GB2208631A

GB2208631A - Aircraft precision approach control system

Info

Publication number: GB2208631A
Application number: GB8819093A
Authority: GB
Inventors: Romeo P Martorella
Original assignee: Grumman Aerospace Corp
Current assignee: Grumman Corp
Priority date: 1987-08-13
Filing date: 1988-08-11
Publication date: 1989-04-12
Anticipated expiration: 2008-08-11
Also published as: IL87382A; DE3827482A1; GB2208631B; GB8819093D0; FR2619458A1; JPH01119500A; CA1317008C

Description

n r - 2 4 U 8 6,5 i 1 - t - AIRCRAFT PRECISION APPROACH CONTROL SYSTEM The

present invention relates generally to a Precision Approach Control (PAC) system for stabilizing an aircraft during landing thereofl such as during a relatively precise landing on an aircraft carrier, and more particularly pertains to a precision approach control system which provides the pilot with precise control over the flight path rate and flight path angle of the aircraft during landing. The precision approach control system also maintains the aircraft at a predetermined angle of attack during landing.

Precise control of the flight path of an aircraft should be maintained throughout a landing approach to an aircraft carrier, which makes this a very demanding task for a pilot. During such a landingt the pilot is presented with a relatively narrow landing window along an ideal glide slope path. The landing is further complicated by uncertain aircraft carrier motions and also by atmospheric and ship-induced turbulences.

The landing approach of high-performance# relatively unstable aircraft on an aircraft carrier is an even more demanding task, requiring precision control of the flight path by the pilot. The prior art has used Stability Augmentation Systems (SAS), Approach Power Compensators (APC)p and Direct Lift Control (DLC) subsystems to allgment the basic aircraft flying qualities and control systems# but using separate design criteria for each of these different subsystems. With the main objectives of these subsystems (short period response, phugoid damping, g control) achieved, the pilot is given improved control over the aircraft. However, in highperfQrmance, relatively unstable aircraft

1 2- requiring exceptional f light path control, this design methodology is generally insufficient since it does not assure precise flight path control.

However, none of the prior art approaches has resulted in an entirely satisfactory solution to the problem of providing a pilot with precise flight path control over an aircraft during a relatively critical landing thereof such as on an aircraft carrier.

Manual and Automatic Carrier Landing (ACL) designs resulting from an integrated approach to the flight path control problem, as well as the application of qualitative flight path control criteria, have achieved superior flight path response in a Grumman P-14 aircraft with minor modifications to its existing hardware, which has been demonstrated in studies and piloted simulations.

The present invention relates to a precision approach control system for an aircraft during landing, comprisingi a said aircraft having a plurality of operating control surfaces thereon, an autosystem for maintaining the aircraft at a predetermined angle of attack during landing thereof, a control system for maintaining the inertial flight path angle of the aircraft constant during landing thereof, And a controller, operated by the pilot, for controlling the flight path rate of the aircraft.

The present invention provides a precision approach control system or mods of operation for an aircraft which allows landing thereof in a more stable and easier manner than with existing and available control systems.

Improved control over an aircraft during landing should result in significantly enhanced flight safety and also in substantial savings in fuel since fewer bolters and wave-offs can be expected, which will result in fewer landing approaches. rower landing approaches combined with 1 c J 3- significantly enhanced aircraft control by the pilot should reduce the number of critical piloting situations, thereby significantly enhancing flight safety.

The subject Invention &loo provides a precision approach control system for an aircraft which essentially provides the pilot with a flight path angle rate controller, and which utilizes the autothrottle system to maintain the aircraft at a predetermined angle of attack ('K)p thereby defining aircraft approach speed with weight.

In one embodiment in the X-29 aircraft# the controller In the cockpit that is normally the pitch rate command stick controller during a Power Approach landing In converted Into a flight path angle rate () controller. The precision approach control mode of the present Invention provides true control of the Inertial flight path and velocity vector of the aircraft, providing the pilot with rapid and precise control over the aircraft during a landing approach. The precision approach control mode also uses the autothrottle control subsystem to maintain the aircraft at a predetermined angle of attack (cK) during landing# which In a particular disclosed embodiment for the X-29 aircraft was selected to be 0.750, which define the aircraft approach speed with weight.

