CA1088655A - Flying method and system using total power for an aircraft - Google Patents
Flying method and system using total power for an aircraftInfo
- Publication number
- CA1088655A CA1088655A CA256,222A CA256222A CA1088655A CA 1088655 A CA1088655 A CA 1088655A CA 256222 A CA256222 A CA 256222A CA 1088655 A CA1088655 A CA 1088655A
- Authority
- CA
- Canada
- Prior art keywords
- signal
- gamma
- flight path
- speed
- path angle
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
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Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/04—Control of altitude or depth
- G05D1/06—Rate of change of altitude or depth
- G05D1/0607—Rate of change of altitude or depth specially adapted for aircraft
- G05D1/0653—Rate of change of altitude or depth specially adapted for aircraft during a phase of take-off or landing
- G05D1/0661—Rate of change of altitude or depth specially adapted for aircraft during a phase of take-off or landing specially adapted for take-off
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/0055—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot with safety arrangements
- G05D1/0077—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot with safety arrangements using redundant signals or controls
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- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
ABSTRACT OF THE DISCLOSURE
A flying method using total energy, in particular for the take-off and overshoot of an aircraft, is disclosed in which the aerodynamic flight path angle .gamma. a is governed by reference to a desired flight path angle .gamma. d which is the potential flight path .gamma. t modulated by the difference between the aircraft speed V and a reference speed V2. An error signal .delta. representative of the difference between the aerodynamic flight path angle .gamma. a and the desired flight path angle .gamma. d is displayed. The display of the desired flight path angle .gamma. d may be by means of the pitch com-mand bar of an attitude director indicator for example.
A flying method using total energy, in particular for the take-off and overshoot of an aircraft, is disclosed in which the aerodynamic flight path angle .gamma. a is governed by reference to a desired flight path angle .gamma. d which is the potential flight path .gamma. t modulated by the difference between the aircraft speed V and a reference speed V2. An error signal .delta. representative of the difference between the aerodynamic flight path angle .gamma. a and the desired flight path angle .gamma. d is displayed. The display of the desired flight path angle .gamma. d may be by means of the pitch com-mand bar of an attitude director indicator for example.
Description
i~ `
. -The present invention relates to a flying rnethod and sys-tem, in particular for take-off or overshoot of an aircraft, using the total energy.
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 an aircraft symbol with respect to a moving sphere and the infor-mation of the attitude director indicator which is visualised by command bars (the information of the attitude director indicator 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 air-::: : . .
craft, (mass, chosen position of the flaps), to the machanics i (occasional conditions such as de-icing etc.. ), to the meteoro-logical 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 Vl, a speed V2 and a speed of rotation Vr which will be referred to hereafter.
- 30 When in possession of these data, the pilot may proceed with the aircraft take-off. To this end, once in the centre of the runway and after having reached an adequate engine condition . '; .~.; :
,~ - 1 -.
` 1(~88655 ` :
~ which generally, in the case of jet engines, corresponds to a ;; maximum thrust or close to the latter, the aircraft 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 eleva-tors and the aircraft takes-off several instants after at a speed Vloof. 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 pitch attitude l = 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 procedure is then as follows:
In the case where, at the time of the breakdown, the air-i .
craft has not yet reached the speed Vl, 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 ; 20 has exceeded the speed Vl, 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 attitude director indicator.
In this case, the pilot's work consists of bringing the pitch command bar of attitude of the director indicator to zero and of keeping it at zero. The movement of the longitudinal bar above or be]ow the aircraft symbol thus indicating an order to ~ pitch up or an order to pitch down.
,: ' r ~ - 2 -The invention thus relates to a method and system for flying an aircraft by means of an attitude director indicator, in particular as regards take-off and overshoot.
According to the invention, this method consists essen-tially of governing the aerodynamic flight path angle y a of the . aircraft by a desired flight path angle y d which is in turn the potential flight path y t modulated by the variation of the aerodynamic speed V with respect to a reference value V2 and of : displaying the error signal ~ between the aerodynamic flight path 10 angle y a and the desired flight path angle y d by means of the pitch command bar of an attitude director indicator.
The method thus uses the potential flight path y t which represents the flight path angle which the aircraft should assume , . . .
: to maintain its constant speed in rectilinear flight and which is expressed as follows:
yt = g (jx cosa - jz sin~) in which: ~ is the aerodynamic angle of attack of the aircraft, ~:i jx is the acceleration on the axis of bank, jz is the accelera-.. tion on the axis of yaw and g is the acceleration of gravity.
