Document
NASA Technical Memorandum 104255
Flight Testing and Simulation of an
F-15 Airplane Using Throttles for
Flight Control
Frank W. Burcham, Jr., Trindel Maine, and Thomas Wolf
N92-32864 (NASA-TM-IO4255) FLIGHT TESTING AND SIMULATION OF AN F-15 AIRPLANE USING THROTTLES FOR FLIGHT CONTROL Unci as (NASA) 21 p G3/O@ 0116915 August 1992 National Aeronautics and Space Administration
NASA Technical Memorandum 104255
Flight Testing and Simulation of an
F-15 Airplane Using Throttles for
Flight Control
Frank W. Burcham, Jr., Trindel Maine, and Thomas Wolf NASA Dryden Flight Research Facility, Edwards, California National Aeronautics and Space Administration Dryden Flight Research Facility Edwards, California 93523-0273
FLIGHT TESTING AND SIMULATION OF AN F-15 AIRPLANE
USING THROTTLES FOR FLIGHT CONTROL
Frank W. Burcham, Jr.* Tn_n del Maine** Thomas Wol_ NASA Dryden Flight Research Facility Edwards, California Abstract VC airspeed, kts a angle of attack, deg Flight tests and simulation studies using the throt- tles of an F-15 airplane for emergency flight control
Introduction
have been conducted at the NASA Dryden Flight Re- search Facility. The airplane and the simulation are ca- A multi-engine aircraft with a major flight-control pable of extended up-and-away flight, using only throt- system failure (such as loss of hydraulic pressure) may tles for flightpath control. Initial simulation results use throttle manipulation for emergency flightpath con- showed that runway landings using manual throttles- trol. Differential throttle control generates yaw, which only control were difficult, but possible with practice.
through dihedral effect, results in roll. Collective throt- Manual approaches flown in the airplane were much tle inputs may be used to control pitch. The DC-10, more difficult, indicating a significant discrepancy be- B-747, and L-1011 aircraft have had to use throttles tween flight and simulation. Analysis of flight data for emergency flight control. 1 and development of improved simulation models that resolve the discrepancy are discussed. An augmented To study the use of the propulsion system for throttles-only control system that controls bank angle emergency flight control, the NASA Dryden Flight and flightpath with appropriate feedback parameters Research Facility at Edwards, California, conducted has also been developed, evaluated in simulations, and flight, ground simulator, and analytical studies. The is planned for flight in the F-15.
study had three objectives. The first objective was to determine the degree of control power available for Nomenclature various classes of airplanes. Results from this objective have shown a surprising amount of control capability CG center of gravity for most multi-engine airplanes) The second objec- CAS control augmentation system tive was to investigate control modes that could be DEEC digital electronic engine control developed for future airplanes. An augmented control system that uses pilot flightpath inputs and feedback EMD engine model derivative control to provide throttle commands for emergency HUD heads-up display landings has been developed. This augmented system PCA propulsion controlled aircraft has been evaluated on a transport airplane simulation,: and an F-15 simulation. 3 A flight evaluation on an PLA power lever angle, deg F-15 is planned. The third objective was to provide PLF power for level flight, deg awareness of throttles-only control capability and sug- gested manual throttles-only control techniques for pi- lots. Reference 1 presents Dryden results of simulation and flight studies of several airplanes, including the *Chief, Propulsion Branch. Associate Fellow, AIAA.
** Aerospace engineer.
B-720, Lear 24, F-15, ]3-727, C-402, and B-747.
t Simulation engineer.
Copyright (_)1992 by the American Institute of Aeronau- More recently, additional flight tests have been flown tics and Astronautics, Inc. No copyright is asserted in the to investigate the details of throttles-only control for United States under Title 17, U.S. Code. The U.S. Govern- the F-15 airplane, and to develop data to compare ment has a royalty-free license to exercise all rights under with the F-15 simulation. Significant discrepancies the copyright claimed herein for Governmental purposes. All were found when the flight data were compared with other rights are reserved by the copyright owner.
F-15 simulation data. Additional flights and a series engines. These engines (company designation of improvements to the simulation have been made to PW-1128) include a redesigned fan and other improve- resolve the flight-to-simulation discrepancies.
ments. The F100 EMD engines are controlled by a dig- ital electronic engine control (DEEC). Interim control This paper reviews the principles of throttles-only system software was incorporated in these EMD en- control, recent results of propulsion-only flight con- gines. This software produces slower, nonproduction trol for the F-15, comparisons of flight to simulation engine response characteristics at low power settings data, and simulation upgrades. Although the FI00 en- that make it more representative of higher bypass tur- gines are equipped with afterburners, all tests discussed bofan engines.
in this paper were limited to nonafterburning power.
