Flying method and system using total energy for an aircraft

Abstract

A flying method using total power energy, in particular for the take-off and overshoot of an aircraft, is disclosed in which the aerodynamic .[∇]. flight path angle γa is governed by reference to a desired .[∇]. flight path angle γd which is the total gradient potential flight path γt modulated by the difference between the aircraft speed v and a reference speed v₂. An error signal δ representative of the difference between the aerodynamic .[∇]. flight path angle γa and the desired .[∇]. flight path angle γd is displayed. The display of the desired .[∇]. flight path angle γd may be by means of the pitching tendency bar of the artificial horizon pitch command bar of an attitude director indicator for example.

Description

The present invention relates to a flying method and system, in particular for take-off or overshoot of an aircraft, using the total (longitudinal tendency bar) (pitch command bar) of a flight control horizon 3 of convential conventional type. It is known that the position of this tendency command bar 2, with respect to the model aircraft symbol 4, makes it possible to give the pilot either an order to dive pitch down or an order to fly tail-down or even, when the tendency pitch command bar 2 is superimposed on the model aircraft symbol 4, that the aircraft 1 is in the desired configuration.

Depending on the indications of the artificial horizon attitude director indicator 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 potential flight path γt and of the aerodynamic .[∇]. flight path angle γa which, as above mentioned, are the three essential parameters used in the flying system using total power energy according to the invention.

The total gradient potential flight path signal γt which may be obtained by a computer, from the actual incidence aerodynamic angle of attack and two accelerometers whose perceptible axes are respectively parallel to the axis of bank (direction Jx) and the axis of yaw (detection Jz) is transmitted to an adder 6 which receives, at its second input 7, a signal corresponding to the variation of the speed V with respect to a reference value, for example V₂ +10 knots or even only V₂. In the example illustrated, the speed signal V coming from an anemometer is transmitted to a subtractor 8 which receives, at its second input 9, a signal corresponding to the reference speed. The signal resulting from this difference is amplified (unit 10) and is then transmitted to an amplitude limiter 11 which effects a limitation of the minimum and maximum values of the difference. The amplitude limiter 11 is connected to the adder 6 whose output is connected to a subtractor 12 which receives a signal corresponding to the aerodynamic .[∇]. flight path angle γa at its second input 13.

A signalδ provided at the output of the substractor 12 is transmitted to a comparator voter 14 (whose function will be explained hereafter), which controls the pitching tendency pitch command bar 2 of the artificial horizon attitude director indicator 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 gusts. 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 voter 14 is not taken into consideration and the latter is replaced by a simple electrical connection to the member for controlling the pitching pitch command bar 2 of the flight control horizon attitude director indicator 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 pitch command bar 2 is located below the model aircraft symbol 4, which corresponds to an order to dive pitch down.

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 pitch command bar 2 of the flight control device thus indicates an order to dive pitch down to restore γa to the value of γt in order to prevent deceleration;

b. In the case where V>V V<Vreference and γa=γt, the pitching pitch command bar 2 of the flight control device attitute director indicator then indicates an order to dive pitch down to momentarily reduce γa in order to restore V to V ref.

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

The state δ>o δ<o 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 γ<γt, the pitching tendency pitch command bar 2 thus indicates an order to fly tail-down pitch up in order to restore γa to the value of γt in order to prevent acceleration;

d. In the case where V>V reference and γa=γt, the pitching tendency pitch command bar thus indicates an order to fly tail-down pitch up 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 voter 14.

This comparator voter 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 pitch attitude), 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 pitch command bar 2 of the flight control horizon attitude director indicator 3, the signal whose value, at a given time, is comprised between the value of the two other signals, at the same time.

It should be noted that when the signal received by the input 16 (or stable position pitch attitude signal) is selected, if θ₁ >18°, the pitching tendency pitch command bar 2 gives an order to dive pitch down, on the other hand, if θ₁ >18°θ₁ <18°, the pitching tendency pitch command bar 2 gives an order to fly tail-down pitch up.

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 aerodynamic angle of attack 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 during the take off roll with the under-carriage collapsed main under-carriage compressed (for example by means of a relay 26 controlled by a detector associated with the under-carriage) a signal representative of the stable position pitch attitude θ₁ in place of the incidence aerodynamic angle of attack 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 when the aircraft is not in clean configuration (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 aircraft is in clean configuration, 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 amplitude 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 acquiring 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 when the aircraft is not in clean configuration (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 flight path angle subsequent to an engine failure on take-off. In the configuration of the leading edges retracted In clean aircraft configuration (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 voter 45, also by means of an adaptation amplifier 46.

