GB2134866A

GB2134866A - Angle of attack based pitch generator and head up display

Info

Publication number: GB2134866A
Application number: GB08407239A
Authority: GB
Inventors: Hans R Muller
Original assignee: Sundstrand Data Control Inc
Current assignee: Sundstrand Data Control Inc
Priority date: 1980-11-28
Filing date: 1984-03-20
Publication date: 1984-08-22
Also published as: GB2134866B; GB8407239D0

Abstract

A pitch calculator system for an aircraft comprising an error correction circuit consisting of an error signal generator (74), the output of which is integrated over time by an integrator (80) to obtain a correction signal, and a hold circuit for preventing loading of said integrator with false information in the event of an unusual attitude of said aircraft, wherein the hold circuit (92-98) comprises: a source of aircraft flight signal representing the aircraft pitch, the aircraft vertical acceleration or the aircraft longitudinal acceleration; comparison means coupled to said flight signal source for providing an output signal in the event said flight signal exceeds a predetermined value; and decoupling means (78) responsive to said comparison means for disconnecting said integrator from said error generator when said output signal is provided. <IMAGE>

Description

SPECIFICATION Angle of attack based pitch generator and head up display Background of the invention This invention relates to means for generating a pitch signal which is unaffected by wind shears and turn errors and more particularly to generating a pitch stabilization signal for a head up display which has a long term component derived from an angle of attack signal and a short term signal derived from a gyroscope signal.

This invention is divided from our copending British Patent Application No.8130899 (Serial No. 2088310) which describes and claims a pitch calculating circuit for providing a pitch signal representative of the pitch of an aircraft, comprising: means for developing a first signal functionally related to the angle of attack of said aircraft; means for developing a second signal representing the air mass flight path angle of said aircraft; and means for combining said first and second signals to obtain a first pitch signal representative of the pitch of the aircraft which is not subject to acceleration and turn errors.

Prior art pitch computation circuits are disclosed in Muller U.S. Patent No. 3,851,303 and Muller U.S. Patent No. 4,095,271 both assigned to the assignee of this application.

In Muller U.S. Patent 3,851,303, a longitudinal accelerometer signal is modified by a differentiated air speed signal and is combined with a gyroscope signal to provide an indication of the pitch attitude of the aircraft. However, the differentiated air speed signal is subject to wind shear error, thereby decreasing the accuracy of the generated pitch signal.

Muller U.S. Patent No. 4,095,271, discloses a pitch generator circuit which generates a pitch signal derived from a head up display mounted accelerometer which is calibrated to the display reference axis so that it indicates the pitch angle of the reference axis during steady state, i.e.

unaccelerated flight conditions. The accelerometer output is compensated for horizontal acceleration by subtracting the air speed rate to generate a display reference computed pitch signal. This accelerometer derived pitch signal is used as a long term pitch reference and is combined with a gyroscope output for short term pitch excursions. This is accomplished by correcting the gyro pitch signal at a slow and limited rate to the long term reference.

A system disclosed by Greene U.S. Patent No.

4,012,713 utilizes the output of a longitudinal accelerometer and a differentiated air speed signal to provide a wind shear signal. The signal is fed to appropriate indicator means to alert the pilot or other aircraft of the existence of a dangerous wind shear condition.

Summary of the invention The present invention provides a pitch calculator system for an aircraft comprising an error correction circuit consisting of an error signal generator, the output of which is integrated over time by an integrator to obtain a correction signal, and a hold circuit for preventing loading of said integrator with false information in the event of an unusual attitude of said aircraft, wherein the hold circuit comprises: a source of aircraft flight signal representing the aircraft pitch, the aircraft vertical acceleration of the aircraft longitudinal acceleration; comparison means coupled to said flight signal source for providing an output signal in the event said flight signal exceeds a predetermined value; and decoupling means responsive to said comparison means for disconnecting said integrator from said error generator when said output signal is provided.

