Document
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United States Patent [19]
ill] Patent Number: 6,041,273
Burken et al.
[451 Date of Patent: Mar. 21, 2000
EMERGENCY CONTROl, AIR(:RAFT [541 4.764.872 8/1988 Miller ...................................... 364/433 SYSTEM USING THRUST MODULATION 5.330.131 7/1994 Burcham et al ...................... 244/75 R 5,337.982 8/1994 Sherry ..................................... 244/186
[751
Inventors: John J. Burken. Tehachapi: Frank W.
FOREIGN PATENT DOCUMENTS Burcham, Jr.. Lancaster. both of Calif.
3635899 10/1986 Gemaany [73] Assignee: The United States or America as represented by the Administrator of Primary Examiner--Will Jam A. Cuchlinski, Jr.
the National Aeronautics and Space Assistant Examiner--Yonel Beaulieu Administration. Washington. D.C.
Attorne); Agent, or Firm--John H. Kusmiss [57] A BSTRA CT [21] Appl. No.: t18/886,656 [22] Filed: JuL 1, 1997 A digital longitudinal Aircraft Propulsion Control (APC) system of a multiengine aircraft is provided by engine thrust [51] InL CI. 7 ..................................................... B64C 19/I)0 modulation in response to comparing an input flightpath [521 U.S. CI ............................... 701/3: 701/120; 701/121; angle signal _ from a pilot thumbwheel or an ILS system 244/12.4; 244/12.5; 244/51:244/75 R with a sensed flightpath angle _, to produce an error signal "re [58] Field of Search ................................ 701/3, 120. 121; that is then integrated (with reasonable limits) to generate a 244/51, 75 R, 181. 188, 12.3. 12.4, 12.5 drift correction signal to be added to the error signal "le after first subtracting a lowpass filtered velocity signal Veil for References Cited [56] phugoid damping. The output error signal is multiplied by a constant to produce an aircraft thrust control signal ATC of U.S. PATENT DOCUMENTS suitable amplitude to drive a throttle servo for all engines, 3.096.050 7/1963 Snowdon .................................. 244/77 each of which includes its own full-authority digital engine 3591,110 7/1971 Dramer et al ......................... 244/77 D control (FADEC) computer. An alternative APC system 3.618.878 11/1971 Klein et al ............................ 244/77 D omits sensed flightpath angle feedback and instead controls 3,627.238 12/1971 Menn .................................... 244/77 D the flightpath angle by feedback of the lowpass filtered 3.665,465 5/1972 Miller ................................... 343/7 ED velocity signal Veil which also inherently provides phugoid 3.908.934 911975 Schloeman ............................ 244/77 D damping. The feature of drift compensation is retained.
4,357.663 11/1982 Robbins et al ......................... 364/433 4.471.439 9/1984 Robbins et al ......................... 364/433 4.536,843 8/1985 Lambregts .............................. 364/434 9 Claims, 2 Drawing Sheets
_[-_--] .[--_ / THROTTLE MULTI [ENGINE
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DYNAMICS
L_J- I S+g-O AIRCRAFT
SENSORS
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U.S. Patent
Sheet I of 2
Mar. 2 !, 2000
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U.S. Patent Mar. 21, 2000 Sheet 2 of 2 6,041,273
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Flightpath 15F ('-Y
man-- ............... .
deg
5 ,,_
angle, 1.0 _ -
I I I I I
FIG 3a o _--" I
Velocity
change,
Kn
FIG. 3b
0 i ---_I I I I
Pitch
attitude
,1
rate
dag/sec
-.1 I I I I I
FIG 3c
Engine
_j_ Left engine thrust Right engine thrust
thrust 4000
change,
thrust
Ib
I Y I I I I I
-2000
FIG 3d
Altitude
change,
ft
i I I I
FIG 3e 0
10 20 30 40 50 60
Time, sec
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R EMERGENCY CONTROL AIRCRAFT tem (ILS). or the like. using a feedback signal from a SYSTEM [ISING THRUST MODULATION flightpath angle sensor, or its reasonably equivalent altitude if there is not available on the aircraft a flightpath angle ORIGIN OF INVENTION signal source aided by a feedback signal from a velocity 5 sensor or. in a less complex method with some loss in The invention disclosed herein was made by employees of performance, using only the sensed velocity signal.
the United States Government and may be manufactured and STATEMENT OF THE INVENTION used by or for the Government for governmental purposes without the payment of any royalties therefor.