Generally, the measurement of y t may be undertaken in a computer, from the real aerodynamic angle of attack 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 energy variometer may be used to advantage, which provides in particular:
Information relating to the potential flight path y t, and Information relating to the aerodynamic flight path angle y a, (which is equal to the algebraic difference between the 30 pitch attitude ~1 and the aerodynamic angle of attack ~. -Furthermore, it is known that aircraft are equipped more and more frequently for inertial navigation and consequently com-~ 3 -:
.- ~ . .
:
`` 1(~88655 ~ prise 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 potential flight path y t.
The invention therefore relates to the realisation of a circuit making it possible to calculate the potential flight path .` y t 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.
.:: 10 To achieve this result, the invention uses mainly the fol-lowing formulae, assuming zero side-slip of the aircraft and that ; the wings are horizontal, y t = _ ddt + sin y a ` y t ~ gl ddt + r a .~ According to these relationships, the value of y t 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 : i~
~ of the aerodynamic flight path angle y a which, as above-mentioned -1 20 may be obtained by forming in conventional manner, the algebraic ;
difference between the pitch attitude ~1 and the aerodynamic : angle of attack ~.
It should be noted in this respect that in the simplest : case and in the hypothesis where the calculation of the value of y t is not indispensable and does not serve to ensure the detec-tion of an engine failure, the calculation of y a is not neces-: ' sary. In fact, if we substitute the value g dt + Y a for the value y t, the control yd - ya = 0 may be written successively in an equivalent manner:
~`: 30 y t - k(V-V2) - y a = 0 g dt + Y a - k(V-V2) - ya = 0 g dt ~ k(V-V2) = 0 ~ - 4 -.; ~
, 1~88655 which consists of controlling the derivative dVt 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 0 = 18, the error signal ~ is transmitted to a voter which compares it with a difference signal ~1 proportional to (l - 18) and to a reference signal ~2~ corresponding~for example to a constant order to pitch down by 20 and which effects the control of the pitch command bar of the attitude director indi-cator by selecting from the signals ~ 2~ 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 energy, in particular when they are trans-lated by values of the signal ~ causing erroneous orders to pitch down, which lead the pilot to impart a dangerous negative aero-, . ~
dynamic flight path angle to the aircraft.
Incontrast toan activebreakdown whichbecomes apparentand as soonas itappears disturbsthe membercontrolled possiblythe pitch commandbar ofthe attitudedirector indicator adormantbreakdownis intendedto mean a breakdown affecting a circuit which operates andacts on the controlled member only under certain particular conditionswhich 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 ~. .~ ,.
. .
1~88~S5 serious consequences from the point of view of safety.
The ob~ect 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 ~, to prevent the pitch command bar from leading to information less than a minimum safety flight path angle 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 correc-tions.
To achieve these results, the invention proposes a device .~,..
providing protection against abnormal negative flight path angles - of the aircraft due in particular to poor operation of the flying .: .
system providing the signal ~ and which uses the properties of said voter.
According to this device, the input ~2 of the voter is connected to a circuit providing a signal K (ya - yo) in which K
is an adjustment gain, yo is a value representing the minimum safety flight path angle, 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 voter, the presence of a signal ~2 when the latter is selected and on the other hand, to automati-cally disconnect the entire flying system when the signal ~2 remains selected at the end of a predetermined period of time.
A circuit of this type may comprise a subtractor arranged between the input ~2 and the output of the voter and a logic circuit comprising a delay which disconnects the flying system when the differential signal provided by said subtractor is can-celled and remains zero for a period of time greater than a pre-determined period of time ~.
Embodiments of the invention will be described hereafter, as non-limiting examples, with reference to the accompanying , :
` 1~8~6S~
drawings in which:
Fig. 1 is a circuit diagram making it possible to illus-trate the principle of the flying system according to the inven-~ .
tion, in the case where signals representative of the values y a ~ and y t are available on board the aircraft;
- Fig. 2 is a circuit diagram of a preferred embodiment of the system illustrated in Fig. l;
. 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-engined - aircraft with failure of one of the engines at the speed Vl, 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 illus-trate the principle of the flying system according to the inven-tion, 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 sys-tem using total energy, 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 j of the flying system occurs; the curve representative of the sig-: i .
nal ~ 1 ~ 18) 1.5; the curve representative of the signal ~; the curve representative of the signal ~2 = K (ya - yo) and the curve representative of a signal ~'2 corresponding to an order to pitch down by analogy with the signal ~2 mentioned in the main patent.