The inlets are mounted on the sides of the forward Plans for implementation of an augmented system for flight on the NASA F-15 are also discussed. fuselage, and are external compression horizontal ramp inlets with variable geometry. A variable capture-area Description of F-15 Airplane and capability exists in which the inlet cowl rotates about Instrumentation a point near the lower cowl lip. At subsonic speeds, the inlet cowl angle is normally positioned by a control The F-15 airplane (Fig. 1) is a high-performanee system as a function of angle of attack. The cowl may fighter airplane with a maximum Mach capability of be moved to the full-up inlets emergency position by 2.5. The F-15 (McDonnell Aircraft (McAir) Division the pilot.
of the McDonnell Douglas Corp., St. Louis, MO) has The NASA F-15 flight-control system has the stan- a high wing with 45 ° of leading-edge sweep and twin dard mechanical flight-control system and a digital con- vertical tails. It is powered by two Pratt & Whit- trol augmentation system (CAS). For throttles-only ney (West Palm Beach, FL) F100 a_erburning tur- control research, the CAS can be turned off and the bofan engines mounted close to the centerline in the mechanical system can be operated in an emergency aft fuselage. The thrust-to-weight ratio is very high, mode. This eliminates any flight-control system mo- approaching 1 at low altitudes with maximum after- tion except that caused by pilot inputs.
burning power. The NASA F-15 is the number 8 pre- production F-15A, has no weapons systems installed, The F-15 is equipped with a heads-up display (HUD) and has additional extensive instrumentation. The which provides flight information such as airspeed and zero-fuel weight is 29,450 lb. Fuel capacity is 11,600 lb.
altitude. A velocity vector symbol is available for de- termining the precise flightpath relative to the ground.
The engines installed in the NASA F-15 are the The F-15 airplane was instrumented to measure the developmental F100 engine model derivative (EMD) Figure 1. Three-view drawing of the F-15 airplane.
The simulationmay be run in a batch (non-real time) parameters required for the throttles-only flights. All mode or may be flown from a simulated cockpitshown typical engine and airplane parameters were measured.
in Fig. 2(a). The cockpit simulatesthe key instruments Data from individual sensors and from the digital con- in the NASA F-15 airplane. An actual F-15 stick and trol system data buses (each engine and the digital throttlequadrant are provided. The controlpanel on flight-control system) were recorded on an onboard the left allows the operator to selectspecialmodes as pulse code modulation system and also telemetered to required.
the ground. Data were presented in a ground control room for real-time monitoring and analysis. An HUD The visualdisplayprovides a limitedout-the-window camera was also provided and the signal was teleme- color view of the world with an optionalHUD overlay.
tered to the ground for real-time display. Data were The HUD information is similarto that available in the also recorded for post-flight analysis.
F-15 airplane, and includesthe velocity vectorsymbol.
The lakebed,main runway, and Edwards area are mod- F-15 Simulation eled with adequate realism for the approach-landing Two F-15 simulations (Fig. 2) were used in this task of this study. Upgrades to the Dryden simulation study,one at NASA Dryden and the other at the McAir .... that have evolved over the course of thisprojectwillbe Simulation Facilityin St. Louis, MO. The NASA discussedlater in the Results and Discussion section.
Dryden F-15 simulation is a fixed-base, full-envelope, Similar tests were conducted at the McAir simula- six-degree-of-freedom aircraft simulation. This model tion(Fig. 2(b)). This fixed-base simulation featuresan contains nonlinear aerodynamics, a nonlinear flight- actual F-15 cockpit and high-fidelity visual equipment control system, and originally, a first-order engine which projectsscenery onto a 40-ft dome. The aerody- response model. It is written in FORTRAN and .... _ ...... namic, controlsystem, and propulsion system models is modular in construction. The integration in- were similarto those at Dryden.
terval is 25 msec. Because it is an engineering ....
simulation, only those elements necessary to sup- port the flight research programs are implemented.
EC 90-227-1 (a) Dryden F-15 simulationcockpit.
Figure 2. F-15 simulation cockpits.
ORIGINAL PAGE BLACK AND WHITE PHOTOGRAP_ ORIGINAL PAGE BLACK AND WHITE PFIOTOGRAPI-, (b) McAir F-15 simulation cockpit.
Figure 2. Concluded.
Over a short period of time (approx 15 sac), added Principles of Throttles-Only Control thrust causes a speed increase, which increases The principles of throttles-only flight control I will be lift, causing a pitch rate increase, and a climb (if reviewed here, using examples for the F-15 airplane.
allowed to continue for a longer period of time, this effect will be oscillatory, see Phugoid, page 5.
