Kalman filter: computer~:
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I 920253 Figure 16. Self-repairing flight control system on F-15 aircraft.
_____ _..,. .___ _________ ......... 1->-1 PIiot Mechanical Stab surface Inputs control actuator Stab CAS servo f F- DFCC To aileron actuators: control I--------< Impairment : laws Correction control 1 commands I (flight test only) Sensor
----- -------
L-- ----------- Inputs -------------,
Hawk/32 processor _ ........ _
I I Fault Mixer I detect I (SIDC) I _____ Aircraft.__ _______ ,_s_u_r_fa_c_e....,__o_s_lt_lo_n__, model Gain estimator (EGE) 1 920254
----------------------------•
Figure 17. Implementation of self-repairing flight control system.
30 X 10 Altitude, ft /ft 2
"
101-----+----W--+--+---+-+-+-Hr--~------,..._-t
q = 350 lb/ft 1.0
• Fllg ht test points Vcas = 360 kn\
\ \I
Mach number 0.4 0.5 0.6 0.7 0.8 0.9 0 200 300 400 500 600 Velocity, calibrated airspeed, kn 900244 Figure 18. Flight demonstration test envelope for the F-15 SRFCS.
Fault Maneuver Failure Indication Subsystem Cause scenario conditions failed major system Connector falls 1 >3g Roll CAS Dynamic pressure disengage undergload sensor 2 1-g small None Stabllator surface Actuator connecting pitch inputs pin 1-g small Hydraulic 3 Pitch, roll CAS Stabllator actuator pitch Inputs disengage Platform stabilization 4 2-g turn Autopilot Inertial navigation disengage system fails under g load 5 5-g turn Pitch, roll CAS Pitch computer Card A loose connection disengage undergload 6 Pullup CAS disengage Right angle-of- Excessive friction In rotor attack sensor 900245 Figure 19. In-flight maintenance diagnosis scenarios.
Fault detection, right stabllator partial missing Correction detection and verification: 60 percent Detection, no verification: 40 percent False detection or verification: O percent Fault detection, right stabllltor locked Correct detection: 100 percent Estimate of remaining stabllator surface Correct value, span missing: 51 percent (±20 percent tolerance) 900246 Figure 20. Summary results of FDIE.
SRFCS , Stabllator hydraulic 1 Off 'I failure Introduced On I I Roll I stick .
.
.
force, ~.... ... ..
lb ... ., ... _,. ....... . ----- --·
0 - -,-.I.,- - - - - _......__ - ........ .._.,,._r. - • .-•- - .£..~ - ... ---· .......... -· ..
I -10 Right \_ Damaged right stabllator stabllator locked at 6° leading edge up position, deg -2 -1 0 1 2 3 4 5 6 7 8 9 10 Time, sec 900247 Figure 21. Flight data from F-15 SRFCS (Mach 0.7, altitude= 20,000 ft).
□ Undamaged
o Damaged and
reconfigured Pilot commands fl. Damaged and bank to not reconfigured bank Final EGE Bank estimate angle, 0111---i?\1--12':l::'-'l Reconfigure deg complete at 3.2 sec -25 EGE determines Initial damage amount mixer on at 2.6 sec SIDC determines right -50 stabllator damage at 2.25 sec Stabllator 80% damage at 1.0 sec -75 ..__...__ ..... __ ..... _....i.. _ ___,, __ .,__..i..--i..---J--~---- 0 2 4 10 12 14 16 18 20 6 8 Time, sec 920255 Figure 22. Bank response comparison for self-repairing flight control system.
Left throttle command Computer Ri ht throttle command Yaw angle Bank angle Bank thumbwheel Roll rate angle !1111111111 III II 1111!----:B=-a~nk=-a-n-g-=-le~~ Bank an le software command I---, Pitch rate Pitch Pitch attitude attitude thumbwheel Pitch attitude software Pitch attitude command 920256 Figure 23. F-15 augmented throttles-only control system function.
• NASA F-15 simulator landed on Edwards runway Landing trouble parameter ci sink rate (ft/sec} + • CAS off, V = 170 knots, no afterburner used bank angle (deg} + up to 30 penalty for touchdown dispersion
- - -
- -
■ Manual mode
\
40 30
o Augmented mode
Landing trouble
\
- parameter
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-
~Ji--------:-- o _, ~I t-10
-
'r-... I
-
~
.n...
Better -v- K> I..,.
Touchdown dispersion penalty
-
-
-
Approximately 1 hr
-
- - I I I I I 1 2 3 4 5 6 7 8 Landing number 920257 Figure 24. Results of PROTECT on the NASA F-15 aircraft.