With reference to Fig. 1, the flying system, in particular '~ .
~ - 7 -.-' .- ' for take-off and overshoot of the aircraft (block 1), acts essentially on the pitching bar 2 (pitch command bar) of a flight control horizon 3 of conventional type. I-t is known that the position of this command bar 2, with respect to the air-craft symbol 4, makes it possible to give the pilot either an order to piteh down or an order to fly tail-down or even, when ,::-.
the pitch command bar 2 is superimposed on the aircraft symbol ,~ 4, that the aircraft 1 is in the desired configuration.
Depending on the indieations of the 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 aireraft 1, and in particular, of the speed V, of the potential flight path ~ t and of the aerodynamic flight path angle ~ a which, as above mentioned, are the three essential parameters used in the flying system using total energy according to the inventlon .
The potential flight path signal ~ t which may be obtained - by a computer, from the actual aerodynamie angle of attaek and two aeeelerometers whose pereeptible axes are respeetively parallel to the axis of bank (direetion Jx) and the axis of yaw (deteetion Jz) is transmitted to an adder 6 whieh receives, at its seeond input 7, a signal eorresponding to the variation of the speed V with respeet to a referenee value, for example V2 +
10 knots or even only V2. In the example illustrated, the speed signal V eoming from an anemometer is transmitted to a subtraetor 8 whieh reeeives, at its seeond input 9, a signal eorresponding to the referenee speed. The signal resulting from this differenee is amplified (unit 10) and is then transmitted to an amplitude limiter 11 whieh effeets a limitation of the minimum and maximum values of the differenee. The amplitude limiter 11 is eonneeted to the adder 6 whose output is eonneeted to a subtraetor 12 whieh ''' ~4 ~
' ' ' .
`` 1~886S~>
receives a signal corresponding to the aerodynamic flight path angle r a at its second input 13.
A signal provided at the output of the subtractor 12 is transmitted to a voter 14 (whose function will be explained here-; after), which controls the pitch command bar 2 of the attitude director indicator 3.
In this respect, it will be noted that one of the consider-able advantages of this system consists in that it combines accelerometric information (calculation of r t) and angular infor-mation (calculation of r a) with anemometric information (calcu-lation V) which react conversely at the time of 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 aneomometric information and consequently, the flying system is not subject to any considerable disturbances.
If the operation of the voter 14 is not taken into consid-eration and the latter is replaced by a simple electrical connec-tion to the member for controlling the pitch command bar 2 of the .
attitude director indicator 3, the indications of the bar 2 may . i-- 20 be interpreted as follows: in the case where the signal ~ pro-vided by the subtractor 12 is positive, the pitch command bar 2 is located below the aircraft symbol 4, which corresponds to an order to pitch down.
The state ~ > o signifies that y a is greater than y d and, consequently, inter alia, may result from two situations of the following types:
a) In the case where V = V reference and where y a > y t, 'r,: the pitch command bar 2 of the flight control device thus indi-` cates an order to pitch down to restore y a to the value of y t in order to prevent deceleration;
b) In the case where V < V reference and y a = y t, the pitch command bar 2 of the attitude director indicator then :' .
. .
1~88655 :;.
j......... indicates an order to pitch down to momentarily reduce y a in -- order to restore V to V ref.
.. In the case where the signal ~ provided by the subtractor 12 is negative, the pitch command bar 2 is located above the air-craft symbol 4, which corresponds to an order to pitch up.
The state ~ < o signifies that y a is smaller than y d .~ and consequently, inter alia, may result from two situations of the following types:
- c) In the case where V = V reference and where y a < y t, . .
the pitch command bar 2 thus indicates an order to pitch up in . order to restore y a to the value of y t in order to prevent acceleration;
d) In the case where V > V reference and y a = y t, the pitch command bar thus indicates an order to pitch up in order to momentarily increase y a 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 voter 14.
This 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 ~1 proportional to the difference ~1 ~
: 18 (91 being the pitch attitude, whereas the third 17 receives a reference signal ~2 corresponding in the embodiment described to an order to dive by 20.
