## Patents

 Publication number USRE31159 E Publication type Grant Application number US 06/125,879 Publication date Feb 22, 1983 Filing date Feb 29, 1980 Priority date Jul 6, 1976 Publication number 06125879, 125879, US RE31159 E, US RE31159E, US-E-RE31159, USRE31159 E, USRE31159E Inventors Jean-Luc Sicre Original Assignee Societe Francaise d'Equipement pour la Navigation Aerienne Export Citation Patent Citations (4), Referenced by (2), External Links:
Flying method and system using total energy for an aircraft
US RE31159 E
Abstract
Images(6)
Claims(22)
What is claimed is:
2. A flying method according to claim 1, wherein said step of displaying includes displaying said error signal δ by means of the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar of an .[.artificial horizon.]. .Iadd.attitude director indicator.Iaddend..
3. A flying method according to claim 1, and further including calculating the .[.total gradient.]. .Iadd.potential flight path .Iaddend.γt from an accelerometric box.
4. A flying method according to claim 1, and further including calculating the .[.total gradient.]. .Iadd.potential flight path .Iaddend.γt from the speed Vs of the aircraft with respect to the ground, and providing the value of this speed Vs by a central inertial unit.
5. A flying method according to claim 4, and further including calculating γt from the formula: ##EQU4## (g being the acceleration due to gravity), or the formula: ##EQU5##
6. A flying method according to claim 5 and further including governing the derivative dVs/dt by a magnitude proportional to the difference (V-V2) in which V is the aerodynamic speed of the aircraft and V2 is the reference speed.
8. A flying method according to claim 7, and utilizing a reference .[.position.]. .Iadd.attitude .Iaddend.which corresponds to θ1 =18° and a constant reference signal which corresponds to 20°.
9. A flying method according to claim 1, and utilizing a reference speed of at least two values, (V2 +10 knots) or V2, these two values being interchangeable according to whether the .[.total gradient.]. .Iadd.potential flight path .Iaddend.γt is above or below a calibrated value.
10. A flying method according to claim 9, and utilizing a calibrated value of 6°.
11. A flying method according to claim 1, and including limiting the speed difference signal between positive and negative values.
14. A system according to .[.one of.]. claim 12, wherein the signal corresponding to the variation between the speed V and the reference speed comprises a subtractor which subtracts from the speed V, provided by an anemometer, a reference speed V2, the subtractor being connected to a first input of a second subtractor, having a second input connected to a source of reference voltage by means of a commutator controlled by a level detector of .[.total gradient.]. .Iadd.potential flight path .Iaddend.γt, the second .[.comparator.]. .Iadd.subtractor .Iaddend.being connected to a non-linear member which serves as a limiter and which provides said speed variation signal.
15. A system according to claim 12, wherein the level detection circuit which operates from the signal γt is composed of a low pass filter connected to an adder by a direct connection and by a derivation circuit comprising a differentiation and filtering member followed by a diode, the output of the adder being connected to a level detector which ensures the control of said commutator.
16. A system according to claim 15, wherein the level detector is adapted to control the non-linear member.
17. A system according to claim 12, wherein the .[.comparator.]. .Iadd.voter .Iaddend.is connected to a member for controlling the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar of the .[.artificial horizon.]. .Iadd.attitude director indicator .Iaddend.by means of a limiter.
18. A system according to claim 12, wherein, in order to ensure protection against abnormal negative .[.gradients.]. .Iadd.flight path angle .Iaddend.of the aircraft, due in particlar to poor operation of the flying system circuit providing the signal δ, the input of the .[.comparator.]. .Iadd.voter .Iaddend.corresponding to the signal δ2 is connected to a circuit providing a signal K (γa-γo), in which K is an adjustment gain and γo is a value representative of a minimum safety .[∇]. .Iadd.flight path angle.Iaddend., below which one should not drop.
19. A system according to claim 12, comprising a detection device serving to detect at the output of the .[.comparator.]. .Iadd.voter .Iaddend.the presence of the signal δ2 when the latter is selected, and to disconnect the entire flying system when the signal δ2 remains selected at the end of a predetermined period of time.
20. A system according to claim 12, wherein the input δ2 of the .[.comparator.]. .Iadd.voter .Iaddend.is connected to an amplifier for an adjustment gain K, which receives the differential signal of a subtractor, which in turn receives a signal representative of the aerodynamic .[∇]. .Iadd.flight path angle .Iaddend.γa and, a signal γo representative of a .[∇]. .Iadd.flight path angle .Iaddend.of predetermined value.
21. A system according to claim 20, wherein the predetermined .[∇]. .Iadd.flight path angle .Iaddend.value is 0.5°.
22. A system according to claim 20, wherein the amplifier for the gain K is connected to a subtractor which also receives the output signal of the .[.comparator.]. .Iadd.voter.Iaddend., and the differential signal provided by said subtractor is transmitted to a logic unit which acts on the .[.flight control horizon.]. .Iadd.attitude director indicator .Iaddend.in order to cause the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar to disappear when said differential signal remains zero for a period of time greater.[.,.]. than a predetermined period.
Description

The present invention relates to a flying method and system, in particular for take-off or overshoot of an aircraft, using the total .[.power.]. .Iadd.energy.Iaddend..