The predetermined angle of attack ( cp<) would normally be different for different types of aircratti and could even be designed to be variable and selected by the pilot.

The precision approach control system or mode of operation of the present invention In designed to control the approach of an aircraft during landing to provide a more stable flight path and easier inode of landing, which In very important In critical landing situations# such an during the landing of an aircraft on an aircraft carrier or on a relatively short runway.

1 1 4_ in operation of the precision approach control systemp when the aircraft is subjected to vertical or horizontal winds or wind shear, the system controls the aircraft to maintain the inertial flight path angle constant, which essentially defines operation in the precision approach control mode.

One embodiment of the precision approach control system was designed for operation in an existing aircraft, the Grumman X-291 and the particular PAC system implemented therein used the existing controls and control subsystems onboard that aircraft. The Grumman X-29 aircraft to designed with three pilot-operated controllers, a throttle, a control command stick, and rudder pedal controls.

Disengagement of operation in the PAC mode will cause operation to revert to normal operation in the power approach mode. The precision approach control mode in the X-29 aircraft in designed to be capable of being overridden by the pilot by engaging the throttle controller with a force in excess of a given threshold forcel such as above eight pounds. Moreover# the PAC mode of operation is designed to be disengaged by closure of weight-on-wheels switches on the aircraft, which indicates landing contact. Accordingly# it should be recognized that the PAC mode of operation of the present invention can be designed to be suspended by a higher priority operating system or subsystem or by the pilot.

In the precision approach control embodiment in the X-29 aircraft, the PAC mode of operation was designed to be engaged by first selecting a normal power approach mode of operation, then by engaging the autothrottle systemo and then by engaging the PAC mode, with all engagements being by normal electrical switches in the cockpit. for engagement to be complete, several other conditions must exist within 1 M5M 1 proper predefined limitst such as angle of attack probe data# attitude reference data, nomal acceleration data, etc. The trim button is then operated to stabilize the rate-of-climb (descent) of the aircraft# which is shown on a needle gage, and additional trim control should not normally be required thereafter. This trim requirement in only required in the X-29 control arrangement embodiment# and alternative embodiments do not necessarily require this feature, The stick controller which is normally the pitch-command stick, controller in the cockpit is then operated in the PAC mods by the pilot to control the rate of descent of the aircraft.

The present invention for a PAC system is designed to reduce pilot workload by minimizing aircraft flight pathdeviations caused by atmospheric disturbances, by maintaining a stable, trimmed approach airspeed and by providing an optimum flight path response to pilot commands through the pitch-command stick controller (single control input), with response characteristics which are more easily perceived and predicted by the pilot, The improved performance is attained while providing for acceptable transient excursions in angle-of-attack and control surfaces rolative to aerodynamic limits, in engine thrust (throttle) variations, and also in short-period attitude excursions and damping.

The precision approach control syBtem implemented in the X-29 aircraft automatically modulates the thrust through the throttle to hold the angleof-attadk of the aircraft constant and hence the aireped constant. This provides the pilot with direct control over the aircraft's flight path angle and speed with the pitch-command center stick controller. Moreover, improved phugoid damping is obtained via thrust modulation, eliminating any tendency for an oscillatory flight path. Direct lift control is obtained 1 J M6- with incremental flap motion upon stick movement# with the canard cancelling flap pitching moments with a flap canard interconnect other systems, such as the F-14p use spoilers as the direct lift command controllers which would move with the center stick.

The present invention for an aircraft precision approach control system may be more readily understood by one skilled in the art with reference being had to the following detailed description of a preferred embodiment thereofp taken in conjunction with the accompanying drawings, wherein like elements are designated by identical reference numerals throughout the several views, and in whichs

Figure 1 is a schematic illustration of an exemplary embodiment of a canard-equipped aircraft, such as the Grumman X-29, which can be operated in a PAC mode of operation pursuant to the teachings of the present inventionj Figure 2 is a functional block diagram of a PAC outer loop control system pursuant to the subject iftventiont and Figure 3 is a functional block diagram of a PAC modified auto throttle system in accordance with the present invention.