The function of the voter 14 is to compare the signals applied thereto by its three inputs and to transmit to the member for controlling the pitch command bar 2 of the 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 pitch attitude signal) is selected, if ~1 > 18, the pitch command bar 2 gives an order to pitch down, on the other hand if ~1 < 18, the pitch command bar 2 gives an order to pitch up.
- ' , .:
1~88655 . .
With reference to Fig. 2, which is a circuit diagram of one embodiment of the invention, the signal ~ ~ is transmitted by the intermediary of a low pass filter 21, to a subtractor 22 whose second input 23 receives the signal y a which, as above-mentioned, is equal to (~ ) Consequently, this signal y a is obtained by means of a subtractor 24 which receives, on the one hand, a signal representative of ~1 and, on the other hand, a signal representative of the aerodynamic angle of attack ~
which is filtered by means of a low pass filter 25. In order ~; 10 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 ~ during rotation, the circuit ; providing the signal representative of ~ comprises a commutation device which makes it possible: on the one hand, to transmit : . .
~ to the subtractor 24 during the take off roll with the main under-;:, carriage compressed (for example by means of a relay 26 controlled by a detector associated with the under-carriage), a signal repre-sentative of the pitch attitude ~1 in place of the aerodynamic angle of attack ~ 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 dependlng 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 ,., . . .
: ~ :
~886~5 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 y t 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 filter-ing member 40 followed by a diode 41 connected to the adder 38.
- This derivation is particularly provided in order to increase the ~ 10 sensitivity of the level detector 36 when the derivative of y t 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, when the aircraft is not in clean configura-tion (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 y t drops below the value 6~, the level detector 36 acts on the double commutator to interrupt the 10 knots signal.
When the 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 mernber 42, which serves as an 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 acquiring the reference speed. The selector 53 makes it possible to modify the value of the limitations of .. ~
:
",, 88655~
42 depending on the state of the level detector 36, on the posi-tion of the leading edges (and possibly on the condition "air-craft on the ground") In particular, when the aircraft is not in clean configura-tion (case of take-off), the level detector36 imposes wide limits when the output of the adder 38 exceeds 6 or narrow limits when ~ t drops below 6. In this latter case, the limitation of nega-tive sign is adjusted in order to constitute an implicit protec-.
, tion of the system against trajectories having a negative flight path angle subsequent to an engine failure on take-off. In clean ~ aircraft configuration (use during cruising) the level detector : 36 is inoperative and wide limitations are imposed.
i~ 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 voter 45, also by means of a matching amplifier 46.
, . .
At its second input 47, the voter 45 receives a signal proportional to (~1 ~ 18) obtained by means of a subtractor 48 receiving the signal ~1 and the signal corresponding to 18.
20 This subtractor ~8 is connected to the voter 45 by means of a matching amplifier 49.
, The third input 50 of the voter 45 receives a constant ` reference signal corresponding for example to an order of 20.
The output of the voter 45 is connected to the member for controlling the pitch command bar of the attitude director indi-cator by means of a matching amplifier 51 and a limiter 52 which makes it possible toauthorise maximum displacement of the 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 .' '~ .
.
1(~88ti5~
:.
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: palm switch 6f of the throttle levers, etc.) The attitude director indicator thus has the configuration illustrated in Fig. 3, in which the pitch command bar gives an order to pitch down, since only the limited term (V - V2) is ' different from zero.
:; 10 During the take-off roll, as soon as the brakes are , :
released and the take-off thrust established, the potential flight path y t increases (approximately 12) and, consequently the level detector 36 causes the selection of V2 + 10 knots.
The command bar which is located approximately in the upper posi-~,.
tion thus indicates an order to pitch up and is maintained inthis position approximately until rotation (Fig. 4).
During rotation, owing to the fact that the potential :.
flight path decreases and the aerodynamic flight path angle increases, the pitch command bar begins to drop and this move-ment is accentuated upon take-off (Fig. 5).
- The pllot must then act on the flying controls in order to maintain the pitch command bar in coincidence with the air-craft symbol (Fig. 6).
Maintaining the position of the tendency bar firstly ,~ ensures the acquisition of an acceleration to V2 + 10 knots then; ~;~
either the holding of V2 + 10 knots (al being less than 18); or the holding of (~1 at 18) (with acceleration).
Once this flying configuration has been achieved, the pilot may proceed to accelerate from V2 + 10 (or from ~1 = 18) with the hyperlift devices retracted.
