WO1988004444A1

WO1988004444A1 - Flight control optimization system for multi-control surface aircraft

Info

Publication number: WO1988004444A1
Application number: PCT/US1987/003264
Authority: WO
Inventors: Romeo P. Martorella; Jimmie Chin
Original assignee: Grumman Aerospace Corporation
Priority date: 1986-12-11
Filing date: 1987-12-10
Publication date: 1988-06-16
Also published as: IL84785A0; JPH01501461A; EP0292558A1; US4797829A; EP0292558A4; AU1221788A

Abstract

A flight control system utilizing multi-controlled surfaces (2a, 2b, 2c) to provide the necessary control power and feedback logic to stabilize the aircraft (1) with minimal noise amplification. In addition, a unique performance optimization loop (88) is integrated into the stability logic to position the multi-control surfaces for optimum maneuver performance.

Description

Title of the Invention: FLIGHT CONTROL OPTIMIZATION

SYSTEM FOR MULTI -CONTROL SURFACE AIRCRAFT

FIELD OF THE INVENTION The present invention relates to flight control systems, and more particularly to the use of multi-control surfaces to provide system stability and optimized performance in a multi-control aircraft which is basically unstable.

BACKGROUND OF THE INVENTION

_≠ Modern aircraft designs can include unstable . airframe configurations such as the X-29 aircraft and can include multi-control surfaces utilized for the foregoing function. Due to the location of the center of gravity of such an aircraft, there is inherent instability which must be carefully attended to by a computer-aided flight control system. Feedback in the flight control system is provided with normal acceleration and pitch rate parameters, derived from accelerometers and gyros. This feedback data is supplied to a servo system which is intended to stabilize the aircraft.

In multi-control surface high performance aircraft such as the canard-equipped X-29, the craft is inherently unstable, and dependence upon conventional flight control system technology has raised the problems of attaining stability margin while minimizing the sensitivity of the flight control system to noise. At high operational speeds, these factors detract from the effectiveness of such an aircraft.

SUBSTITUTE SHEET BRIEF DESCRIPTION OF THE PRESENT INVENTION A flight control system for stabilizing such an aircraft is disclosed utilizing multi-control surfaces to provide the necessary control power and a feedback logic to stabilize the aircraft with minimal noise ampli ication. In addition, a unique performance optimization loop is integrated into the stability logic to position the multi-control surfaces for optimum maneuver performance. The strake, flap^" and canard control surfaces of the aircraft are driven by an error signal which is comprised of a pilot stick command signal and f edback ^••components including measured vertical acceleration and pitch rate of the aircraft and a third feedback component derived from a complementary filter. The filter has a high-pass filter section which operates upon canard position data and a low-pass filter section which operates upon pitch rate.^'- A summation of the signals passing both filtering sections is summed with the vertical acceleration and pitch rate feedback signals to for a combined feedback signal. As a result, the stability margin of an inherently unstable aircraft is increased; while noise effects are kept to a minimum.

This invention has as its main function the purpose of stabilizing a basically unstable multi-controls surface aircraft. Once stability is achieved for such an aircraft, it is important to maximize the performance which is also a main objective of this invention. This is achieved by adding a separate feedback loop between the input actuators of the multi-control surfaces and the outputs from control surface sensors. Look-up tables are incorporated in the feedback loop for generating error signals as a function of Mach, altitude and angle of attack. The result is a smaller and lighter, more efficient, aircraft that may be fabricated at reduced costs with the attendant advantage of more efficient fuel consumption.

SUBSTITUTE SHEET BRIEF DESCRIPTION OF THE FIGURES The above-mentioned objects and advantages of the present invention will be more clearly understood when considered in conjunction with the accompanying drawings, in which:

FI.G..1 is a schematic illustration of a canard-equipped aircraft incorporating a basic flight control system, shown in block diagram form;

FIG. 2 is a schematic illustration of a complementary filter, incorporated in a flight control digital computer;

FIG. 3 is a block diagram of a flight control system incorporating a stability feedback loop;

FIG. 4 is a block diagram of that portion of the flight control system which incorporates a performance optimization feedback loop.