Roll: Differential thrust generates sideslip, which, The degree of change to the flightpath angle is through dihedral effect, results in roll. Roll is con- proportional to the difference between the initial trolled to establish a bank angle, which results in a turn trim airspeed and the current airspeed, hence, the and change in aircraft heading. Figure 3 shows a typ- change in flightpath angle tends to increase as ical roll response to differential throttle. Once the dif- speed increases.
ferential throttle is applied, the differential thrust be- gins to increase, inducing sideslip and roll. As sideslip , Pitching moment due to thrust line offset. If the increases, the airplane directional stability generates a engine thrust line does not pass through the center moment equal to the moment from differential thrust, of gravity (CG), there will be a pitching moment and equilibrium is reached (in this case for the F-15) introduced by thrust change. For many transport with approximately 12 deg/sec of roll rate.
aircraft, the thrust line is below the CG, and in- Pitch: Pitch control due to throttle changes is more creasing thrust results in a nose-up pitching mo- ment, the magnitude being a linear function of complex. There are several effects that may be present, the thrust change. This is the desirable geom- depending on the aircraft characteristics. These effects are shown in concept in Fig. 4(a). etry for throttles-only control, because a thrust change immediately starts the nose in the same direction as will be needed for the long-term flight- 1. Flightpath angle change due to speed stability.
path angle change. The effect is more a function Most airplanes exhibit positive speed stability.
which kinetic and potential energy (speed and al- of change in thrust than change in speed, and oc- curs near the time of the thrust increase, as seen in titude) are traded. The degree of oscillation in speed and altitude is related to the speed stabil- Fig. 4(a). High mounted engiries result in a pitch down, which counters the effects of speed stabihty. ity. The phugoid oscillation is excited by a pitch, or velocity change, and will have a period of ap- Pitching moment due to thrust will cause a change in angle of attack, and hence, lift. For the F-15, proximately 1 rain., and may or may not damp the thrust line passes within + 1 in. of the verti- naturally. Figure 5 is an example of the phugoid response from the F-15 simulation in its initial cal CG, depending on fuel quantity, and this effect issmall.
configuration as excited by a 10°-step increase in PLA. The flightpath angle increase results in a , Flightpath angle change due to the vertical compo- steepening climb and speed peaks, and begins to nent of thrust.Ifthe thrustlineisinclinedto the decrease after about 15 sec, oscillating about the flightpath, as iscommonly the case,an increasein initial trim speed. In the oscillatory phugoid mo- thrust will cause a direct increase in vertical ve- tion, pitch rate is in phase with velocity, while locity, i.e., rate of climb, and a resulting increase flightpath angle (and rate of climb) lags by 90 °, in flightpath angle. For a given aircraft configu- and altitude lags by 180 °. Although a very small ration, this effect will increase as angle of attack amplitude phugoid is nearly a constant angle-of- increases (i.e., as speed decreases).
attack motion, for the size phugoid oscillations typically seen in throttles-only control, pitch rates Figure 4(b) is an actual time history of pitch rate for are significant, as shown. This results in a varia- the F-15 for a throttle increase to intermediate power. tion in angle of attack, in this case varying over It shows the overall result of the effects previously men- a 2- to 3°-range. Properly sized and timed throt- tioned, with a maximum pitch rate of 2 deg/sec. tle inputs can be used to damp unwanted phugoid oscillations. I 4. Phugoid. The phugoid is the longitudinal long pe- riod oscillation of an airplane. It is a motion in 100 -- Left "" x 80 -- .. .......... -- ]z 103 Throttle 60 Net -- _ /' -- 8 angle, i ,,"" _ thrust per deg 4020 -- I_'-..... ............ Right _ 2 englne'lb
o I I .I I I I
Roll rate, Sideslip, deg/sec 8 _- _....... - ..... 1 deg
16 / I I i i/_ 2
-1 0 1 2 3 4 5 Time, sec g10_7 Figure 3. Roll control resulting from differential thrust.
._- Commanded Actual Phugold Change in thrust . response Change In airspeed " Speed stability Change in Thrust offset fllghtpath Vertical angle component due to: of thrust Overall 0 2 4 6 8 10 12 14 16 Time, sec (a) Schematic pitch effects of thrust increase.
Measured pitch rate, 1 deg/sec 0 I I I I I I I I J 0 2 4 6 8 10 12 14 16 Time, sec _2o3.
(b) Flight data, NASA F-15, VC = 170 kts.
Figure 4. Pitch effects of a step increase in thrust on both engines.
16000 12000 ARItude, ft Pitch rate, deg/sec i Rate of climb, ft/mln.
-500 deg FIIghtpath angle, deg -10 VC, kts 100 I I I I I 0 40 80 120 160 200 Time, sec 9L:_Q91 Figure 5.