For this, as soon as the aircraft reaches a predetermined altitude, since it is stablised at V2 + 10 knots (or at l = 18), ~ . ~
v~ - 14 -. ..
1~88 ',' 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 y a to 4 below y t. 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 pltch command bar gives an order to pitch down, i.e. an order to vary the pitch attitude in order to bring y a to 4 below y t (Fig. 7).
After the retraction of the flaps, the pitch command bar gives an order to 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 inven-. tion 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 Vl, or during the stage ,~ , succeeding take-off and up to the stabilisation of the aircraft at a speed of V2 + 10 knots (or ~1 = 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 potential flight path y t is less than the flight path angle y a and gives an order (a) to pitch down, whereas the speed variation signal which passes from v - (V2 + 10) to (V - V2) gives an order (b) to pitch up. In this case, the order (a) is preponderant and consequently the pitch command bar gives an order to pitch down.
The pilot consequently acts to bring the pitch command bar into coincidence with the aircraft symbol in order to achieve and maintain the configuration y t = y a and V = V2.
In the case where the engine failure occurs between the ~t - 15 -~ 1~88ti55 . speed Vl and take-off, at the time of the failure, the drop of ~ t 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 rota-:~, ; tion. At the time of rotation and take-off, the command bar con-tinues to drop.
The pilot must act on his flying controls to bring and : maintain the command bar in coincidence with the 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 refer-ence 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 palm switches, implanted on the throttle .~ 20 lever and provided for this operation. This action returns to ..
the logic of the syfitem, 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 pitch ~ -command bar is virtually zero (during a rotation at average speed) (Fig. 12).
After rotationof the pitch attitude, by means of the sys-tem according to the invention, the pilot is able to acquire and maintain a speed of V ref. + 10 knots (or ~1 = 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 , ': " ' ' ;88~55 .: :
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 over-shoot, 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 ~ t < 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 pi~lot system for example.
.
" The indications of the pitch command bar make it possible to achieve and maintain the reference speed displayed. Cruising is distinguished by the system by the condition of 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 air-craft acts in a similar manner to that previously described, on the pitching bar 62 of the attitude director indicator 63.
As previously mentioned, the aircraft shown diagrammati-cally with its control members, respectively by the units 64 and 65, comprises a central inertial unit providing a signal repre-sentative of the speed Vs. After having been derived with respect ' :
~886~5 :~, to time and multiplied by a coefficient equal to -, 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.
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 voter 72 which controls the pitch command bar 62 of the attitude director indicator 63.
!- The 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 ~1 proportional to the dif-ference l ~ 18 (~1 being the pitch attitude), whereas the third input 77 receives a reference signal ~2 corresponding in the embodiment described to an order of 20.
The function of the 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 pitch command bar 62 of the attitude director indicator 63, the signal whose value, at a given time, is comprised between the value of the two other sig-nals, 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 lS identical to that shown in Fig. 2, apart from the fact that instead of using a signal ~ t, provided by an accelero-, . .' ' :
:;
1~886S5 metric box provided for this purpose, this signal y t is calcu-lated from the speed signal Vs supplied by a central inertial unit 80.
,., This signal Vs coming from the central unit 80 is trans-.: .
~: mitted to a derivation circuit 81 also acting as a high pass ~ilter for the transfer function 1 + s (s being the Laplace oper-ator) and whose output is connected to an amplifier 82 for the `~ gain -. ;
~ This amplifier 82 is in turn connected to an adder 83 i`, 10 which at its second input also receives a signal representative of the flight path angle y a. This signal y a is obtained by forming the difference ~ - a in a conventional manner, by means of a subtractor 84, after having filtered the signals ~ and ~ by , means of low pass filters 85 and 85l having a transfer function ''~". 1 i ~ 1 5 -The signal ~ t 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.
` 20 In the system for flying the aircraft illustrated, with its controls, by the blocks 86 and 86' (Fig. 17), the signal of potential flight path y t which may be obtained in conventional manner from the actual 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 varia-tion 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 . -- 19 . ., .
~ 1~88655 ''`
minimum and maximum values of the variation. The amplitude limiter 92 is connected to the adder 88 whose output is con-nected to a subtractor 93, which receives a signal corresponding to the aerodynamic flight path angle y a at its second input 94.
The signal ~, provided at the output of the subtractor 93 is transmitted to a voter 95, which controls the pitch command bar of the attitude director indicator 87.