DETAILED DESCRIPTION OF THE INVENTION FIG. 1 schematically illustrates an aircraft 1 having a control surface 2a, which may be a canard, a control surface 2b, which may be a wing flap and 2c, which may be a strake flap, such as employed in the X-29 jet aircraft. Actuators 3 variably position the control surfaces 2, 2b and 2c by conventional means. A flight control digital computer 4 of known design includes a number of inputs including a pilot command input and data inputs from aσcelerometers and gyros, collectively referred to by reference numeral 5. Thus far, the system described, employs components and subsystems to achieve stability via multi-control surfaces. However, the illustrated system also incorporates position data from control surface 2a, which serves as an additional input 6 to digital computer 4 which performs a complementary filtering function, in order to obtain a synthetic pitch acceleration signal with minimum noise. The

SUBSTITUTE SHEET complementary filter is in accordance with co-pending application Serial No. 797,089.

FIG. 2 is a basic schematic illustration of digital computer 4 which is seen to include the necessary memory and control for achieving the function of a complementary filter 7. By being provided with canard position data as well as pitch rate data, the complementary filter 7 estimates pitch acceleration; further, the filter simultaneously reduces flight control system noise and improves aircraft stability margins.

An implementation of the aircraft stability feedback loop, utilizing three control surfaces, is illustrated in FIG. 3. The illustrated loop serves to maintain the aircraft in a stable mode, particularly when it encounters-a serious perturbation, such as severe wind shifts. The flight control circuit illustrated drives a strake actuator 8, flap actuator _>9, -and canard actuator 10 of the aircraft 1, which is schematically illustrated in FIG. 1 and indicated as a block in FIG. 3.. A pilot command signal on input line 38 is a pilot stick signal ( *) which is multiplied in amplifier 40 by a gain. The resulting amplified signal is fed to difference point 28 where a number of feedback signals are subtracted from the amplified stick signal. These feedback signals include vertical acceleration n₂, which is detected along output line 70 of a conventional aircraft accelerometer (not shown) followed by amplification in amplifier 30 by a gain factor K_nz. This amplified vertical acceleration feedback s.ignal is summed at point 44 with a pitch rate signal, the latter measured by a conventional aircraft gyroscope (not shown) and amplified at amplifier 32 by a gain Kg. A third -feedback signal is provided from complementary filter 7 to summation point 46 where it is summed with the previously discussed feedback signals that were summed at point 44. Before being summed at point 46, this third feedback signal undergoes

JBSTITUTE SHEET amplification in amplifier 34 by a gain Kg. The tot feedback signal 43 is fed into a lead/lag filter 42. T filter provides lead compensation to the control syst to stabilize the aircraft. The filtered signal is f into the difference point 28. The error signal appeari at the output 50 undergoes a di ferentiation and amplification by a gain K_p 52. The output is summed wi the error signal at 56. This signal goes to zero steady state. The resulting error signal is fed along line in parallel to amplifiers 58 and 60 which respective amplify the inputs thereto by gains K_s and Kf. The signals are combined with the corresponding performan optimization loop signals at summations 92 and 10 respectively, in .FIG. 4.. The respective signals a integrated via integrators 98 and 120. The outputs fr the integrators are signals which are used to drive t strake and flap actuators (8, 9) , respectively. T error signal 57 is integrated via integrator 54, and t output 62 is used to drive the canard actuator 10. Thu for a particular error signal present at summing poi 56, individual gains are provided to the control surfa actuators in order to maintain stability for aircraft The actual gains may be empirically derived for particular aircraft in accordance with well-kno techniques .

The outputs from the strake, flap and cana actuators are respectively indicated by the angul quantities δ_s, δp and <5_C existing on individual parall actuator control lines 64, 66 and 68 to caus corresponding angular displacement of the contro surfaces thereby varying the flight contro characteristics for aircraft 1. As an aircraft chang its flight path, new vertical • acceleration n₂ an pitch rate q occur and are fed back. The pitch rate q i

SUBSTITUTE SH not only fed back to the amplifier 32, but also forms input to the complementary filter 7.

At the filter, the pitch rate signal from line is connected to an input terminal 74 of a two-po estimator 80 which performs a well-known calculation" estimate pitch acceleration from pitch rate. T quantity is then fed to the input of a low pass filter which basically attenuates the high frequency no signal.