Phugoid oscillation from the F-15 simulation, VC = 170 kts, 10°-increase in PLA at 0 sec.
maximum negative pitch rate. Tests were repeated over Speed Control a range of speeds, and in some cases, for a suitable Once the flight-control surfaces of an airplane are range of fuel quantities (with resulting CG positions).
locked at a given position, the trim airspeed of most Another test was the full-differential throttle test, airplanes is only slightly affected by engine thrust. Re- used to determine the maximum roll rate. The airplane trimming to a different speed may be achieved by other was gently rolled to 30°-bank, then full-differential techniques, such as variable stabilizer control, CG con- throttle was applied, and the airplane rolled back trol, lowering of flaps, landing gear, etc. In general, through level and to at least 30 ° in the other direction.
the speed will need to be reduced to an acceptable This test was also conducted over a range of speeds.
landing speed; this implies developing nose-up pitch- ing moments. Methods for doing this include moving The small throttle movement test was also performed the CG aft, lowering the flaps, and extending the land- on the F-15 airplane. In this test, beginning at PLF, ing gear. For the F-15, moving the inlets to the full-up the throttles were advanced-retarded by 1 in., and the emergency position reduces the trim speed by 20 kts.
resulting pitch rates were measured. For roll rate tests, the throttles were split by 1 in. These results are more Thrust Response like the types of throttle movements that are commonly Thrust response of turbofan engines may be slow rel- used in engines-only flight control.
ative to piston or turbojet engines. The F100 EMD engine controllers in the NASA F-15 have interim soft- Typical pilot-in-the-loop maneuvers were also used to evaluate throttles-only control capability of the ware, and respond quickly at higher thrust levels, but F-15. With the flight-control surfaces fixed, the pilot at low thrust levels, respond more slowly. Idle to in- was asked to fly tests which included (1) achieve and termediate power throttle snaps take approximately maintain level flight, (2) turn to and hold a given head- 2.5 sec. Reductions to idle power exhibit a rapid re- sponse until low thrust is reached, but a very slow ing, (3) initiate and attempt to maintain a constant rate of descent, (4) use various techniques to damp a spooldown taking up to 10 sec occurs before idle thrust is reached. phugoid oscillation, and (5) make approaches to a run- way. In the simulator, the pilot was also asked to make Effects of Speed on Propulsive Control Power landings on a runway and make go-arounds from a low- For turbine-powered airplanes, engine thrust is not a approach situation.
strong function of airspeed, however, the stabilizing ef- Results and Discussion fects of vertical and horizontal stabilizers are a function of dynamic pressure, and are inversely proportional to This section discusses the development of the sim- the square of airspeed. The result of these characteris- ulation and flights of the NASA F-15 airplane for tics is that the relative propulsion system control power throttles-only control in chronological order. All data increases as airspeed decreases.
presented are with the landing gear down. Also in- cluded are the plans to implement the augmented Test Techniques throttles-only control system on the NASA F-15 airplane.
Test techniques were developed to assess the throttles-only control capability of the F-15 airplane The initial throttles-only control tests were con- and simulation. To avoid flight-control system inputs, ducted on the NASA Dryden F-15 simulation. It was the CAS was turned off, and the emergency mode was found that the F-15 had pitch capability at speeds be- selected for the mechanical system. In this mode, the low 300 kts, and roll capability at all speeds. The air- flight-control surfaces would not move as long as the plane was quite stable in the initial simulation configu- pilot did not move the stick or rudder pedals. One test ration. Flightpath control with throttles worked well; if used was the full-throttle (maximum nonafterburning) the HUD velocity vector was below the desired flight- range test. Although full throttles are rarely used dur- path, the pilot simply added thrust until it reached ing throttles-only flight, this test provides an assess- the desired position. If the flightpath was higher than ment of the maximum capability, and an easily repeat- desired, the pilot reduced the thrust until the desired able metric with which to make comparisons between flightpath was reached. With some practice, the F-15 flight and simulation.
simulation could be landed repeatedly on a runway, a Some initial throttle step tests were also conducted. At From power for level flight (PLF) conditions, both this point, initial flight tests were flown on the NASA throttles were advanced to intermediate power (max- F-15 airplane. Open-loop tests, including full-throttle imum nonafterburning) _ determine the maximum steps, were flown and control capability appeared like pitch rate capability. The same test was then repeated the simulation.
by going from PLF to idle power to determine the Full-Throttle Steps more PLA increase is available than decrease, and that more nose-upthannose-down control isavailable. The Typicalresults from the full-throttle step tests on flight data for thrust decreases are less than predicted the F-15 airplane and simulation are shown in Figs. 6 by thesimulation and will be discussed later.
and 7. The flight and simulation maximum-minimum pitch rates are shown in Fig. 6, and exhibit a response Full-differential thrust test results are shown in inversely proportional to the squareof thespeed.For Fig. 7. Again,the inverse squarerelationship to speed the F-15 and most other turbine-powered airplanes, is evident. The flight datashow somewhat less roll rate PLF in the approach to landing phase is rather low, than thesimulation results.
typically 25 to 35 percent of thrust. This means that Simulation FIIgM data m With separate gross • _ PLF to Intermediate Idle thrust and ram drag With gross thrust and ram drag assumed 2 - coincident Pitch rate, 1 - deg/sec T PLA from PLF to intermediate PLF to idle PLA from "1 -- I l I I -2 150 200 250 300 VC, kts Effect of calibrated airspeed on pitch rate for the F-15 flight and simulation.
Figure 6.