Brief description of the drawings Fig. 1 is a diagrammatic illustration of an aircraft with a head up display for pitch related visual information used in monitoring the approach of the aircraft to a landing; Fig. 2 is a functional block diagram of a circuit for generating a pitch signal; Fig. 3 is a schematic diagram showing the alignment geometry of the head up display mounted accelerometer and the angle of attack vane; Fig. 4 is a functional block diagram of a circuit for generating a pitch stabilization signal from the computer pitch signal of Fig. 2; Fig. 5 is a schematic diagram showing the alignment geometry of an angle of attack probe mounted accelerometer and a head up display mounted accelerometer; Fig. 6 is a diagrammatic illustration of an angle of attack vane with an accelerometer mounted thereon; and Fig. 7 is a functional block diagram of an alternative embodiment of a circuit for generating a display stabilization signal from the computed pitch signal of Fig. 2.

Description of the preferred embodiment The present invention is illustrated and described herein for use with a head up display system which provides pitch and flight path information to a pilot for assisting the guidance of the aircraft. However, some features of the invention are useful in providing a head up display of other pitch related information or for generating an accurate pitch signal for other purposes, such as a flight guidance system.

Referring to Fig. 1, an aircraft 20 has a head up display 22 which projects pitch related information onto a combiner screen 24 located between the pilot and the outside world. The head up display, or HUD 22 may be of the fon-n disclosed in Bateman U.S. Patent No. 3,654,806; Bateman U.S. Patent No. 3,686,626; Kirschner U.S. Patent No. 3,816,005 or Muller U.S. Patent No. 3,851,303 which are assigned to the assignee of this application.

An angle of attack derived pitch signal, denoted 0,,, utilizes a relationship between a body angle of attack, denoted a, which is the air velocity vector in relation to the fuselage reference line, or FRL, and an air mass flight path angle y. For the aircraft pitch altitude shown in Fig. 1: a8=Oa-y or, rearranging #α=αB+γ The body angle of attack a, may be computed from a local air flow angle a which is measured by an angle of attack vane 26 located on the fuselage 28 of the aircraft 20.In general, the angle av measured by the angle of attack vane 26, is related to the local air flow angle aL by the following equation: av=a+Aav or, rearranging αL=αv-#αv The local air flow angle a:, is also related to the body angle of attack a by the following equation: a=ao+Ka Rearranging and substituting for a: av-A av-ao αB= K where ao and K are aerodynamic constants determined during empirical flight testing and Aav is the angle of attack probe 26 error relative to the reference dataum to which ab and K have been determined.

The quantity stB may then be substituted into the equation for 0f, to obtain the result: av-ao - Aav #α= +γ K K The quantity y may be obtained by dividing the barometric altitude rate, denoted HBARO by the true air speed, or VTRUE and multiplying the result by 57.3. The true air speed may be obtained from an air data computer or by other sources of information. The barometric altitude rate may be obtained by differentiating the output of a barometric altimeter (not shown) which may be a part of the instrumentation of the aircraft 20.

The long term accuracy of the angle of attack based pitch signal 8, depends upon the effect that horizontal and vertical winds and shears have on the angle of attack a, and upon the air mass flight path angle y. Generally, wind and shear errors tend to cancel in the computation of the angle of attack based pitch signal #α.

Fig. 2 illustrates a pitch calculator system incorporating the invention wherein the output from the angle of attack vane 26 is utilized to provide a long term pitch signal and wherein the output from a gyroscope 29 is utilized to provide short term pitch information.

The output av of the angle of attack vane 26 is coupled to a summing circuit 40 where it is decreased by an amount equal to the constant a:O.

This signal is then multiplied by a factor of K by a multiplier circuit 42 resulting in a signal equal to: αv-αo K which in turn is equal to: #αv αB+ K This signal is added to an adding circuit 44 to the air mass flight path angle y, which is obtained by dividing the barometric altitude rate HBARO by the true air speed VTRUE and by multiplying the quotient by 57.3 in a multiplier circuit 46.

The summation of the two signals in the summing circuit 44 produces an output signal which is equal to: #αv #αv αB+ +γ=#α+ K K This signal is coupled to a summing circuit 50 through a switch contact 48 and is limited by a limiter 52 to a plus or minus 3 maximum swing.

This signal is then integrated by an integrator 54 to filter the short term dynamic disturbances of the signal.

The integrated signal from the integrator 54 is summed with the gyro output, denoted 0"Vro' in a summing circuit 56 to provide a calculated pitch output Aav K which is indicative of the pitch attitude of the aircraft 20 but which contains a constant component equal to zero K The output signal from the summing circuit 56 is coupled back to the summing circuit 50, which subtracts this signal from the output of the summing circuit 44 to produce an error signal.