Generally, apart from radio guidance apparatus or the like, aircraft which are equipped for flying blind necessarily comprise three basic instruments, in particular as regards take-off, namely:

a flight control horizon which makes it possible to visualise the attitudes of the aircraft and which consequently provides information about the stability materialised by the position of .[.a model.]. .Iadd.an aircraft symbol .Iaddend.with respect to a moving sphere and the information of the .[.flight control device.]. .Iadd.attitude director indicator .Iaddend.which is visualised by .[.tendency.]. .Iadd.command .Iaddend.bars (the information of the .[.flight control device.]. .Iadd.attitude director indicator .Iaddend.being provided by a computer).

an anemometer which indicates the relative speed of the aircraft with respect to the air, and

a variometer which measures the vertical component of the speed, i.e. the ascending speed or descending speed.

Thus, by using these three instruments, the pilot is able to undertake all the take-off operations manually, even in the case of poor visibility (nevertheless complying with the minimum limits of visibility on the runway from the flight manual.)

Before effecting take-off, it is known that the pilot is in possession of a certain number of data relating to the aircraft, (mass, chosen position of the flaps), to the mechanics (occasional conditions such as de-icing etc. . . ), to the meteorological or geographical conditions, to possible obstacles on the ground and finally to the runway, these data enabling him to calculate three essential speeds for take-off, namely a speed V1, a speed V2 and a speed of rotation Vr which will be referred to hereafter.

When in possession of these data, the pilot may proceed with the aircraft take-off. To this end, .[.one.]. .Iadd.once .Iaddend.in the centre of the runway and after having reached an adequate engine condition which, generally, in the case of jet engines, corresponds to a maximum thrust or close to the latter, the .[.aircract.]. .Iadd.aircraft .Iaddend.begins the stage of travelling along the ground, during which its speed increases progressively.

In the case of a normal take-off, as soon as the aircraft reaches the speed of rotation Vr, the pilot actuates the .[.rudders and ailerons.]. .Iadd.elevators .Iaddend.and the aircraft takes-off several instants after at a speed Vloff. From the time when the aircraft has taken-off, the pilot must scan alternately the three afore-mentioned navigational appliances, in order to reach and maintain a speed V2 +10 knots or to maintain a .[.stable longitudinal position O1 .]. .Iadd.pitch attitude θ1.Iaddend. =18° which thus corresponds to a speed greater than V2 +10 knots.

If there is a breakdown in one of the engines at the time of take-off, the .[.proceedure.]. .Iadd.procedure .Iaddend.is then as follows:

In the case where, at the time of the breakdown, the aircraft has not yet reached the speed V1, the pilot must abandon take-off and must brake in order to stop the aircraft before the end of the runway.

In the case where the breakdown occurs after the aircraft has exceeded the speed V1, since the aircraft is no longer able to stop under normal conditions before the end of the runway, the pilot must continue take-off and, after rotation, the pilot must seek to reach and maintain the speed V2.

To obtain and maintain the speed V2 +10 knots, in the case of normal take-off, or the speed V2 in the case where an engine fails, the surveyance which the pilot must assume may be considerably simplified by using the .[.flight control device.]. .Iadd.attitude director indicator.Iaddend..

In this case, the pilot's work consists of bringing the .[.longitudinal tendency bar (pitching).]. .Iadd.pitch command bar of attitude .Iaddend.of the .[.flight control device.]. .Iadd.director indicator .Iaddend.to zero and of keeping it at zero. The movement of the longitudinal bar above or below the .[.model.]. .Iadd.aircraft symbol .Iaddend.thus indicating an order to .[.fly tail-down.]. .Iadd.pitch up .Iaddend.or an order to .[.dive.]. .Iadd.pitch down.Iaddend..

The invention thus relates to a method and system for flying an aircraft by means of .[.a flight control device.]. .Iadd.an attitude director indicator.Iaddend., in particular as regards take-off and overshoot.

The method thus uses the .[.total gradient.]. .Iadd.potential flight path .Iaddend.γt which represents the .[∇]. .Iadd.flight path angle .Iaddend.which the aircraft should assume to maintain its constant speed in rectilinear flight and which is expressed as follows: ##EQU1## in which: α is the .[.incidence.]. .Iadd.aerodynamic angle of attack .Iaddend.of the aircraft,

jx is the acceleration on the axis of bank,

g is the acceleration of gravity.

Generally, the measurement of γt may be undertaken in a computer, from the real .[.incidence.]. .Iadd.aerodynamic angle of attack .Iaddend.and an accelerometric box comprising at least two accelerometers whose perceptible axes are respectively parallel to the axis of bank (detection jx) and the axis of yaw (detection jz).

To this end, the information of a total .[.power.]. .Iadd.energy .Iaddend.variometer may be used to advantage, which provides in particular:

Information relating to the aerodynamic .[∇]. .Iadd.flight path angle .Iaddend.γa, (which is equal to the algebraic difference between the .[.stable position.]. .Iadd.pitch attitude .Iaddend.θ1 and the .[.incidence α (French Convention sign).]. .Iadd.aerodynamic angle of attack α.Iaddend..