At its second input 47, the comparator voter 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 voter 45 by means of an adaptation amplifier 49.

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

The output of the comparator voter 45 is connected to the member for controlling the pitching tendency pitch command bar of the flight control device attitude director indicator by means of an adaptation amplifier 51 and a limiter 52 which makes it possible to authorise maximum displacement of the pitching tendency pitch command 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 palm switch 6f of the throttle levers, etc.)

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

Whilst travelling on the ground During the take-off roll, as soon as the brakes are released and the take-off thrust established, the total gradient potential flight path γt increases (approximately 12°) and, consequently, the leve level detector 36 causes the selection of V₂ +10 knots. The tendency command bar which is located approximately in the upper position thus indicates an order to fly tail-down pitch up and is maintained in this position approximately until rotation (FIG. 4).

During rotation, owing to the fact that the total gradient potential flight path decreases and the aerodynamic .[∇]. flight path angle increases, the pitching tendency pitch command 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 pitch command bar in coincidence with the model aircraft symbol (FIG. 6).

Maintaining the position of the tendency bar firstly ensures the attainment acquisition of an acceleration at to V₂ +10 knots then;

either the retention holding of V₂ +10 knots (θ₁ being less than 18°);

or the retention holding of (θ₁ at 18°) (with acceleration).

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 pitch command bar gives an order to dive, i.e. an order to vary the stable position pitch attitude in order to bring γa to 4° below γt (FIG. 7).

After the retraction of the flaps, the pitching tendecy pitch command bar gives an order to fly tail-down pitch up 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 failure occurs once the aircraft has been stabilised, after take-off at the speed V₂ +10 knots, the system automatically selects (by means of the level detector 36) the speed V₂ as the reference speed. At this time, the total gradient potential flight path γt is less than the .[∇]. flight path angle γa and gives an order (a) to dive pitch down, whereas the speed variation signal which passes from V-(V₂ +10) to (V-V₂) gives an order (b) to fly tail-down pitch up. In this case, the order (a) is preponderant and consequently the pitching pitch command bar gives an order to dive pitch down. The pilot consequently acts to bring the pitching tendency pitch command bar into coincidence with the model aircraft symbol in order to achieve and maintain the configuration γt=γa and V=V₂.

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 command bar continues to drop.

The pilot must act on his flying controls to bring and maintain the tendency command bar in coincidence with the model aircraft symbol, 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 palm switches, 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 pitch command bar is virtually zero (during a rotation at average speed) (FIG. 12).

After rotation of the stable position pitch attitude, by means of the system according to the invention, the pilot is able to pick-up acquire 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 procedure 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 pitch command 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) aircraft in clean configuration.

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 attitude director indicator 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 V₂.

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 voter 72 which controls the pitching tendency pitch command bar 62 of the flight control horizon attitude director indicator 63.

The comparator voter 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,) pitch attitude), 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 voter 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 pitch command bar 62 of the flight control horizon attitude director indicator 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 derivation circuit 81 also acting as a low high 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 .[∇]. flight path angle 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.

In the system for flying the aircraft illustrated, with its controls, by the blocks 86 and 86' (FIG. 17), the signal of total gradient potential flight path γt which may be obtained in conventional manner from the actual incidence aerodynamic angle of attack and two accelerometers, is transmitted to an adder 88 which receives, at its second input, a signal corresponding to the speed variation V with respect to a reference value, for example V₂ +10 knots or even only V₂. To this end, the speed signal V, coming from an anemometer, is transmitted to a subtractor 89 which receives, at its second input 90, a signal V₂ corresponding to the reference speed. The differential signal, provided by the subtractor 89, is amplified (unit 91) and is then transmitted to an amplitude limiter 92 which carries out a limitation of the minimum and maximum values of the variation. The amplitude limiter 92 is connected to the adder 88 whose output is connected to a subtractor 93, which receives a signal corresponding to the aerodynamic .[∇]. flight path angle γa at its second input 94.

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

The comparator voter 95 comprises three inputs, whereof one 96 is connected to the subtractor 93, the second 97 is connected to a circuit providing a signal δ₁ proportional to the difference θ₁ -18° (θ₁ being the longitudinal stable position pitch attitude, whereas the third input 98 receives a signal K (γa-γo) emanating from an amplifier 99 for an adjustment gain K, which receives the differential signal of a subtractor 100. On the one hand, this subtractor 100 receives a signal representative of the aerodynamic .[∇]. flight path angle γa taken for example from 94 and on the other hand, a signal γo representative of a gradient an angle of predetermined value, for example 0.5°.