-, O
Flight Simulation %% %%%% m With separate gross 2O %% thrust and ram drag %% With gross thrust and ram drag assumed coincident Maximum roll rate, deg/sec 0 I l I I I 100 150 200 250 300 350 VC, kts Q2031Q Effect of calibrated airspeed on maximum roll rateforF-15flight and simulation, full differential thrust.
Figure 7.
Pilot-in-the-Loop Tests The engine gross thrust and ram drag terms needed to be separated since the inlet and noz- In the next flight phase, the manual throttles-only zle axes were significantly displaced. This was flight tests (with the pilot actively controlling flight- done on the Dryden simulation, and resulted in ap- path in a closed-loop fashion) were flown. These tests proximately 10-percent less roll due to differential showed that the F-15 airplane was much more difficult thrust, slightly less pitch up due to increased thrust, to fly than the simulation. Figure 8 shows a compari- and significantly more pitch down due to decreased son of approaches to a runway for the F-15 airplane thrust. The same change was also made to the and simulation. The simulation is relatively stable, McAir simulation. The effect of the changes to and only small PLA changes were required. The ac- separate the gross thrust and ram drag effects on tual F-15 airplane was never stabilized, large throttle pitch rate and roll rate is shown in Figs. 6 and 7.
Flight --- Simulation Pitch attitude, 4 deg I I I 1 I -4 Bank angle, deg -10 7O Average S $ %% I " deg 40 -- - -- power 50 " "_/_ "_1 setti ng, / 'v / 20 I 200 - kts 180 VC, 190 _1 0 20 40 60 80 100 Tlme, sec Figure 8. Comparison of flight and simulation results for a landing approach, landing gear down, VC = 170 kts.
excursions were evident, and the flightpath control was some of the small throttle step maneuvers were used to much poorer. The pilot reported strong coupling be- compare the simulated response to that of the actual tween the pitch and roll axes, large thrust lags, and aircraft. The following describes the method used to mismatches between engines. Even maintaining level make the comparisons. The simulator was set to at- tain a straight and level trim that matched the flight flight was difficult; it was not possible to attain a hands-off trim condition for more than a few seconds, Mach, altitude, and fuel weight with the CAS off and even in perfectly smooth air. The flightpath control flight control in emergency, the inlets in the emergency technique in which thrust was modulated relative to position, the gear down, and the speed brake in the the velocity vector position resulted in a large ampli- proper position. The pilot had been asked to re-trim tude oscillation. the aircraft before each maneuver, and for this study an effort was made to select time segments that started The McAir F-15 simulation was flown by the same with the aircraft more or less in trim.
pilot who had flown the NASA F-15 airplane. The To avoid step jumps caused by any mismatch be- McAir simulation flew much like the Dryden simula- tween the simulation trim and the flight trim, the initial tion, and also did not predict the great difficulty found values of the left and right PLA from flight were sub- in the flights.
tracted from the respective time histories to create in- Since the F-15 simulation model was being used to cremental PLA time histories. These incremental PLA design and evaluate the augmented mode, it was criti- time histories were then added to the simulation trim cal to resolve the major differences between the flight values to drive the simulation. The flight time histories and simulation pilot-in-the-loop results. First, the en- were plotted with the time histories generated by the gine model in both simulations was improved to in- simulation for a variety of variables characterizing the corporate the nonlinear response characteristics of the response of the aircraft. There were several problems F100 EMD engines present at low throttle settings.
with this analysis. Since this is an open-loop compar- This made the F-15 simulation more difficult to fly, but ison between the flight data and the simulation, even with practice, it was still possible to make repeatable small differences between the model and flight tend to runway landings in the Dryden and McAir simulations.
accumulate and become large with time. Thus, these There was an additional destabilizing effect in the air- comparisons are only potentially useful for short-term plane not being modeled in the simulation that made responses. Second, there is no record of the random the airplane much more difficult to control.
external forces acting on the aircraft available to drive the simulation. The pilot reported still air during these Additional effects were modeled, including engine maneuvers so it can be assumed that the effects of un- gyroscopic moments, which were found to be insignif- icant. Vertical CG effects were also investigated. Ex- modeled atmospheric disturbances are at a minimum.
treme values (thrust line 6 in. above the vertical CG) Third, reflecting the overall difficulty of flying the air- could destabilize the simulation to the degree seen in craft engines-only, the pilot had considerable difficulty flight, but the actual range of vertical C(] travel is only establishing a trim condition prior to the step inputs.
=t: lin.