10 Canard position data <$c Present on control l

68 is connected in parallel to the complementary filte so'-as' o'-'form a^' second input thereto. This input connected to a high-pass filter 78, via connection l 76. The high-pass filter has a gain M. which relates τ_5 a well-known flight control- moment parameter. " complementary filter 7 is so named due to the fact t the low-pass filter 82 operates on pitch rate to prov pitch acceleration information in the low frequency ra while the high-pass filtered canard deflection provi

2o the complementary high frequency information. The t constants τ for filters 78 and 82 are chosen to achi maximum stability margin and minimum noise.

The outputs from filters 78 and 82 are added summing point 84 to form a filter output signal q_E, al

25 line 86, which is then input to the amplifier resulting in an amplified feedback signal from complementary filter which is added at summing point to the other two feedback signals from summing point The resultant feedback signal is connected via line 48

30 the lead/lag filter 42, and then to point 28, wher total error signal is formed. The lead/lag fil compensates for system hardware lags.

FIG. 4 illustrates a block diagram for performance optimization loop -88 which precis

35 positions the control surfaces of an aircraft af stability has been obtained by the loop of FIG. Strake and flap signals undergoing gain amplification at ₅8 and 60 (common to the loop of FIG. 3) input parallel signals to the loop 88. The first strake signal from amplifier 58 is fed, along line 90, to a first input of summing point 92. The output from the summing point is fed, via line 96, to an integrator 98 which generates a strake command δ _sσ on line 100 which is in turn introduced to strake actuator 8. The strake command during normal operation of an aircraft will include a stabilizing signal introduced to summing point 92 from amplifier 58. A performance component is introduced to the second input 94^' of summing point 92.

In order to better appreciate how the signal at the second input 94 of summing point 92 is generated, continued reference to FIG.' 4 is made.

The stabilizing signal developed from amplifier 60 (common with FIG. 3) is fed along line 102 to the performance optimization loop'88. There it is fed to a first input of summing point 104 and the output- line 118 from the summing point 104 is connected to flap actuator 9 via integrator 120, the output of which is the flap actuator command δ_fc. This command is connected in parallel along line 122 that serves as an input to a look-up table 124 which stores data as a function of Mach, altitude, and δ_fc. Conventional altimeters and air speed sensor in aircraft 1 provide input data at each moment of time relative to Mach number and altitude. An output signal from table 124 will be generated along line 126 as a result of the three inter-related instantaneous parameters. The strake command for performance optimization along line 128 is compared with the total strake command at point 129. The error from the difference point 129 is serially fed to a limiter 130 which limits the swing of the signal output from the difference point 129 prior to further processing by the loop. The output from limiter 130 is connected to contacts 132^' and 136 of a switch 134 which functions as a single pole, double throw switch. This switch normally resides in the illustrated position; but when the flaps of an aircraft are moved to a saturated, fully displaced position,- switch 134" is changed to its other state so as to eliminate feedback from terminal 132. The strake signal for optimum performance 136 is amplified by the amplifier 138 and summed with the stability loop strake signal at summation point 92. The total signal 96 is 0 passed through an integrator 98. The output 100 is the signal which drives the strake actuator 8.

The optimization loop 88 is activated by the canard command signal 62. This signal is compared with a reference canard position 110 at the difference point τ_5 141-. The reference -canard position comes from a look-up table 108, which is a function of Mach, altitude and angle of attack. The error signal from 141 undergoes amplification at amplifier 112. The output of the amplifier is connected via line 114 to a limiter 116

2o which limits the rate of change of the signal on line 114. The output from limiter 116 along line 106 is summed with the stability loop (FIG. 3) flap command signal 102 at summation point 104. This second input to the summing point is parallel connected to terminal 140

25 of switch 134 so that the signal generated as a result of look-up table 108 may be provided to switch 134 when the switch has been changed from the state illustrated, which occurs when the flaps of an aircraft are fully displaced to a saturated position, as previously explained. The

30 output from switch 134 is connected along line 136 to amplifier 138, the output of which is connected along line 94 to the second input of summing point 92 thereby completing the performance optimization loop 88.

In operation of the loop - shown in FIG. 4, the

35 error signals to limiters 116 and 130 are driven to nulls as a result of continual updating of loop 88 by look-up

SUBSTITUTE SHEET tables 124 and' 108. The flap and strake commands illustrated provide trim signals to the strake and flap actuators 8 and 9 to trim their positions after stability of the aircraft has been achieved so as to optimize performance of the aircraft.

It should be understood that the invention is not limited to the exact details of construction shown and described herein, for obvious modifications will occur to persons skilled in the art.