Referring to the drawings in detail, figure I illustrates a canardequipped aircraftt such as the Grumman X-29 aircraft, and illustrates in a schematic manner an aircraft having a canard control surface 2a# a wing flap control surface 2b# and a atrake flap 2c, all of which are employed in the X-29 jet aircraft, Actuators 3 variably position the control surfaces 2al 2b and 2c. A flight control digital computer 4 of known design has a number of inputs thereto including pilot command inputs, and data inputs from accelerometers and gyros,, collectively referred 1 1 to by reference numeral S. The X-29 control system employs known components and subsystems to achieve stability for an inherently unstable aircraft by multi-control surfaces.

Figure 1 also illustrates on the right side thereof the centerline t of the aircraft, the horizon, the velocity Vector V of the aircraft# the flight path angle qr of the aircraft, and the angle of attack o4 of the aircraftf all as are well known in the art.

It should be realized that the PAC mode of operation is applicable to many different types of aircraft other than the canard-equipped aircraft of Figure 1. Moreover, the particular design of a PAC system for a particular type of aircraft will depend to a largo extent on the operating and control systems already existing onboard that aircraftj and the extent to which the design can be implemented from an existing design or from an original design.

The following description is specifically with reference to a PAC system implemented in a Grumman X-29 aircraft. Upon engagement of the PAC mode# a PAC mode light indicator in the cockpit is energized. If the speed stability mode which is part of normal power approach had been previously selected by the pilot, the speed stability switch will go off. If speed stability is automatically engaged because the aircraft speed is below 148 Xts, when the PAC mode is selected, speed stability will also disengages Disengagement of the PAC mode and reversion back to the normal power approach mode can be achieved by overriding the throttle motion with a pilot force in excess of eight pounds. Upon disengagement, a PAC solenoid held switch will disengage. Re-engagement of the PAC mod can be achieved only by pilot action through reselecting the PAC mode via the

1 4- 8.

PAC switch. The PAC raode will also be disengaged upon closure of a weight on wheals switch on the aircraft.

The operational procedure to engage the PAC mods in to first engage the normal/PA mode (flap handle in MCC# thumbwheel switch - TW-9) Next,, the boost switch and autothrottle in engaged# and then the PAC switch is engaged. The trim button is then operated to stop any motion of the ratof-climb needle W. Further trim control should not be needed thereafter. The desired rate of descent and flight path angle is then controlled by operation of the stick controller.

Figure 2 illustrates a functional block design of one embodimentof a precision approach control outer loop control system, which illustrates the control systems of the longitudinal control surfaces. Referring to the left side of Figure 2 j a or stick command signal from the controller is multiplied at 20 by a gravity constant 0 divided by the aircraft speed CAPV to obtain a signal M DL# which is then multiplied at 22 by a constant representing the stick gear gain to obtain a k command signal. This multiplication is used because the stick throw was originally scaled for incremental load factor (DNZ). An upper control branch multiplies the command signal by a constant at 24 to provide a lead of the4tr stick command signal.

A DNE signal (DNZBody cos 0 coo) represents a f6edback signal, which is also multiplied at 26 by the same constant as at 20 to provide a Ir actual signal# which is then summed at 26 with the V command signal. The signal could also be obtained directly from an inertial navigation system, when available onboard the aircraft. The output of 28 is integrated at 30 to obtain an integrated signal which is multiplied by a constant at 32. The actual signal 1 1 1 -g- is also multiplied by a constant at 34 to provide damping and stability. The three signals from 24, 32 and 34 are then summed at 36 to obtain a PAC command signal.

A 0 signal, representing the pitch rate of the aircraftj if also directed to a washout filter multiplier 38j which stabilizes the signal to mero at steady state, and the output thereof is then multiplied by a constant at 40 to provide additional damping. The output thereof Is summed at 42 with the PAC command signal from 36 to provide a PAC inner loop pitch rate command signal, This signal in then ' applied to the existing X-29 inner control loop of the X-29 aircraft, and the output thereof is applied to a summing circuit 50.