In accordance with the present invention, longitudinal to control of a multiengine aircraft is provided in response to TECHNICAL FIELD a flightpath angle command, yc. produced by a pilot- operated thumbwheel, instrument landing system, or the The invention relates to an emergency control aircraft like. and compared to a sensed ffightpath angle. ,{. scaled by system that uses autothrottle thrust modulation for control of a constant. Kgma. to render the feedback signal proportional an aircraft in the event of a primary flight control system 15 to the command signal. The comparison produces an error failure.
signal. "re. which is in turn scaled by a constant Ke before BACKGROUND ART subtracting a sensed velocity signal. Vel. after lowpass filtering and scaling it by a constant Kvel to render the Aircraft flight control systems are designed with extensive velocity feedback signal proportional to the scaled ffightpath redundancy to ensure a low probability of failure. During 20 error signal. This subtraction of the scaled and lowpass recent years, however, several airplanes have experienced a filtered velocity feedback signal produces a phugoid- major flight control system failure leaving engine thrust as damped ¥ error signal. The undamped 7 error signal is the only mode of flight control. In some of these emergency integrated and scaled by a constant Kvi to produce a drift situations, engine thrust was modulated by the pilot to correction signal for addition to the phugoid damped error maintain control only of the flightpath angle ,/. but in other 25 signal, which after scaling by a constant Kve. produces an situations lateral control was also required. In the majority engine servo control signal that is then used to control the of the cases surveyed, crashes resulted and over 1.200 engine through a cockpit throttle servo that is part of a flight people died.
control computer. All of these functions involved in produc- Thus. the challenge was to create a sufficient degree of ing a servo control signal from the flightpath angle com- thrust modulation control by the Flight Control Computer 3o mand and the feedback signals ('{ and Vei) are programmed (FCC) to safely fly and land an airplane using an Instrument into a cockpit flight control computer, including the scaling Landing System (ILS). Meeting this challenge was the of the servo control signal in order for it to be compatible objective of a program to develop a Propulsion-Controlled with the servo requirements. Typically there is but one Aircraft (PCA) emergency backup system. The proposed throttle servo controlling all engines of a multiengine PCA system required that the aircraft have at least two 3s airplane, but in the event there is a separate servo for each engines, one on each wing. and that the normal control engine, the same control signal is applied to each servo.
surfaces not be locked in a hardover position which could For a less complex longitudinal control system, only the exceed the moments that could be created from the con- velocity feedback signal is used for comparison with the trolled thrust of the engines.
flightpath angle command, yc. In this case. the command The results of analytical studies and flight tests show that 40 signal is scaled by a constant Kf for it to be proportional to conlxol can be obtained by autocontroi of the engine throttles the lowpass filtered signal Vel scaled by the constant Kvel.
through the FCC. However. making a safe runway landing The error signal produced by this comparison is thus a is exceedingly difficult because of low phugoid damping and phugoid damped flightpath angle error signal. That error high pilot work load near the ground. The problem has been signal is then integrated and scaled by the constant Kvi for to not only improve emergency control performance but also 45 addition to the phugoid damped error signal for drift cor- reduce pilot work load using a PCA system. The end goal rection. This single velocity feedback correction will pro- was to make landing as well as climbing, cruising at altitude.
vide for flightpath angle control because, as the actual and descending a viable task by using the PCA system with aircraft flightpath angle increases or decreases from the minimal or no previous pilot training required.
conu'nanded angle, the aircraft velocity will decrease or 50 increase, respectively. Correction may be slower with only In the recent past. a propulsion controlled system has been this velocity feedback, but it does have the redeeming developed and disclosed in U.S. Pat. No. 5330.131 by Frank W. Burcham et al. (incorporated herein by this reference) feature of providing phugoid damping.