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

The differential signal provided by the subtractor 102 is transmitted to a logic unit 103 which acts on the flight control horizon attitude director indicator 87, in order to cause the pitching tendency pitch command 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 pitch attitude θ₁ =18°;

a second part 106 corresponding to maintaining holding 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 voter and a horizontal line 114 corresponding to a constant signal δ'₂ indicating a considerable and constant order to dive pitch down .

It will thus be seen that after reaching the speed V₂ +10 knots or θ₁ =18°, during the phase of maintaining holding 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 pitch down and by a decrease in the .[∇]. flight path angle 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 .[∇]. flight path angle γa and equal to γo. In practice the value of γo which constitutes the bottom aerodynamic .[∇]. flight path angle, below which one should not fall, is approximtely approximately nd∅5°.

Nevertheless, it is important to state that this bottom aerodynamic .[∇]. flight path angle 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 pitch command bar of the flight control horizon attitude director indicator 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 pitch command 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.

Claims

1. A flying method using total power energy, for an aircraft, and comprising the following steps, governing the aerodynamic .[∇]. flight path angle γa by reference to a desired .[∇]. flight path angle γd, obtaining the desired .[∇]. flight path angle γd by modulating the total gradient potential flight path γt by the difference between the aircraft speed v and a reference speed v₂, and displaying an error signal δ representative of the difference between the aerodynamic .[∇]. flight path angle γa and the desired .[∇]. flight path angle γd.

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 pitch command bar of an artificial horizon attitude director indicator.

3. A flying method according to claim 1, and further including calculating the total gradient potential flight path γt from an accelerometric box.

4. A flying method according to claim 1, and further including calculating the total gradient potential flight path 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.

7. A flying method according to claim 1, and further including taking into account either the variation δ, intended to subject the speed v to the control of the reference value, or a variation of the stable position pitch attitude with respect to a reference position, and including transmitting the error signal δ to a comparator voter, comparing δ in the comparator voter with a signal δ₁ corresponding to said variation of position attitude and with a constant reference signal δ₂ and controlling the longitudinal tendency pitch command bar of the artificial horizon attitude director indicator by the intermediate signal of the group (δ, δ₁, δ₂) i.e. the signal whose instantaneous value is between those of the other two signals.

8. A flying method according to claim 7, and utilizing a reference position attitude 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 potential flight path 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.

12. A system for applying a flying method using total power energy, for an aircraft, and comprising the steps of governing the aerodynamic gradient δa flight path angle γa by reference to a desired gradient δd flight path angle γd, obtaining the desired gradient δd flight path angle γd by modulating the total gradient δt potential flight path angle γc by the difference between the aircraft speed B v and a reference speed v₂, and displaying an error signal δ representative of the difference between the aerodynamic gradient δa flight path angle γa and the desired gradient δd flight path angle γd, and wherein a signal representative of the total gradient δt potential flight path angle γc is transmitted to a first input of an adder which receives at a second input a signal corresponding to the difference between the speed v and a reference speed, the adder output being connected to a first input of a subtractor which receives at a second input, a signal corresponding to the aerodynamic graident δa flight path angle γa, the subtractor output being transmitted to a comparator voter which is connected to a member for controlling the pitching tendency pitch command bar of the artificial horizon attitude director indicator, and the comparator voter also receiving a signal representative of a variation of stable position pitch attitude and a reference signal.

13. A system according to claim 12, wherein the signal representative of the aerodynamic .[∇]. flight path angle is provided by a circuit which forms the difference between the stable position pitch attitude θ₁ and the incidence aerodynamic angle of attack α, this circuit being able to comprise a commutation device controlled by a signal representing the state of the under-carriage main undercarriage, in order to substitute the stable position pitch attitude θ₁ for the incidence aerodynamic angle of attack αwhen said under-carriage main undercarriage is collapsed compressed, in order to cancel out said difference.

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 potential flight path γt, the second comparator subtractor 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 voter is connected to a member for controlling the pitching tendency pitch command bar of the artificial horizon attitude director indicator by means of a limiter.

18. A system according to claim 12, wherein, in order to ensure protection against abnormal negative gradients flight path angle of the aircraft, due in particlar to poor operation of the flying system circuit providing the signal δ, the input of the comparator voter 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 .[∇]. flight path angle, 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 voter 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 voter 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 .[∇]. flight path angle γa and, a signal γo representative of a .[∇]. flight path angle of predetermined value.

21. A system according to claim 20, wherein the predetermined .[∇]. flight path angle 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 voter, and the differential signal provided by said subtractor is transmitted to a logic unit which acts on the flight control horizon attitude director indicator in order to cause the pitching tendency pitch command bar to disappear when said differential signal remains zero for a period of time greater, than a predetermined period.