There were three cases where both throttles were in- creased about 1 in. In all three cases the simulation Fuel slosh was investigated. It was thought that in- creasing power would move the fuel aft, adding more properly predicted the direction of the response, but somewhat underpredicted the pitch rate. The throt- nose-up pitching moment, and adding to the pitch re- sponse. In the roll axis, differential thrust could move tle step also excited roll rate oscillations in all three fuel in the wing tanks in a direction to reduce the cases. A typical case is shown in Fig. 9. Fan RPM is rolling moment. shown responding to the throttle increase, along with the corresponding pitch rate, roll rate, and angle of An additional flight was flown and small (approxi- attack. These small roll oscillations resulting from en- mately 1-in.) throttle steps were tested. In addition, gine mismatches were adequately modeled in the simu- tests were flown at high, medium, and low fuel levels to lation, the primary difference was that the oscillations investigate the effects of fuel quantity. The amount of in the simulation damped out more quickly than those fuel affects fuel slosh and horizontal and vertical CG, in the airplane. These differences are in accord with pi- but only small effects of fuel quantity were seen.
lot comments on the differences observed between the A batch version (non-real time) of the Dryden F-15 flight and simulation. Note that only a very small de- simulation was modified to permit throttle positions crease in angle of attack occurred, whereas the simula- measured inflight to drive the simulation. This way, tion showed a larger decrease.
lO°-step Increase 1.5 Pitch 1.0 rate, deg/sec .5 In b°th thr°ttles I i_!
" I I l m.5 1.0 _\ Flight --'_i1_ 1 JL',, .5 Roll rate, deg/sec --.5 I I -1.0 Flight _ (x, deg Simulation J I I I Right _ ....................
,..o...........-°°. ....... _ ....... .°- ,., _ J Fan RPM -'--_ Left 1 I I 5 10 15 Time, sec Figure 9. Comparison of flight and simulation data for a 10°-step increase in throttle setting, VC = 175 kts.
conditions also showed the same initial pitch up and Figure 10 shows results for a typical PLA reduc- angle-of-attack increase. These results showed a serious tion. The pitch rate comparisons of flight and simula- tion data are shown where both throttles were reduced discrepancy between the simulation and flight. Fan RPM and thrust take almost 9 sec to stabilize because from PLF to idle. While the long-term response of the of the slow responding engine control logic. Fan RPM flight data was the expected pitch down, there was a and angle of attack show a direct inverse relationship.
significant initial pitch up. There was also a signifi- Figure 11 shows a cross plot of fan RPM and angle cant increase in angle of attack. Data at other flight and inertias. The right scale of Fig. 11 is the approxi- of attack for the data of Fig. 10 and also for several mate pitching moment that is required to obtain such a other cases, including another step throttle reduction change in angle of attack. Although there is some vari- and phugoid damping tests. These data represent a ability in the data, the trend with fan RPM is clear.
range of airplane weights and therefore, CG positions Throttles PLF to Idle .5 0 "_1_ "_ Pitch ""- Simulation " ' rate, -.5 ........... ----_ ............. " deg/sec -1.0 I -1.5 I " "1 T'n'_-rl.l-II- , I 3000 - ... .....
if- Right Net 2000 Left : thrust, !
Ib 1000 -1000 I I I I J
eooo
Left Fan RPM 6000 4000 I - : ......... ; ............. I-............ i 11 - deg ..................... ...................................
7 I I. ..... I I I I 0 5 10 15 20 25 30 Time, sec Figure 10. Comparison of flight and simulation data for a step throttle decrease to idle, VC = 175 kts. (Simulation without inlet airflow effect modeled).
--- Step throttle reduction Phugold damping test - .006 Change - .004 Change in In angle pitching moment of attack, deg -.002 coefficient -0.5 I 1 I I I 4 5 6 7 8 9 10x10 3 Fan RPM Figure 11. Effect of fan RPM on change in angle of attack and pitching moment coefficients, landing gear down, VC -- 175 kts, angle-of-attack range 7.7 to 11% Effects of Inlet Airflow drag, and also increases the overall airplane lift, drag, and pitching moment (this would be expected with the Since the fan RPM is proportional to engine airflow, overhanging ramp configuration of the F-15 inlet). The possible airflow effects of the inlet on airplane pitching wind-tunnel pitching moment coefficient data is shown moment were investigated. There had been extensive in Fig. 12 for the inlet ramp-full-up emergency posi- wind-tunnel tests previously conducted on the effects tion. The fairing extrapolates, based on other data, of inlet airflow on F-15 inlet and overall airplane drag, to higher values of mass flow ratio that occur at lower llft, and pitching moment. 4 These data show that speeds. This pitching moment effect would produce an reducing the inlet airflow increases the inlet lift and -.018 • Test data Assumed fairing -.020 -.022 Pitching moment coefficient -.024 -.026 "t -.028 f i 1 i _ .20 .30 .40 .50 .60 .70 .80 Inlet mass flow ratio _ Figure 12. Pitching moment due to inlet mass flow ratio, F-15 7.5-percent wind-tunnel model test results, a = 8°, Mach= 0.6.
well. Although only one case is shown, similar results effect in accordance with the flight data, i.e., a throttle were observed for all other tests. This airflow effect reduction would result in a pitch up and an increase in has also been incorporated in the piloted simulation.
angle of attack, which would eventually be overcome The pilot commented that with the inlet airflow effects by the speed stability effects as the velocity is reduced.
modeled, the simulator flies much more like the air- The lowest Mach number in the wind-tunnel study plane. Attempts are continuing to refine this pitching was 0.6. It is not clear how to extrapolate the results to moment effect to better match the flight data.