The novel features that are considered characteristic of based upon a program modification of an otherwise normal FCC that. in the event of a failure of the normal flight control this invention are set forth with particularity in the appended system of a multiengine airplane, substitutes normal pitch ss claims. The invention will best be understood from the axis control with symmetric control of the engines through following description when read in connection with the their FCC driven servos using pitch attitude and pitch rate accompanying drawings.
sensed by gyros to provide the feedback signals necessary to BRIEF DESCRII:q'ION OF THE DRAWINGS track a pitch command signal. That patented flight control 60 FIG. I is a functional block diagram of a PCA system system requires modification not only of the FCC but also of which is the subject of the present invention to be imple- each full-authority digital engine control (FADEC) com- mented by programming the system into the FCC of the puter for each engine. An objective of this invention is to aircraft using both sensed flightpath angle. "{. and sensed provide a longitudinal emergency backup system for a PCA velocity. Vel. feedback signals.
that requires only program modification of the FCC without 65 FIG. 2 is a functional block diagram of a modified PCA any changes in the separate FADEC computers of the system using only the feedback signal Vel from a velocity engines based upon tracking a flightpath angle command sensor.
(_') from a pilot thumbwheel, an Instrument Landing Sys-
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IqGS. 3A-3E present a time history of a PCA control of because of the e.g. offset, and a trim speed stability change.
flightpath angle, y. in response to a step change in the If an engine were mounted above the e.g.. as is the case with conmmndcd llightpath angle, yc. and recorded dynamic the MD-II tail engine, an increase in thrust would cause a changes in velocity, pitch attitude rate. engine thrust, and pitch down moment until the trim speed overcomes the altitude.
5 nosedown dynamics. Other effects, such as ram drag and engine inlet location, are other important considerations in DIZI'AILED DESCRIIYrlON OF THE the aircraft dynamics.
INVENTION PCA Control System Design The test vehicle for the invention to be described in detail All large civil transports have at least two engines; 10 therefore, the design philosophy was to make the PCA was the MD-I 1 wide-body transport which has a mechanical longitudinal control program for the MD-11 generally appli- flight control system with hydraulically powered actuators.
cable not only to work primarily with the two wing engines The MD-II airplane is equipped with a Flight Control but also to include the third engine as though it were an Computer (FCC) which provides a Flight Management additional pair of wing engines so that if a civil transport has System (FMS) that integrates autopilot, navigation, and 15 more engines, the aircraft could take advantage of this PCA autoland functions. The autopilot control includes a thum- bwheel for commanding flightpath angle. _/,-,,a during all control system. If all the aircraft engines do not lie on the same horizontal plane, pitching moment and velocity (Vel) phases of flight, namely climbing, cruising and descending.
changes can be compensated independently. The MD-11 test The hydraulic power for control of all aircraft flight aircraft falls into this category since it has a center engine effectors is provided by three independent systems, each 20 above the e.g. the thrust of which can be used directly for designed for fail-safe capability, i.e.. for not failing in a trim speed and dynamic control of flightpath angle. Lateral hardover position. Essential flight control functions may be control may be effected by differential control of the two maintained by any one of these three independent systems.
wing engines while longitudinal control is taking place, but Pitch control is provided by dual elevators, one on each the present invention relates only to the problem of longi- horizontal stabilizer, and pitch trim is provided by a movable 25 tudinal propulsion control.
horizontal stabilizer. Inboard and outboard ailerons supple- The longitudinal control laws developed assume not only mented by wing spoilers provide roll control. A dual rudder that the normal control surfaces are not functioning but also mounted on a single vertical stabilizer provides yaw control.