Furthermore, it is known that aircraft are equipped more and more frequently for inertial navigation and consequently comprise a central inertial unit providing, inter alia, a signal representative of the ground speed Vs of the aircraft, from which may be calculated in a relatively simple manner, the value of the .[.total gradient.]. .Iadd.potential flight path .Iaddend.γt.

The invention therefore relates to the realisation of a circuit making it possible to calculate the .[.total gradient.]. .Iadd.potential flight path .Iaddend.γt from the speed Vs of the aircraft, whose value is provided by a central inertial unit provided in the aircraft, in order to prevent the use of a redundant accelerometric box.

To achieve this result, the invention uses mainly the following formulae, assuming zero side-slip of the aircraft and that the wings are horizontal, ##EQU2##

According to these relationships, the value of γt may be obtained by forming the sum of a magnitude proportional to the derivative with respect to time of the speed Vs and of the value of the aerodynamic .[∇]. .Iadd.flight path angle .Iaddend.γa which, as above-mentioned, may be obtained by forming in conventional manner, the algebraic difference between the .[.stable position.]. .Iadd.pitch attitude .Iaddend.θ1 and the .[.incidence.]. .Iadd.aerodynamic angle of attack .Iaddend.α.

It should be noted in this respect that in the simplest case and in the hypothesis where the calculation of the value of γt is not indispensable and does not serve to ensure the detection of an engine failure, the calculation of γa is not necessary. In fact, if we substitute the value ##EQU3## .[.of.]. .Iadd.for .Iaddend.the value γt, the control γd-γa=0 may be written successively in an equivalent manner: which consists of controlling the derivative dVs/dt by a magnitude proportional to the difference (V-V2).

According to a feature of the invention, in order to take into account the two previously mentioned imperatives, namely the speed V2 +10 knots (or V2 in the case of an engine failure) and the angle θ=18°, the error signal δ is transmitted to a .[.comparator.]. .Iadd.voter .Iaddend.which compares it with a difference signal δ1 proportional to (θ1 -18°) and to a reference signal δ2, corresponding for example to a constant order to .[.dive.]. .Iadd.pitch down .Iaddend.by 20° and which effects the control of the .[.longitudinal tendency.]. .Iadd.pitch command .Iaddend.bar of the .[.flight control device.]. .Iadd.attitude director indicator .Iaddend.by selecting from the signals δ, δ1, δ2, the signal whose instantaneous value is comprised between that of the other two.

The present invention also relates to the realisation of a device for detecting and protecting against active or dormant breakdowns, which could occur in the previously described flying system using total .[.power.]. .Iadd.energy.Iaddend., in particular when they are translated by values of the signal δ causing erroneous orders to .[.dive.]. .Iadd.pitch down.Iaddend., which lead the pilot to impart a dangerous negative aerodynamic .[∇]. .Iadd.flight path angle .Iaddend.to the aircraft.

In contrast to an active breakdown which becomes apparent and as soon as it appears disturbs the member controlled, possibly the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar of the .[.flight control horizon.]. .Iadd.attitute director indicator.Iaddend., a dormant breakdown is intended to mean a breakdown affecting a circuit which operates and acts on the controlled member only under certain particular conditions which only occur occasionally.

A breakdown occurring in the circuit for detecting engine failures in the aircraft may be mentioned as an example of a dormant breakdown in the flying system in question. For example, a breakdown of this circuit causing, under all conditions, the multi-engined state, will have an effect only when an engine failure actually occurs. It is thus clear that if they are not detected in time, such dormant breakdowns may have particularly serious consequences from the point of view of safety.

The object of the invention is therefore, on the one hand, in the case where an active breakdown occurs and in particular a breakdown occurring in the chain providing the error signal δ, to prevent the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar from leading to information less than a minimum safety .[∇]. .Iadd.flight path angle .Iaddend.and on the other hand, to indicate the existence of this breakdown in order that the pilot is not wrongly informed and may thus make the necessary corrections.

To achieve these results, the invention proposes a device providing protection against abnormal negative .[.gradients.]. .Iadd.flight path angles .Iaddend.of the aircraft due in particular to poor operation of the flying system providing the signal δ and which uses the properties of said .[.comparator.]. .Iadd.voter.Iaddend..

According to this device, the input δ2 of the .[.comparator.]. .Iadd.voter .Iaddend.is connected to a circuit providing a signal K .[.(γA-δo).]. (.Iadd.γa-γo.Iaddend.) in which K is an adjustment gain, γo is a value representing the minimum safety .[∇]. .Iadd.flight path angle.Iaddend., below which one should not descend.

According to another feature of the invention, this device also comprises a circuit making it possible, on the one hand, to detect at the output of the .[.comparator.]. .Iadd.voter.Iaddend., the presence of a signal .[.δ2.]. .Iadd.δ2 .Iaddend.when the latter is selected and on the other hand, to automatically disconnect the entire flying system when the signal δ2 remains selected at the end of a predetermined period of time.

A circuit of this type may comprise a subtractor arranged between the input δ2 and the output of the .[.comparator.]. .Iadd.voter .Iaddend.and a logic circuit comprising a delay which disconnects the flying system when the differential signal provided by said subtractor is cancelled and remains zero for a period of time greater than a predetermined period of time τ.