Mach 0.3 where the flight studies are being conducted, The inlet airflow effect is small, and would often be particularly since the mass flow ratio would have been neglected in an airplane simulation. However, when the higher at the lower Mach number. Two things are only moments being used for control are the small mo- noted from the flight data. First, the significant change ments from the propulsion system, normally neglected in angle of attack as a function of engine RPM seems to effects may become significant. This is particularly be limited to an intermediate range of fan speeds. Sec- true for airplanes with highly integrated propulsion sys- ond, the cases where the engine was stepped up instead tems such as fighters where inlet-airframe interactions of down did not have a comparable initial pitch down are strong. It would likely be less true for subsonic air- or significant angle-of-attack decrease as seen in Fig. 9.
These effects are consistent with the wind-tunnel inlet planes with podded engines where the inlets tend to be simple pitot inlets normal to the flow.
airflow effects shown in Fig. 12.
Differential Throttle Tests Based on these observations, the data from Figs.
11 and 12 were used to develop a piecewise linear There were four cases with primarily differential increment to the pitching moment as a function of throttle input. In all cases, the simulator responded inlet airflow with no increment being added at the with somewhat more roll rate in response to the dif- higher airflow. With this airflow effect, it has been ferential throttle input than the aircraft did. A typ- possible to substantially improve the simulator's abil- ical case is shown in Fig. 14 where the pilot initially ity to match the flight data. The results of this air- split the throttles approximately 2 in. and held that for flow effect are shown in Fig. 13, the flight data of 3 sec, then split the throttles 2 in. in the opposite di- Fig. 10 are shown with the original and updated simu- rection. The yaw rate match is very good. The result- lation. The changes in pitch rate are properly mod- ing roll rate oscillations were comparable in frequency eled, and the trend for angle of attack is predicted Throttles PLF to idle .5
;, F,ght
...... , ,____liar" Ir ]11_.. ,-- Simulation with Pitch "__"-,,, ...... ' __/_% airflow effect _. Slmulatlon..n_..
rate, -.5 deg -1.0 -1.5 - LL li,'_J Flight /- Simulation with 10 - - ...... - deg 8 I I I _ i 0 5 10 15 20 25 30 Time, sec Figure 13. Comparison of flight and simulation results for a throttle step from PLF to idle, VC = 175 kts (simulation with and without inlet airflow effect modeled).
simulation roll rates shown in Fig. 6. The inlet airflow and damping in the flight and the simulator response, effects that are important in pitch have only a minor although the roll rates were higher in the simula- effect on the yawing and rolling moments due to differ- tion than in the flight data. These roll rates agree ential throttle.
with the previously collected data comparing flight and ,:'! " .... _-- Right PLA, deg
Left
_ ............ /--
I I I I I Pitch rate, deg/sec i Bw_ - ', Roll rate, '" F Flight deg/sec ', I I I I I l Yaw rate, -2 .F Flight deg/sec - Simulation _ .............. -4 I I I I I -6 0 5 10 15 20 25 Time, sec Figure 14. Comparison of flight and simulation response for a differential throttle input, VC = 175 kts.
system will provide the feedback signals and digital en- Augmented Throttles-Only Control gine control systems on each engine will be used to System move the throttles to the commanded position. The Manual throttles-only control is difficult for up-and- PCA control logic will reside in the digital flight-control away flight and a successful landing on a runway would computer.
be extremely unlikely for the NASA F-IS, based on pi- Concluding Remarks lot comments. However, an augmented propulsion con- trolled aircraft (PCA) concept 2 shows promise of being A flight and simulation evaluation of the throttles- able to make repeatable runway landings. Figure 15 only control capability of the F-15 airplane has been shows an augmented PCA system designed for the conducted. Principles of throttles-only control have F-15.1 Appropriate feedbacks are used to stabilize been shown. Initial flight-to-simulation comparisons the pitch and roll axes. Thumbwheel controllers re- were good for differential throttle and increasing throt- mind the pilot that the system is a slow-response, low- tle, but were poor for decreasing throttle. Detailed authority system. Initial simulation results based on comparisons of flight and simulation data have revealed the first NASA Dryden and McAir simulation showed an unmodeled pitching moment effect thought to be that the system worked well. More recently, the up- caused primarily by inlet airflow. The inlet airflow ef- dated simulation model, which flies much like the air- fect is small. However, when the only moments being plane and incorporates inlet airflow effects, has been used for control are from the propulsion system, nor- used to evaluate the PCA system. Although phugoid really neglected effects may become significant. This damping is reduced, PCA system performance is still is true for airplanes with highly integrated propulsion adequate at the lower speeds. At higher speeds, gain systems such as fighters where inlet and engine interac- changes and the addition of airspeed feedback make tions are strong, but less true for airplanes with pod- the performance of the PCA system satisfactory. The ded engines. Incorporating this effect into the simu- flight-test control laws have the capability for changing lations has greatly improved the simulation-to-flight gains, which will help with solving problems that oc- comparisons. Based on simulation results, an aug- cur during the flight evaluation. Based on the simula- mented throttles-only feedback control system shows tion, repeatable runway landings with this PCA system promise of making repeatable runway landings of the should be practical.