Flaps are extended rearwardiy and downwardly to increase are not in a hardover position. The PCA uses engine thrust modulation of all three engines driven by a closed-loop wing resistance when required. Lateral dynamics of flae 30 controller programmed into FCC of the PCA to increase bare aircraft is controlled by a yaw damper integrated as part of airframe phugoid damping and thus allow the pilot to land the FMS. and longitudinal stability augmentation is pro- safely by controlling flightpath angle ¥.
vided by the FMS through pitch dynamics. All of the lateral and pitch dynamic surfaces are controlled by hydraulic Symmetric or collective throttle inputs are used for lon- actuators.
gitudinal control. Symmetrical thrust changes cause an 35 initial change not only in velocity but also in pitch and pitch The three engines that power the MD-11 aircraft are Pratt rate. depending on the relative location of the thrust line and & Whitney (PW4460) high-pass ratio turbofan jet engines in the aircraft center of gravity at the moment of thrust change.
the 60.000 lb. thrust class. Two of the engines are mounted It would therefore be desirable to monitor pitch and pitch in underwing pods below the center of gravity (e.g.) of the rate and to so integrate them into the autothrottle control aircraft, one on each wing. and the third engine is located at command (ATC) so as to minimize phugoid oscillations.
the base of the vertical stabilizer above the e.g. The wing However. the present invention integrates only sensed veloc- engines are 121 inches below the nominal e.g.. and the tail ity in the longitudinal control laws that produces the ATC.
engine is 240 inches above the nominal e.g. with its thrust axis inclined 2.5 o down (nozzle pointing down_. Each Classical methods were initially used to provide a linear engine is provided with its own Full-Authority Digital 45 design of longitudinal control using simulators with reason- Engine Control (FADEC) computer driven by the FCC. able first cut results. Later in the flight-test phase, nonlinear time domain methods were employed for rapid setting of As is typical for high-bypass turbofans, thrust response is sealing constants. The nonlinear simulators were also used initially very slow. Once thrust levels are above 20%. the to adjust the initial scaling constants determined from the engine response improves dramatically. An "'approach idle" setting (when the flaps are extended beyond 27 °) maintains 5o linear design. The PCA system was designed with the flexibility to change the control gain in flight by using an the idle revolutions per minute (RPM) at a sufficiently high existing Multifunction Control and Display Unit (MCDU_ level that 8 second response from idle to full-power require- that provides the pilot with "'diM-a-gain" options for the ment can be met. A "'cruise idle" or "'minimum idle" setting PCA.
can require as much as 12 seconds to go from idle to full power, ff the PCA were engaged with "minimum idle." a 55 Pilot Vehicle Interface pilot-induced oscillation (PIt) could occur because of large The Flight Control Panel (FCP) is used for the pilot to set time lags. For this reason, the FCC is programmed to set the the flight path angle "/c for the PCA control. A flightpath FADEC system to "'approach idle" when the emergency angle error signal is produced by subtracting sensed flight- PCA is engaged by the pilot once a failure of the normal path angle and also subtracting from the error aircraft flight control occurs.
60 velocity in the FCC for phugoid oscillation damping. The Pitch effects occur because of thrust changes in the wing pilot could also engage the Instrument Landing System engines located below the e.g. and slightly tilted up. This (ILS) and set the PCA-system with a preselected mode to arrangement is typical of the majority of wing-engine air- control flightpath angle during an ILS approach by pushing craft. Assuming that the aircraft were initially trimmed in an approach/land button on the FCP. but another mode of level flight, an increase in thrust would cause a change in 65 velocity feedback only (described below_ could also be selected through the MCDU depending upon the maneuver flightpath angle caused by the vertical component of thrust.
a moment resulting from the horizontal thrust component the aircraft is engaged in at the time of normal major aircraft
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control failure. When any PCA selected mode is engaged. Kgma= 1.03 the "approach idle" engine setting is automatically set to The engine thrust control output command ATC is applied to the throttle servo 5 of all engines.
achieve faster engine thrust response to PCA commands.
An alternative implementation of the control law defined thus avoiding any possible PCA induced oscillation.
by Equations (1) using only a velocity feedback signal Vel Longitudinal Control of Flightpath Angle to control the flightpath angle y is shown in FIG. 2. The Longitudinal control uses symmetric thrust commands sensed velocity Vel is processed as before, but with the simultaneously directed to all engines through a cockpit constant Kvel equal to 0.15 instead of 0.21 and other throttle servo to control the flightpath angle "y.Although the constants are set differently. The control law in the S plane control law was developed primarily to control the flightpath is as follows: angle for a landing approach, it may also be used to control the flightpath angle for up-and-away as well as level cruise ATC=Ksc*_e(Kvb'Y3+Kve') (2) control in response to flightpath angle and velocity com- mands entered in the cockpit by the pilot, the ILS. or other where: "te=_-B.
system.