Embodiments of the invention will be described hereafter, as non-limiting examples, with reference to the accompanying drawings in which:

FIG. 1 is a circuit diagram making it possible to illustrate the principle of the flying system according to the invention, in the case where signals representative of the values γa and γt are available on board the aircraft;

FIG. 2 is a circuit diagram of a preferred embodiment of the system illustrated in FIG. 1;

FIGS. 3 to 14 show the dial of a flight control horizon during the successive stages of flying an aircraft corresponding successively:

to the normal take-off of a twin-engined aircraft in FIGS. 3, 4, 5, 6, 7 and 8,

to the take-off of a twin-engine aircraft with failure of one of the engines at the speed V1, FIGS. 9, 10 and 11,

to overshoot of a twin-engined plane, FIGS. 12, 13 and 14;

FIG. 15 is a circuit diagram making it possible to illustrate the principle of the flying system according to the invention, in its simplest version and in the case of an aircraft equipped with a central inertial unit;

FIG. 16 is a circuit diagram of a preferred embodiment of the system illustrated in FIG. 15;

FIG. 17 is a theoretical circuit diagram of a flying system using total .[.power.]. .Iadd.energy.Iaddend., equipped with a device for protecting against breakdowns;

FIG. 18 is a diagram representing: the trajectory of an aircraft in the course of take-off, during which poor operation of the flying system occurs;

the curve representative of the signal δ1 =(θ1 -18°) 1.5;

the curve representative of the signal δ;

the curve representative of the signal δ2 =K (γa-γo) and

the curve representative of a signal .[.δ2 .]. .Iadd.δ'2 .Iaddend.corresponding to an order to .[.dive.]. .Iadd.pitch down .Iaddend.by analogy with the signal δ2 mentioned in the main patent.

Depending on the indications of the .[.artificial horizon.]. .Iadd.attitude director indicator .Iaddend.3, the pilot may act on the controls 5 of the aircraft 1 in order to obtain said superimposition and the action of the pilot is translated by a modification of the parameters of the aircraft 1 and, in particular, of the speed V, of the .[.total gradient.]. .Iadd.potential flight path .Iaddend.γt and of the aerodynamic .[∇]. .Iadd.flight path angle .Iaddend.γa which, as above mentioned, are the three essential parameters used in the flying system using total .[.power.]. .Iadd.energy .Iaddend.according to the invention.

A signalδ provided at the output of the substractor 12 is transmitted to a .[.comparator.]. .Iadd.voter .Iaddend.14 (whose function will be explained hereafter), which controls the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar 2 of the .[.artificial horizon.]. .Iadd.attitude director indicator .Iaddend. 3.

In this respect, it will be noted that one of the considerable advantages of this system consists in that it combines accelerometric information (calculation of γt) and angular information (calculation of γa) with anemometric information (calculation V) which react conversely at the time of .[.squalls.]. .Iadd.gusts.Iaddend.. Due to this, at the time of squalls, the differences exhibited by the accelerometric and angular information are compensated for by differences in the anemometric information and consequently, the flying system is not subject to any considerable disturbances.

If the operation of the .[.comparator.]. .Iadd.voter .Iaddend.14 is not taken into consideration and the latter is replaced by a simple electrical connection to the member for controlling the .[.pitching.]. .Iadd.pitch command .Iaddend.bar 2 of the .[.flight control horizon.]. .Iadd.attitude director indicator .Iaddend.3, the indications of the bar 2 may be interpreted as follows:

in the case where the signal δ provided by the subtractor 12 is positive, the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar 2 is located below the .[.model.]. .Iadd.aircraft symbol .Iaddend.4, which corresponds to an order to .[.dive.]. .Iadd.pitch down.Iaddend..

The state δ>o signifies that γa is greater than γd and, consequently, inter alia, may result from two situations of the following types:

a. In the case where V=V reference and where γa>γt, the .[.pitching.]. .Iadd.pitch command .Iaddend.bar 2 of the flight control device thus indicates an order to .[.dive.]. .Iadd.pitch down .Iaddend.to restore γa to the value of γt in order to prevent deceleration;

In the case where the signal δ provided by the subtractor 12 is negative, the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar 2 is located above the .[.model.]. .Iadd.aircraft symbol .Iaddend.4, which corresponds to an order to .[.fly tail-down.]. .Iadd.pitch up.Iaddend..

The state .[.δ>o.]. .Iadd.δ<o .Iaddend.signifies that γa is smaller than γd and consequently, inter alia, may result from two situations of the following types:

c. In the same where V=V reference and where .[.γa>γt.]. .Iadd.γ<γt.Iaddend., the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar 2 thus indicates an order to .[.fly tail-down.]. .Iadd.pitch up .Iaddend.in order to restore γa to the value of γt in order to prevent acceleration;

.Iadd.d. .Iaddend.In the case where V>V reference and γa=γt, the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar thus indicates an order to .[.fly tail-down.]. .Iadd.pitch up .Iaddend.in order to momentarily increase γa in order to restore V to V reference.

As aforementioned, the operation which has been described does not take into account the action of the .[.comparator.]. .Iadd.voter .Iaddend.14.