The command signal from 22 Is also multipliod by a constant at 44 and directed through a limit circuit 46, which provides position limits of f 20 to provide a A flap command signal. The& flap command signal is then multiplied by a constant at 48 to provide a DC PAC signal which is summed at 50 with the PAC mode aignal from 42 to provide a canard command signal.

A primary flap command signal from the pilot is also summed at 52 with the A flap command signal from 46 to provide ad total flap command signal for the Grumman X-29 aircraft, rigure 3 is a functional block diagram of the X-29 auto throttle system modified for the PAC mode of operation. An o(. ref signalt representing 8. 750, is summed at 60 with a signal representing the actual angle of attack of the aircraft. The resultant AC signal in multiplied by a constant at 62 to obtain a K &. P4 oignal which is multiplied by a further constant at 64# integrated at 66, and limited at 68 to obtain a TD thrust signals A DWI signal in also multiplied by a constant at 1 1 0.10%.

70j and then multiplied at 72 by the co ' 8 ( I + 300)1, wherein I is the bank angle of the aircraft, the output of which is summed #t 74 with the signal from 62. The output of 74 in then fed through a lag (I see.) filter 76 to produce a TE signal.

A 0( stick command signal is also multiplied by a constant at 78, the Output of which is fed through a 10 second washout filter 80 to provide a TF signal, which in summed at 62 with the TD and TE signals to provide an incremental thrust signal, The incremental thrust signal is multiplied at 84 to provide an incremental power lever signal.

A pilot thrust command signal (before PAC) is then summed at 86 with the incremental power signal, and the output thereof is limited at 88 to provide limits for the PAC power lever command signal at 90 for the X-29 aircraft.

In summary, Figure 2 illustrates a functional block diagram of the control laws commanding the longitudinal control surfaces, and Figure 3 illustrate# a functional block diagram of the c ontrol laws commanding the autothrottle,

The control law in a Ir command, Y hold, with the autothrottle holding epe. in order to maintain an adequate stall margin for the aircraft, the speed is increased in banked turns, While one embodiment of the present invention for a precision approach control system in described in detail herein along with several variations thereonj it should be apparent that the disclosure and teachings of the prevent invention will suggest many alternative designs to those skilled in the art.

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Claims

Claims:

1. A precision approach control system for an aircraft during landing,. comprising:

a. said aircraft having a plurality of operating control surfaces thereon; b. an autosystem for maintaining the aircraft at a predetermined angle of attack during landing thereof; c. a control system for maintaining the inertial flight path angle of the aircraft constant during landing thereof; and d. a controller, operated by the pilot, for controlling the flight path rate of the aircraft.

2. A control system for an aircraft by which the approach of the aircraft is precisely controlled during landing, said system comprising a controller, operable by the pilot, for controlling the flight path rate of the aircraft; a control system for maintaining the inertial flight path angle of the aircraft constant during landing; and an autosystem for maintaining the aircraft at a predetermined angle of attack during landing.

3. A control system as claimed in claim 1 or 2, in which the controller comprises the normal pitch rate command stick controller employed during a power approach landing and which is converted to a flight path angle rate controller in control.

4. A control system as claimed in any preceding claim, in which the autosystem comprises an autothrottle system which controls the power level of the aircraft.

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5. A control system for an aircraft by which the approach of the aircraft is precisely controlled during landing, said system being substantially as hereinbefore described with reference to the accompanying drawings.

1 z PublishBd 1988 8.+ TIe Paten' Office- SLate Hcusc. 6671 H,9--- Hcbcrn. London 11.701IR 4TF Furthcr ccpe-- ma- be obtame-4 f.-z:m Stc Patent Office Wes Branch, St Mary Cray, Orpington, Kent BR5 3RD. Printed kv Multiplex techniques ltd, St Mary Crky. Kent. Con- V87