15 Ksc and B are as defined above for Equation (1); ATC= Referring to FIG. I. the functional block diagram for the Kfw-- control law in the S-plane to be described below comprises: The constant gains for this embodiment of the PCA are: three summing (5") functions !. 2. and 3; an integrating (l/S) Kf=0.2x2.5 function 4: a cockpit throttle servo 5 that serves to control all K_5 three MD-I1 engines via a multiengiae control system 6; Ks_1000 aircraft dynamics sensors 7; a lowpass filter 8: a flight path Kvel=0.15 angle command _/c source 9 which may be a pilot thumb Operation of the embodiment of FIG. 2 is the same as that wheel or an ILS system: and five scaling (proportioning) of FIG. I except for the elimination of a ",/-feedback and the constants 11-15 that function like signal amplifiers set with different values for the constants Kf in place of Ke, as represented by the functional blocks 14 and I1, respectively.
constant gain values but actually implemented in a digital 25 and the different values for the same constants Kvi, Kve, computer program as a multiplier. The feedbacksignals in Ksc, and Kvel. as represented by the functional blocks 22, the PCA system are sensed flightpath angle "y for tracking 23. 24. and 25 employed to scale the signal values in and air velocity Vel for phugoid damping. A flightpath angle proportion with other values to reach the scale needed for the error signal "/e is produced by comparing the pilot input output signal ATC.
value _ with the feedback signal of sensed flightpath angle 30 FIG. 3 shows a simulation time history of a PCA control _/scaled by a constant Kgma and a velocity feedback signal of all three MD-II aircraft engines in response to a 1° Vel.r is produced by lowpass fiRering the sensed velocity flightpath angle command using the control law of Equation signal Vel. Both the flightpath error signal _ and the filtered (1). These data indicate that it took 20 seconds to reach the velocity signal Vel/ are proportioned by separately and commanded input of 1 °. This speed is slower than with only empirically determined constants Ke and KveL The propor- 35 the wing engines, because the middle engine is producing a tioned flightpath error signal Ke*_ is compared to the nosedown moment since it is above the c.g., while the wing proportioned velocity feedback signal Kvel*Vel_ and the engines below the c.g. are producing a nose-up moment.
difference is scaled by a constant Kve to produce an error Velocity increased approximately 2.5 kn then settled back to signal that is the difference between the two proportioned almost the initial speed with a pitch attitude rate of 0.1 feedback signals. The proportioned error signal Ke*_ is 40 deg/sec. The change in thrust levels for the left and tail processed through a limited integrator 4 to produce an engines are presented for the simulation engine activity. The integral of flightpath angle error which is added to the change in the altitude trace shows constant climb for the 1° proportioned error signal to produce an aircraft throttle flightpath angle command. This control law was not flight command signal that is multiplied by a scaling constant Ksc tested but is presented because of the very simple nature of to meet the input signal amplitude requirements of the 45 the design and implementation. Based on linear simulation cockpit throttle servo 5.
results, this control law could be used to safely land the The control law in the S-plane for maintaining a pilot airplane while the pilot manually provides differential commanded flight angle -/is as follows: throttle inputs for lateral/directional control in the absence of any other provisions.
Although particular embodiments of the invention have (l) 50 ATC=KscfKri/'S _e*Ke+(_te*Ke.-B')Kve) been described and illustrated herein, it is recognized that modifications may readily occur to those skilled in the art.
where: K.sc is a constant for proportioning the control Consequently, it is intended that the claims be interpreted to output signal to the cockpit throttle servo; cover such modifications and equivalents thereof.
ye=yc-y, and y is the sensed flightpath angle scaled by a What is claimed is: constant Kgma to be proportional to the command 5s i. An emergency control servo system for using thrust input ",/c in volts per degree: modulation for longitudinal control of a multiengine aircraft B=Kvel*Vel_ and Vely is sensed velocity Vel lowpass having a throttle servo for control of all engines in unison, filtered in block 8; comprising Ke and Kve are scaling constants; 60 means for independently entering a flightpath angle com- mand to be maintained and producing a flightpath angle Kvi is a scaling constant for the integral _Ke/S; and command signal yc.