The error signal is integrated over time to develop a correction signal which is added to the gyroscope signal 8,,,, to decrease long term errors which may be due to long duration wind shears.

To prevent loading of the integrator 54 with false information, such as during takeoff when the air speed is below a predetermined value, for example, 70 knots, the switch contact 48 disconnects the output of the summing circuit 44 from the input of the summing circuit 50 and couples the output Sgyro of the gyroscope 29 to the input of the summing circuit 50. Under these conditions, the output from the summing circuit 56 is the signal Sgyro only.

The limiter 52 and the time constant z of the integrator 54 are chosen such that the short term dynamic disturbances of the raw 8, signal are filtered adequately, yet typical gyro sources, such as platform erection during acceleration and turn errors, are eliminated. The resulting output signal follows the gyro pitch signal for short term changes and the av based pitch signal for long term variances.

The output from the summing circuit 44 is subtracted from the gyroscope signal 6gvro at a summing junction 49 to obtain a signal representing the vertical shear to which the aircraft is subjected. This signal is filtered in a washout circuit 51 which eliminates the short term dynamic components thereof to obtain a signal VS, representing the long term vertical windshear. The signal VS, is coupled to a comparator circuit 53 which provides an output signal to a vertical shear indicator 55 in the event the vertical windshear exceeds predetermined limits.Moreover, in the event a signal is generated by the comparator circuit 53, a switch contact 57 disconnects the limiter 52 and the integrator 54 from the summing circuit 50 and connects them to ground to prevent the calculated pitch output #αv #α*+ K from being affected by long term vertical windshears. During this time, the output of the integrator 54 is maintained at a constant level by connecting the switch contact 57 to ground.

The offset error - K is a constant which does not vary once the angle of attack probe 26 is installed. The output from the circuit of Fig. 2 may be used as part of a conventional HUD system or may be used in other types of applications which require accurate pitch information, such as flight path guidance system.

Referring now to Figs. 3 and 4, there is illustrated a system which generates a pitch stabilization signal from the output signal of the circuit of Fig. 2 and eliminates the error introduced by the angle of attack probe 26 misalignment. The circuit of Fig. 4 is particularly suited for use in a head up display which requires a pitch signal for stabilization of the display.

In Fig. 3, a longitudinal accelerometer 60 is mounted directly on the HUD platform and the HUD display unit 22 is calibrated such that when it is positioned with its reference axis level and with a zero pitch input signal, the displayed horizon line overlays the true horizon. The output of the longitudinal accelerometer 60, denoted A,1, is calibrated to read true pitch when the HUD reference axis is in nominal alignment to the fuselage reference line FRL while operating under static conditions.Once the HUD 22 and the longitudinal accelerometer 60 are installed in the aircraft, any misalignment error Aa1 of the longitudinal accelerometer 60 must be compensated for by a pitch stabilization signal, deonted oD, which is equal to the sum of a true pitch angle O and the misalignment error Aa1.

The computed pitch stabilization circuit illustrated in Fig. 4 eliminates the angle of attack probe 26 misalignment error of the output signal computer in Fig. 2 by slowly correcting the long term component of the Aav K signal to the angle measured by the HUD reference axis mounted accelerometer 60.

The output signal v oQ,,+ K is multiplied by a factor of 0.53 in a multiplier circuit 62 and is subtracted from the output A,1 of the HUD mounted accelerometer 60 in a summing circuit 64. The output of the summing circuit 64 is filtered by a filter circuit 66, which comprises a portion of a complimentary filter circuit 68. The transfer function for the filter circuit 66 is such that the long term components of the output signal from the summing circuit 64 are eliminated. The resulting high frequency components are then passed to another summing circuit 69.

The air speed of the aircraft is detected by an air speed sensor 84 and is differentiated by a rate circuit 86 to provide an air speed acceleration signal V AIR This signal is an input to a filter circuit 67 which is part of the complimentary filter circuit 68. The output of the filter 67 is added in a summing circuit 69 to the output of the filter 66 to provide a signal denoted V*, which consists of a long term component from the filter circuit 67 and a short term component from the filter circuit 66. The time constant T of the complimentary filter 68 may be made relatively long to minimize the effect of wind shears.