This .[.comparator.]. .Iadd.voter .Iaddend.14 comprises three inputs, whereof one 15 is connected to the subtractor 12, the second 16 is connected to a circuit providing a signal δ1 proportional to the difference θ1 -18° (θ1 being the .[.stable position.]. .Iadd.pitch attitude.Iaddend.), whereas the third 17 receives a reference signal δ2 corresponding in the embodiment described to an order to dive by 20°.

The function of the comparator 12 is to compare the signals applied thereto by its three inputs and to transmit to the member for controlling the .[.pitching.]. .Iadd.pitch command .Iaddend.bar 2 of the .[.flight control horizon.]. .Iadd.attitude director indicator .Iaddend.3, the signal whose value, at a given time, is comprised between the value of the two other signals, at the same time.

With reference to FIG. 2, which is a circuit diagram of one embodiment of the invention, the signal γt is transmitted by the intermediary of a low pass filter 21, to a subtractor 22 whose second input 23 receives the signal γa which, as above-mentioned, is equal to (θ1 -α). Consequently, this signal γa is obtained by means of a subtractor 24 which receives, on the one hand, a signal representative of θ1 and, on the other hand, a signal representative of the .[.incidence.]. .Iadd.aerodynamic angle of attack .Iaddend.α which is filtered by means of a low pass filter 25. In order to take into account the conditions of travelling on the ground and flying conditions, and in order to prevent dragging errors due to high angular speeds of α during rotation, the circuit providing the signal representative of α comprises a commutation device which makes it possible:

on the one hand, to transmit to the subtractor 24 .[.whilst travelling on the ground.]. .Iadd.during the take off roll .Iaddend.with the .[.under-carriage collapsed.]. .Iadd.main under-carriage compressed .Iaddend.(for example by means of a relay 26 controlled by a detector associated with the under-carriage) a signal representative of the .[.stable position.]. .Iadd.pitch attitude .Iaddend.θ1 in place of the .[.incidence.]. .Iadd.aerodynamic angle of attack .Iaddend.α in order to obtain a virtually zero signal at the output of the subtractor 24.

and, on the other hand, to modify the time constant of the low pass filter 25 at the time of take-off. The signal supplied by the subtractor 22 is transmitted to a subtractor 27 after passing through an adaptation amplifier (unit 28). The second input 29 of this subtractor receives a signal depending on the difference V--V reference. The circuit which makes it possible to obtain this signal is composed firstly of a subtractor 31 which forms the difference (V-V2), of a low pass filter 32 connected to the output of the subtractor 31, which serves mainly for filtering the speed signal V indicated. This low pass filter 32 is connected to a subtractor 33 which receives, through the intermediary of a commutation circuit, a constant signal corresponding to 10 knots. This commutation system may be constituted by a double commutator 34 and 34b with a low pass filter 35 controlled by a level detector 36 in order to eliminate the signal 10 knots in the case of an engine failure. The level detection circuit which operates after the signal γt is composed of a low pass filter 37 connected to an adder 38, on the one hand, by a direct connection 39 and, on the other hand, by a derivation circuit comprising a differentiation and filtering member 40 followed by a diode 41 connected to the adder 38. This derivation is particularly provided in order to increase the sensitivity of the level detector 36 when the derivative of γt is positive. The output of the adder 38 is connected to the level detector 36 which intervenes as above explained in the control logic of the double commutator 34 and 34b.

In particular, .[.in the configuration where the leading edges are protruding.]. .Iadd.when the aircraft is not in clean configuration .Iaddend.(in the case of take-off), the operation of this level detector 36 is such that when the output of the adder 38 exceeds a calibrated value, for example 6°, the level detector 36 acts on the double commutator 34 and 34b to send the 10 knots signal to the subtractor 33. Conversely, when the value of γt drops below the value 6°, the level detector 36 acts on the double commutator to interrupt the 10 knots signal.

When the .[.leading edges are retracted.]. .Iadd.aircraft is in clean configuration.Iaddend., the state of the level detector 36 has no effect on the commutator 34; in this case, the system is always governed by the indication V reference, which means that it can be used when cruising, in order to acquire the speeds indicated.

The subtractor 33 is connected by means of an amplifier 43 to a non-linear member 42, which serves as an .[.amplifier.]. .Iadd.amplitude .Iaddend.limiter. The values of the amplitude limitations of 42 determine the rate of acceleration or deceleration imposed by the system during the stages of .[.picking up.]. .Iadd.acquiring .Iaddend.the reference speed. The selector 53 makes it possible to modify the value of the limitations of 42 depending on the state of the level detector 36, on the position of the leading edges (and possibly on the condition "aircraft on the ground")

In particular, .[.in the configuration of the leading edges projecting.]. .Iadd.when the aircraft is not in clean configuration .Iaddend.(case of take-off), the level detector 36 imposes wide limits when the output of the adder 38 exceeds 6° or narrow limits when γt drops below 6°. In this latter case, the limitation of negative sign is adjusted in order to constitute an implicit protection of the system against trajectories having a negative .[.slope.]. .Iadd.flight path angle .Iaddend.subsequent to an engine failure on take-off. .[.In the configuration of the leading edges retracted.]. .Iadd.In clean aircraft configuration .Iaddend.(use during cruising) the level detector 36 is inoperative and wide limitations are imposed.