ATC is the engine thrust control output to throttle servo 5 means for continually sensing said aircraft flightpath that controls all engines.
angle and producing a flightpath angle signal y.
The constant gains for the MD-I1 are: Kv_13.45 65 means for producing a flightpath error signal _ as a Kvi--0.952 fiction of the difference between said flightpath angle Kvel=0.21 command _ and said sensed aircraft flightpath angle "y.
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having aircraft dynamic sensors of flight angle y and aircraft means for forming a limited integral as a function of said error signal and velocity Vel. a flight control computer and a single throttle servo for control of all engines in unison and having means means for adding said integral to said error signal for for both independently entering a flightpath angle conunand producing an output control signal for said throttle servo, thereby to adjust thrust of said engines to reduce 5 to be maintained and producing a flightpath angle command said error signal _/e toward zero. signal 7c said emergency system comprising a closed loop 2. An emergency control servo system as defined in claim servo system for producing a throttle command signal ATC ! wherein all of said means function to produce digital for said throttle servo in accordance with the following signals for digital operation of means receiving said digital control law in the S-plane: signals, to 3. An emergency control servo system as defined in claim I further comprising means for sensing aircraft longitudinal ATC=K..wfKvi/S*"_e*Ke+(Te*Ke-B)Kve) velocity to produce a velocity signal, means for lowpass filtering and scaling said velocity 15 where: Ksc is a constant for proportioning said control signal to produce a filtered signal and output signal of said closed loop servo system to said means for subtracting said faltered velocity signal from throttle servo; said flightpath error signal ye, thereby damping any "_=yc-y. and y is the sensed flightpath angle scaled by a phugoid oscillation of said flightpath error ye.
constant Kgma to be proportional to the command 4. An emergency control servo system as defined in claim 2O input 7c in volts per degree; 3 wherein all of said means function to produce digital signals for digital operation of means receiving said digital B=Kvel*Velt, and Vele equals VeI*50/(S+50).
signals.
Ke and Kve are scaling constants; 5. An emergency control servo system for using thrust Kvi is a scaling constant for the integral yeKe/S.
modulation for longitudinal control of a multiengine aircraft 25 9. An emergency control system for using thrust modu- having a throttle servo for control of all engines in unison, lation for longitudinal control of all engines in unison and a comprising flight control computer and a single throttle servo for control means for independently entering a flightpath angle com- of all engines in unison, means for sensing aircraft velocity mand to be maintained and producing a flightpath angle and producing a signal Vel proportionate thereto, and further command signal yc.
3o having means for both independently entering a flightpath means for sensing aircraft longitudinal velocity and pro- angle command to be maintained and producing a flightpath ducing a lowpass filtered velocity signal.
angle signal 7c, and said emergency control system com- prising a closed loop servo system producing a throttle means for subtracting said lowpass filtered velocity signal command signal ATC for said single servo, a throttle com- from said flightpath command signal 7c to produce a 35 mand ATC for all engines as a function of velocity feedback flight'path error signal ye.
in accordance with the following control law in the S-plane: means for forming a limited integral as a function of said error signal and means for adding said limited integral to said flightpath ATC=Ksc*"ye(Kvi/_S+Kve) error signal ye for producing an output control signal 4o ATC to said throttle servo, thereby to adjust thrust of said engines to reduce said flightpath error signal ye where: _-'/c-B, toward zero.
B=Kvel*Velj, 6. An emergency control system of claim 5 wherein said Vel/=50/(S+50)*Ve, control law is implemented in digital form.
Ksc is a scaling constant, and 7. An emergency control system of claim 5 wherein said control law is implemented in digital form. Kvi is a scaling constant.
8. An emergency control system for using thrust modu- lation for longitudinal control of a multiengine aircraft