F-15 airplane practical.
A flight demonstration of this PCA system on the NASA F-15 is planned. The digital flight-control Left Throttle Command Computer Right Throttle Command Yaw Angle Roll Rate Bank angle Bank angle thumbwheel Bank angle software Bank Angle IUlIIIIIIIIIIIIII| command Pitch Rate Pitch attitude Pitch Attitude thumbwheel attitude Velocity > software Pitch i Pitch attitude command Figure 15. Schematic view of the augmented propulsion controlled aircraft system for the F-15.
References 3Bureham, F., Jr., C. Fullerton, G. Gilyard, T. Wolf, and J. Stewart, "A Preliminary Investigation of the Use IBurcham, Frank W. Jr. and C. Gordon Fullerton, of Throttles for Emergency Flight Control," AIAA-91- Controlling C%/ppled Aircraft - With Throttles, NASA 2222, June 1991.
TM-104238, 1991.
4Kamman, J.H. and H.W. Wallace, "Assessment of 2Gityard, Glenn B., Joseph L. Conley, Jeanette L.
Installed Inlet Forces and Inlet/Airframe Interactions," Le, and Frank W. Bureham, Jr., "A Simulation Eval- AFFDL-TR-76-62, July 1976.
uation of a Four-Engine Jet Transport Using Engine Thrust Modulation for Flightpath Control," AIAA-91- 2223, June 1991.
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L 1 B = • PubliC mporllng burden for this collection ol information II estimld_ to.aver.l_e =hour pet re(kooflle, including the lime for revt4wtngtn4trgctiordt iewchlng existing _ toume$, 0,ethering and maintaining the doll _oded, _¢1. com l_eff.ng _ roy .l_mg m.e ¢p.,eotion of i_ormauon. 8erKI ¢ommonts rngatdin 9 this burden etllmlle or_any other _¢t of mliootion of Inlormllfon. including suggestions mr maucm_ ml O _u_...en. to wunlngton Heliaquatienl Bervfoel. Otmctorme tor Intormllten Operations and He#on,=. 1:_1S Jefferlon DIIVlII .HIQ.h.Wlly, Sub 1204, Arlington, VA 2220;}-4302, rlmd to the Office o! Mlnlgemenl and Budget, PlDerwork Reduction Pr© _ I0704-01118), Washington, DC 20503.
1, AGENCY USE ONLY (Leave blank) I 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED
I August 1992 Technical Memorandum
4. TITLE AND SUBTITLE 5. FIJI_I'DINGNUMBERS Flight Testing and Simulation of an F-15 Airplane Using Throttles for Flight Control :l|. i | RTOP 533-02-36 e. AUTHOR(S) Frank W. Burcham, Jr., Trindel Maine, and Thomas Wolf 6, PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) REPORT NUMBER NASA Dryden Flight Research Facility P.O. Box 273 H-1826 Edwards, CA 93523-0273 10. SPONSORING/MONITORING 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) AGENCY REPORT NUMBER National Aeronautics and Space Administration NASA Tlvi-104255 Washington, DC 20546-0001 11. SUPPLEMENTARY NOTES Prepared asAIAA 92-4109forthe AIAAFlightTest Conference, Hilton Head, SC, August24,1992.
12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE Unclassified -- Unlimited Subject Category 08 13. ABSTRACT (Maximum 200 worda) Flight tests and simulation studies using the throttles of an F-15 airplane for emergency flight control have been conducted at the NASADryden Flight Research Facility. The airplane and the simulation are capable of extended up- and-away flight, using only throttles for flightpath conm31. Initial simulation results showed that runway landings using manuaithmlfles-only control weredifficult, but possiblewithpractice. Manual appmacbes flowninthe airplane were much more difficult, indicating a significant discrepancy between flight and simulation. Analysis of flight data and development of improved simulation models that resolve the discrepancy are discussed. An augmented throttles- only cont_l system that controls bank angle.and flightpath with appropriate feedback parameters has also been developed, evaluated in simulations, and is planned for flight in the F-15.
15. NUMBER OF PAGES 14. SUBJECT TERMS 2] 16. PRICE CODE F-15, flight test, simulation, propulsion-only control A02 17. SECURITY CLASSIFICATION 111. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ASSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified Unlimited Unclassified Unclassified NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-S9) by _s= sty. z_-_ 11)11-I02