The output of the non-linear member 42 is connected to the subtractor 27 by means of a low pass filter 44.

The subtractor 27 is connected to the .[.comparator.]. .Iadd.voter .Iaddend.45, also by means of an adaptation amplifier 46.

At its second input 47, the .[.comparator.]. .Iadd.voter .Iaddend.45 receives a signal proportional to (θ1 -18°) obtained by means of a subtractor 48 receiving the signal θ1 and the signal corresponding to 18°. This subtractor 48 is connected to the .[.comparator.]. .Iadd.voter .Iaddend.45 by means of an adaptation amplifier 49.

The third input 50 of the .[.comparator.]. .Iadd.voter .Iaddend.45 receives a constant reference signal corresponding for example to an order of 20°.

The output of the .[.comparator.]. .Iadd.voter .Iaddend.45 is connected to the member for controlling the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar of the .[.flight control device.]. .Iadd.attitude director indicator .Iaddend.by means of an adaptation amplifier 51 and a limiter 52 which makes it possible to authorise maximum displacement of the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar.

FIGS. 3 to 14 make it possible to illustrate the method of operation of the afore-described system, in particular as regards take-off and overshoot.

In a preliminary stage of take-off, before releasing the brakes, the pilot must record the speed V2 which was previously established on the speed module of the control station of the automatic flying system. Then the system is activated by the action of the pilot (for example: .[.engagement lever or blades of the throttle levers.]. .Iadd.palm switch 6f of the throttle levers, .Iaddend.etc.)

The .[.artificial horizon.]. .Iadd.attitude director indicator .Iaddend.thus has the configuration illustrated in FIG. 3, in which the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar gives an order to .[."dive,".]. .Iadd.pitch down.Iaddend., since only the limited term (V-V2) is different from zero.

During rotation, owing to the fact that the .[.total gradient.]. .Iadd.potential flight path .Iaddend.decreases and the aerodynamic .[∇]. .Iadd.flight path angle .Iaddend.increases, the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar begins to drop and this movement is accentuated upon take-off (FIG. 5.)

The pilot must then act on the flying controls in order to maintain the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar in coincidence with the .[.model.]. .Iadd.aircraft symbol .Iaddend.(FIG. 6).

either the .[.retention.]. .Iadd.holding .Iaddend.of V2 +10 knots (θ1 being less than 18°);

Once this flying configuration has been achieved, the pilot may proceed to accelerate from V2 +10 (or from θ1 =18°) with the hyperlift devices retracted.

For this, as soon as the aircraft reaches a predetermined altitude, since it is stabilised at V2 +10 knots (or at θ1 =18°), the pilot indicates a speed greater than V2 on the speed module (for example 250 knots).

As soon as this speed is indicated, the system controls a constant acceleration which may correspond for example to keeping γa to 4° below γt. This control is carried out, even under transitory operating conditions, such as the retraction of the hyperlift members.

Thus, at the beginning of acceleration, the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar gives an order to dive, i.e. an order to vary the .[.stable position.]. .Iadd.pitch attitude .Iaddend.in order to bring γa to 4° below γt (FIG. 7).

After the retraction of the flaps, the .[.pitching tendecy.]. .Iadd.pitch command .Iaddend.bar gives an order to .[.fly tail-down.]. .Iadd.pitch up .Iaddend.in order to adjust the position of equilibrium, subsequent to the retraction of the flaps (FIG. 8).

The operation of the flying system according to the invention will now be studied in the case where there is a failure of one of the aircraft engines, which could occur either at the time of take-off, after reaching the speed V1, or durng the stage succeeding take-off and up to the stabilisation of the aircraft at a speed of V2 +10 knots (or θ1 =18°).

In the case where the engine failure occurs between the speed V1 and take-off, at the time of the failure, the drop of γt is preponderant with respect to the reduction of control over the speed variation and consequently the bar drops (FIGS. 9 to 10) which has the effect of warning the pilot of the failure and thus of the precautions which he must take at the time of rotation. At the time of rotation and take-off, the .[.tendency.]. .Iadd.command .Iaddend.bar continues to drop.

The pilot must act on his flying controls to bring and maintain the .[.tendency.]. .Iadd.command .Iaddend.bar in coincidence with the .[.model.]. .Iadd.aircraft symbol.Iaddend., which corresponds to achieving and maintaining the speed V2 (FIG. 11).

As afore-mentioned, the application of the flying system according to the invention is not limited to take-off manoeuvres. This system may also be suitable in the event of an overshoot during an approach.

In this case, the pilot no longer displays V2 but a reference speed for the approach: V ref. On the speed module of the automatic flying system. When the pilot initiates the operation of overshoot he actuates .[.blades.]. .Iadd.palm switches.Iaddend., implanted on the throttle lever and provided for this operation. This action returns to the logic of the system, which may thus be used for controlling this stage.

Thus, at the time of rotation, which may be carried out manually or by following the flight control device, the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar is virtually zero (during a rotation at average speed) (FIG. 12).