The signal V* from summing circuit 69 is subtracted from the output signal A,1 of the HUD mounted accelerometer 60 which, after being multiplied by a factor of 1.78 in a multiplier circuit 72, provides an output signal oAL which represents the long term pitch signal measured by the HUD mounted accelerometer 60.

An output O", which is the pitch stabilization signal output, is subtracted from the signal oAL in a summing circuit 74 to provide a second error signal which is limited by a limiter circuit 76. The output of the limiter 76 is integrated and further limited by an integrator circuit 80 to develop a second correction signal AOD which is approximately equal to the quantity Aav K This signal is subtracted in a summing circuit 82 from the output signal Aav Oa*+ K from the circuit shown in Fig. 2. The output OD* of the summing circuit 82 is then equal to the true pitch O plus the HUD alignment error Aa1.

To prevent loading the integrator 80 with false information which would cause the 0 signal to be slewed to the dynamic vertical measured by the HUD accelerometer 60 in the event of an unusual attitude or dynamic flight condition, logic circuits are provided to disconnect the integrator 80 from the limiter 76 under a specified set of circumstances.

The signal V* from the summing circuit 69 is passed through an absolute value circuit 88 and is coupled to a comparator 90, which provides an output signal in the event that V* rises above 0.3 ft/sec2. The output of the comparator 90 is then coupled to one input of a NOR gate 92.

Other inputs to the NOR gate 92 are provided by a series of comparator circuits 94, 96 and 98.

The comparator circuit 94 provides a signal in the event that the roll angle rises above a predetermined upper limit, such as 150. Similarly, the comparator circuit 96 receives as its input the output signal A,1 from the HUD mounted accelerometer 60 and provides an output when the pitch angle rises about a particular limit, such as 200. The comparator circuit 98 provides an output to the NOR gate 92 in the event that the vertical acceleration exceeds an upper limit, such as 0.2 times the acceleration of gravity. The roll angle and the vertical acceleration may be provided by an Air Data computer or by accelerometers mounted with the aircraft.

The NOR gate 92 will cause a switch contact 78 to disconnect the limiter 76 from the integrator 80 in the event that one of the comparators 90, 94, 96 or 98 indicates that an unusual attitude or a dynamic flight condition exists. This logic circuitry avoids slewing of the 0D signal to the dynamic vertical measured by the HUD mounted accelerometer 60 by preventing the error signal AOD from accumulating to an abnormally high value.

Once the logic circuitry detects that an unusual condition no longer exists, the NOR circuit 92 causes the switch contact 78 to reconnect the limiter 76 to the integrator 80, allowing resumption of normal operation.

The accuracy of the circuit of Fig. 4 depends upon the fact that the two alignment error angles Aav K and lScr, do not change at all or very little during approach to the runway. The resulting output signal SD is then equal to the true pitch 0 plus the HUD alignment error Aa1 and may be used to pitch stabilize the HUD symbology.

Referring to Figs. 5, 6 and 7, a second embodiment of a pitch stabilization circuit is shown which utilizes an angle of attack probe 26 mounted accelerometer 1 00. This embodiment of the computed pitch circuit results in generation of a 0o signal completely free of acceleration and therefore shear errors. This embodiment of the invention may be used in conjunction with the circuit shown in Fig. 2 in place of the circuit of Fig.

4.

The angle of attack probe mounted accelerometer 100, shown in Figs. 5 and 6, is mounted on a probe body 27 of the angle of attack probe 26 and is aiigned to give a zero indication under static conditions.