SUBSTITUTE SHEET

Claims

1. In a flight control system for an aircraft (1) having three control surfaces (2a, 2b, 2σ) , interconnected feedback control loops for stabilizing and optimizing the performance of the aircraft, the stabilizing loop comprising: first means (44) for summing signals representing pitch rate and vertical acceleration of the aircraft; complementary filter means (7) connected at its input to the pitch rate and canard position signal (86) for reducing signal noise; second means (46) for summing the filtered pitch rate and canard position signal (86) and an output signal from the first summing means (44);

.filter means (42) for providing lead compensation of the signal resulting at the output of the second summing means (46) ; means (28) for obtaining an error signal (50) between the compensated signal and a pilot command input signal (38) ; and parallel means (57, 90, 102) distributing the error signal as respective stabilization error signals for the three controlled surfaces; means (98, 120, 54) for integrating the stabilization error signal; and means (106, 66, 62) for connecting the stabilization error signals to actuators of corresponding control surfaces.

2. In a flight control system for an aircraft having three control surfaces (2a, 2b, 2c) , interconnected feedback control loops for stabilizing and optimizing performance of the aircraft, the performance loop comprising: a first look-up table (108) for generating a first look-up signal (110) dependent upon Mach number, altitude and angle of attack; first means (141) for subtracting a third surface command signal (62) from the first look-up signal (110) ; first means (104) for summing the resulting subtracted signal with a first surface rate stabilization error signal (102) to form a summed signal (115) ; first means (120) for integrating the summed signal to form a first surface command signal (66) ; means connecting the first command signal to a first actuator (9) for repositioning a first control surface (2b) ; second look-up table, means (124) having an input thereof (122) connected to the first surface command signal for gen-erating a second look-up signal (126) dependent upon Mach number, altitude and the first surface control signal; second means (92) for summing the second look-up table signal and a second surface rate stabilization error signal (90) ; second means (98) for integrating the second summed signal to form a second surface command signal (100); and means connecting the second command signal to a second actuator (8) for repositioning a second control surface (2σ) .

3. The structure set forth in claim 2 together with a second subtracting means (129) having a first input (126) connected to the output of the second look-up table and further having a second input connected to the output (128) of the second integrating means (98) ; means (130) for limiting the swing of the second look-up signal having its input connected to the output of the second subtracting means (129) ; and means for connecting the limiting means output to a second input (94) of the second summing means (92) for establishing a feedback loop through the limiting means.

4. The structure set forth in claim 3 wherein the first control surface is an aircraft flap (2b) .

5. The structure set forth in claim 3 wherein the second surface is an aircraft strake (2c) .

6. The structure set forth in claim 3 wherein the first surface is an aircraft flap (2b) and the second surface is an aircraft strake (2c) .

7. The structure set .forth in claim 3 together with switching means (134) , responsive to displacement of the first control surface tα an extreme position, for switching an input (94) of the second summing means (92) from the limiting means output to the first look-up signal (110) .

8. The structure set forth in claim 6 together with means (130) connected in circuit with the output of the first subtracting means (129) for limiting the swing of the first look-up signal.

9. The structure set forth in claim 1 together with a performance loop comprising: a first look-up table (108) for generating a first look-up signal (110) dependent upon Mach number, altitude and angle of attack; first means (141) for subtracting a third surface command signal (62) from the first look-up signal (110) ; first means (104) for summing the resulting subtracted signal with a third surface rate stabilization error signal (102) to form a summed signal (118) ; first means (120) for integrating the summed signal to form a first surface command signal (66) ; means connecting the first command signal to a first actuator (9) for repositioning a first control surface (9b) ; second look-up table means (124) having an input thereof (122) connected to the first surface command signal for generating a second look-up signal (126) dependent upon Mach number, altitude and the first surface control signal; second means (92) for summing the second look-up table signal and a second surface rate stabilization error signal (90) ; second means (98) for integrating the second summed signal to form a second surface command signal (100) ; and means connecting the second command signal to a second actuator (8) for repositioning a second control surface (2c) ; wherein the performance loop predominates system operation when two of the preselected parallel distributed error signals (90, 102) are zero, thereby indicating stable operation.

10. The structure set forth in claim 9 wherein the three control surfaces are canard, flap and strake surfaces (2a, 2b, 2σ) , and further wherein the two preselected parallel distributed error signals are those corresponding to strake and flap surfaces.