After rotation of the .[.stable position.]. .Iadd.pitch attitude.Iaddend., by means of the system according to the invention, the pilot is able to .[.pick-up.]. .Iadd.acquire .Iaddend.and maintain a speed of V ref.+10 knots (or θ1 =18°) (FIGS. 13-14.).

In this respect, in the case of a strong head wind, for example greater than 10 knots, it will be noted that the flying .[.proceedure.]. .Iadd.procedure .Iaddend.may command the display on the speed module of the flying system of: V ref plus an increase depending on the wind. It is thus this reference value (with or without 10 knots) which will thus serve for the system.

In the case where an engine failure occurs during overshoot, the control takes place at V ref in place of V ref+10 knots, as previously.

This transformation takes place from the beginning if the engine failure occurs at the initial time of overshoot; or as soon as γt<6 degrees if the failure occurs several instants after overshoot.

Finally, the flying system according to the invention may be used to acquire a cruising speed. In this case, in a manner similar to the preceding, the speed to be attained is displayed on the speed module of the automatic pilot system for example. The indications of the .[.pitching.]. .Iadd.pitch command .Iaddend.bar make it possible to achieve and maintain the reference speed displayed. Cruising is distinguished by the system by the condition of .[.leading edges retracted (or any equivalent condition).]. .Iadd.aircraft in clean configuration.Iaddend..

Finally, it will be noted that the system according to the invention facilitates more flexible and more reliable piloting of the aircraft. In particular, it makes it possible to pass asymptotically from one speed to another (for example from the speed V2 +10 knots to the speed V2 during an engine failure) and this is without any oscillation.

With reference to FIG. 15, the flying system of the aircraft acts in a similar manner to that previously described, on the pitching bar 62 of the .[.flight control horizon.]. .Iadd.attitude director indicator .Iaddend.63.

As previously mentioned, the aircraft shown diagrammatically with its control members, respectively by the units 64 and 65, comprises a central inertial unit providing a signal representative of the speed Vs. After having been derived with respect to time and multiplied by a coefficient equal to 1/g, this signal Vs is transmitted to an adder 66 which receives, at its second input 67, a signal corresponding to the difference between the anemometric speed V and a reference value .[.V2.]. .Iadd.V2.Iaddend..

In the example shown, the speed signal V coming from an anemometer is transmitted to a subtractor 68 which receives, at its second input 69, a signal corresponding to the reference speed.

The signal resulting from this difference is amplified (unit 70) and is then transmitted to an amplitude limiter 71 which carries out a variation limitation. The amplitude limiter 71 is connected to the adder 66, whose output is connected to a .[.comparator.]. .Iadd.voter .Iaddend.72 which controls the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar 62 of the .[.flight control horizon.]. .Iadd.attitude director indicator .Iaddend.63.

The .[.comparator.]. .Iadd.voter .Iaddend.72 comprises three inputs, whereof one 75 is connected to the adder 66, the second input 76 is connected to a circuit providing a reference signal δ1 proportional to the difference θ1 -18° (θ1 being the .[.stable position,).]. .Iadd.pitch attitude).Iaddend., whereas the third input 77 receives a reference signal δ2 corresponding in the embodiment described to an order of 20°.

The function of the .[.comparator.]. .Iadd.voter .Iaddend.72 is to compare the signals which are sent to the latter on its three inputs and to transmit to the member for controlling the .[.pitching.]. .Iadd.pitch command .Iaddend.bar 62 of the .[.flight control horizon.]. .Iadd.attitude director indicator .Iaddend.63, the signal whose value, at a given time, is comprised between the value of the two other signals, at the same time.

The operation of the system which has been described is strictly identical to that relating to FIG. 1 and consequently will not be described again.

As shown in FIG. 16, the wiring diagram of the flying system is identical to that shown in FIG. 2, apart from the fact that instead of using a signal γt provided by an accelerometric box provided for this purpose, this signal γt is calculated from the speed signal Vs supplied by a central inertial unit 80.

This signal Vs coming from the central unit 80 is transmitted to a .[.shunting device.]. .Iadd.derivation circuit .Iaddend.81 also acting as a .[.low.]. .Iadd.high .Iaddend.pass filter for the transfer function s/1+s (s being the Laplace operator) and whose output is connected to an amplifier 82 for the gain 1/g.

This amplifier 82 is in turn connected to an adder 83 which at its second input also receives a signal representative of the .[∇]. .Iadd.flight path angle .Iaddend.γa. This signal γa is obtained by forming the difference θ-α in a conventional manner, by means of a subtractor 84, after having filtered the signals θ and α by means of low pass filters 85 and 85' having a transfer function 1/1+s.

The signal γt provided by the adder 83 is transmitted to the low pass filters 21 and 37 of an identical circuit to that illustrated in FIG. 2. This circuit whose various parts have the same references as those in FIG. 2, will not be described again.

The signal δ, provided at the output of the subtractor 93 is transmitted to a .[.comparator.]. .Iadd.voter .Iaddend.95, which controls the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar of the .[.flight control horizon.]. .Iadd.attitute director indicator .Iaddend.87.

The signal provided by the amplifier 99 is also transmitted to a subtractor 102 which also receives the output signal of the .[.comparator.]. .Iadd.voter .Iaddend.95.