The angle of attack probe 26 is normally calibrated to the wind chord plane through the use of a pair of probe reference pins 27a and 27b; however, small alignment errors in relation to the FRL are possible. The accelerometer 100 output, denoted A,v, indicates this misalignment error Aav when compared with the HUD unit mounted accelerometer 60. The ideal pitch stabilization signal OD could be computed form 0, if Aav and Aa1 were known. The two misalignment angles Aav and Aa1 are not known directly, however, the difference between the two may be computed from the following equations: ALv=g(0+Av)+AH A1=g(O+Aa1)+A where AH is the horizontal acceleration of the aircraft, g is the gravitational constant and 0 is the true pitch. Subtracting A,v from AL1 yields the result: AL1-ALV=g(Aα1-#αV) or, rearranging: AL1-ALV #α1-#αV= 9 As illustrated in Fig. 5, the ideal pitch stabilization signal OD is equal to the true pitch angle plus the misalignment angle Aa, of the HUD mounted accelerometer 60.If instead, the angle of attack based pitch signal #α from Fig. 2 is used, then: αV αo Aav #D= +1;+Aa1 K K K where a0 and K are constants measured during flight testing.

The quantity v Aa1- K is not directly known; however, if the local air flow angle to body angle scale factor K were equal to 1, then OD would be equal to: #D=αV-αo+γ+#α1-#αV And, since: AL1-ALV #α1-#αV= g 0D may be computed from measurements of av and A,1A,v. However, since K is normally between 1.5 and 2, only an approximation to 0o may be calculated.

An approximate pitch stabilization signal l}D may be calculated by utilizing the relation: #αV #α1-#αV Aa1- = K K2 where K2 is a constant gain factor independent of the constant K. By substituting the approximation into the equation for OD, an equation of #D is obtained: αV αo #α1-#αV #D= + γ + K K K2 To obtain the magnitude of the error of the approximation for #D, ODI is subtracted from SD: 1 1 1 A00=00-001=Aa1(1- )+#αV( -- ~~ K2 K2 K The error signal ## thus contains two components, one proportional to Aav and the other proportional to Aa1.

The gain factor K2 may be chosen such that the error contribution of Aa,, and A are equal. If K2 is so chosen, misalignment errors of the angle of attack probe accelerometer 100 and of the HUD display mounted accelerometer 60 are reduced approximately by a factor of 5. Assuming that K is equal to 1.8, then equal error contributions of #αV and #α1 result in the value of K2 being equal to 1.285.

Referring to Fig. 7, the output ALV from the angle of attack probe mounted accelerometer 100 is subtracted from the output AL1 of the HUD mounted accelerometer 60 in a summing circuit 102. The resultant signal, denoted hA", is multiplied by a factor of 57.3 and is divided by the gravitational constant g in a multiplier circuit 104.

The output of the multiplier circuit 1 04 is then equal to Aa1-Aa and this signal is divided by the gain factor K2, which is equal to 1.285, in a circuit 106. The output of the divider circuit 106 is modified by a filtering circuit 108 which eliminates the high frequency components of the signal, and is then added to the output signal #αV #α*+ K of the circuit shown in Fig. 2 in a summing circuit 110.

The output of the summing circuit 110 is then equal to the approximate display stabilization signal oD, and is completely independent of horizontal accelerations, due to the fact that A" and ALV are subtracted and hence the horizontal acceleration term AH is cancelled. This independence relies upon the assumption that the output signal from the circuit of Fig. 2 is not affected by winds and shears due to the cancellation effect of y and 0UB

Claims

1. A pitch calculator system for an aircraft comprising an error correction circuit consisting of an error signal generator, the output of which is integrated over time by an integrator to obtain a correction signal, and a hold circuit for preventing loading of said integrator with false information in the event of an unusual attitude of said aircraft, wherein the hold circuit comprises: a source of aircraft flight signal representing the aircraft pitch, the aircraft vertical acceleration or the aircraft longitudinal acceleration; comparison means coupled to said flight signal source for providing an output signal in the event said flight signal exceeds a predetermined value; and decoupling means responsive to said comparison means for disconnecting said integrator from said error generator when said output signal is provided.

2. A pitch calculator system according to claim 1, wherein the aircraft flight signal is an aircraft pitch signal.

3. A pitch calculator system according to claim 1, wherein the aircraft flight signal is an aircraft acceleration signal.

4. A pitch calculator system according to claim 1, wherein the hold circuit comprises: a source of aircraft pitch signal; a source of aircraft vertical acceleration signal; a source of aircraft longitudinal acceleration signal; first comparison means coupled to said pitch signal source for providing an output signal in the event said pitch signal exceeds a first predetermined value; second comparison means coupled to said vertical acceleration signal source for providing an output signal in the event said vertical acceleration signal exceeds a second predetermined value;