The differential signal provided by the subtractor 102 is transmitted to a logic unit 103 which acts on the .[.flight control horizon.]. .Iadd.attitude director indicator .Iaddend.87, in order to cause the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar to disappear when said differential signal remains zero after a period of predetermined time, for example five seconds.

It must be stressed that the diagram which has been described is a very simplified theoretical diagram and that it may clearly be completed by all the devices afore-described. In particular, it may be equipped with a commutation device controlled by a circuit for detecting an engine failure, which enables the reference speed to assume two values, namely the value (V2 +10 knots) in the case of normal operation or the value V2 in the case of an engine failure.

The diagram illustrated in FIG. 18 makes it possible to illustrate the action of the afore-described protection device, during the take-off of an aircraft, in the course of which a failure or poor operation of the flying system occurs during take-off.

This diagram firstly illustrates the trajectory of the aircraft, a curve which comprises:

a first part 105 corresponding to the phase during which the pilot seeks to reach the speed V2 +10 knots or the .[.position.]. .Iadd.pitch attitude .Iaddend.θ1 =18°;

a second part 106 corresponding to .[.maintaining.]. .Iadd.holding .Iaddend.a speed of V2 +10 knots;

a third part 107 in which are located:

in broken line the trajectory 108 corresponding to normal operation,

in full line the trajectory 104 corresponding to the system according to the main patent disturbed by poor operation,

in broken line the trajectory 109 corrected by the protection action according to the invention in the case of poor operation.

This diagram also shows, in correlation with the trajectories 104, 108 and 109, the curves 111, 112 and 113 representing the signals δ1, δ and δ2, at the input of the .[.comparator.]. .Iadd.voter .Iaddend.and a horizontal line 114 corresponding to a constant signal δ'2 indicating a considerable and constant order to .[.dive.]. .Iadd.pitch down .Iaddend..

It will thus be seen that after reaching the speed V2 +10 knots or θ1 =18°, during the phase of .[.maintaining.]. .Iadd.holding .Iaddend. this speed, the signals δ1, δ and δ2 are maintained at a substantially constant level and in the order δ2 >δ>δ1 and, consequently, it is the signal δ which is selected (curve shown in broken line 115). As soon as a failure occurs in the circuit producing the signal δ and which is translated by an order to .[.dive.]. .Iadd.pitch down .Iaddend.and by a decrease in the .[∇]. .Iadd.flight path angle .Iaddend.of the aircraft, the value of the signal δ2 =(γo-γa) K decreases and, at the end of a certain period of time, becomes less than the value of the signal δ. In other words, there is an intersection of the curves δ and δ2. Beyond the point where the value of δ2 becomes less than the value of δ, the value of δ2 is located between that of δ and δ1 and due to this, it is the signal δ2 which is selected (see curve shown in broken line 115).

Consequently, the flying system is governed by K(γa-γo), which makes it possible to maintain an aerodynamic .[∇]. .Iadd.flight path angle .Iaddend.γa .[.and.]. .Iadd.equal to .Iaddend.γo. In practice the value of γo which constitutes the bottom aerodynamic .[∇]. .Iadd.flight path angle.Iaddend., below which one should not fall, is .[.approximtely.]. .Iadd.approximately .Iaddend.0.5°.

Nevertheless, it is important to state that this bottom aerodynamic .[∇]. .Iadd.flight path angle .Iaddend.corresponds to an abnormal flying configuration and consequently should not be maintained beyond a predetermined period of time. This is the reason why the logic unit 103 shown in FIG. 17 acts in a manner to cause the .[.pitching tendency.]. .Iadd.pitch command .Iaddend. bar of the .[.flight control horizon.]. .Iadd.attitude director indicator .Iaddend.87 to disappear, when the signal δ2 is selected at the end of a predetermined time, for example five seconds.

It is clear that this disappearance of the .[.pitching tendency.]. .Iadd.pitch command .Iaddend.bar on the flight control horizon 87 warns the pilot of a failure existing in the flying system and causes him to undertake the necessary corrections immediately.

Although the present invention has been described with reference in particular to take-off and overshoot of an aircraft, it is equally applicable to any other situation in which the throttle of an aircraft is acted upon to increase the speed of the aircraft.

Patent Citations
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US3681580 *Aug 1, 1969Aug 1, 1972Teledyne IncRotation,climbout,and go-around control system
US3691356 *Dec 10, 1970Sep 12, 1972Sperry Rand CorpSpeed command and throttle control system for aircraft
US3800127 *Sep 11, 1972Mar 26, 1974Us Air ForceVertical rate reference computer for control of a vtol or stol aircraft at a constant flight path angle
US3822047 *Dec 14, 1972Jul 2, 1974Collins Radio CoTakeoff and go-around climb-out profile pitch command formulation for aircraft
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US8290639Nov 4, 2010Oct 16, 2012The Boeing CompanyManaging control surfaces for an aircraft
EP0120855A1 *Sep 30, 1982Oct 10, 1984Boeing CoTotal energy based flight control system.
Classifications
U.S. Classification701/15, 244/181
International ClassificationG06G7/78, G05D1/06
Cooperative ClassificationG06G7/78, G05D1/0607
European ClassificationG05D1/06B, G06G7/78