The voter 95 comprises three inputs, whereof one 96 is .` connected to the subtractor 93, the second 97 is connected to a 10 circuit providing a signal ~1 proportional to the difference 91 ~ 18 (~1 being the longitudinal pitch attitude), whereas the third input 98 receives a signal K (ya - yo) emanating from an ~; amplifier 99 for an adjustment gain K, which receives the differ-; ential signal of a subtractor 100. On the one hand, this sub-tractor 100 receives a signal representative of the aerodynamic : flight path angle ya taken for example from 94 and on the other - hand, a signal yo representative of an angle of predetermined value, for example 0.5.
The signal provided by the amplifier 99 is also trans-mitted to a subtractor 102 which also receives the output signalof the voter 95.
The differential signal provided by the subtractor 102 is transmitted to a logic unit 103 which acts on the attitude direc-tor indicator 87, in order to cause the 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 des-cribed 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 con-trolled by a circuit for detecting an engine failure, which enables the reference speed to assume two values, namely the . . ~
' ' :.
:
l~ss6ss ~- 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 possibl~ to illustrate the action of the afore-described protection device, ,. .
.j~ 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 r,` aircraft, a curve which comprises: a first part 105 correspond-~j 10 ing to the phase during which the pilot seeks to reach the speed ,;:
V2 + 10 knots or the pitch attitude ~1 = 18; a second part 106 ~'?;'`'~ corresponding to holding a speed of V2 + 10 knots; a third part 107 in which are located in broke line the trajectory 108 cor-~ responding to normal operation, in full line the trajectory 104 : i corresponding to the system according to the main patent dis-, turbed by poor operation, in broken line the trajectory 109 cor-' rected by the protection action according to the invention in the case of poor operation.
This diagram also shows, in correlation with the trajec-~;i 20 tories 104, 108 and 109, the curves 111, 112 and 113 representing ; , the signals ~ and ~2~ at the input of the voter and a hori-zontal line 114 corresponding to a constant signal ~'2 indicating a considerable and eonstant order to pitch down.
,~:
` It will thus be seen that after reaching the speed V2 +
10 knots or ~1 = 18, during the phase of holding this speed, j the signals ~ and ~2 are maintained at a substantially con-stant level and in the order ~2>~>~1 and, consequently, it is ` the signal ~ which is selected (curve shown in broken line 115).
: .
; As soon as a failure occurs in the circuit producing the signal ' 30 ~ and which is translated by an order to pitch down and by a ; deerease in the flight path angle of the aircraft, the value of the signal ~2 = (ro - ~a) K decreases and, at the end of a cer-. .
,' ~
:, 1~88655 :
tain period of time, becomes less than the value of the signal ~.
In other words, there is an intersection of the curves ~ and ~2.
Beyond the point where the value of ~2 becomes less than the value of ~, the value of ~2 is located between that of ~ and ~l and due to this, it is the signal ~2 which is selected (see curve shown in broken line 115).
Consequently, the flying system is governed by K(ya - yo), which makes it possible to maintain an aerodynamic flight path angle ya equal to yo. In practice, the value of yo which consti-tutes the bottom aerodynamic flight path angle, below which one should not fall, is approximately 0.5.
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 pitch command bar of the attitude director indicator 87 to disappear, when the signal ~2 is selected at the end of a predetermined time, for example five seconds.
;~! 20 It is clear that this disappearance of the 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.
`:. :
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"
: . ' :
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. -The present invention relates to a flying rnethod and sys-tem, in particular for take-off or overshoot of an aircraft, using the total energy.
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 an aircraft symbol with respect to a moving sphere and the infor-mation of the attitude director indicator which is visualised by command bars (the information of the attitude director indicator 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 air-::: : . .
craft, (mass, chosen position of the flaps), to the machanics i (occasional conditions such as de-icing etc.. ), to the meteoro-logical 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 Vl, a speed V2 and a speed of rotation Vr which will be referred to hereafter.
- 30 When in possession of these data, the pilot may proceed with the aircraft take-off. To this end, once in the centre of the runway and after having reached an adequate engine condition . '; .~.; :
,~ - 1 -.
` 1(~88655 ` :
~ which generally, in the case of jet engines, corresponds to a ;; maximum thrust or close to the latter, the aircraft 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 eleva-tors and the aircraft takes-off several instants after at a speed Vloof. 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 pitch attitude l = 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 procedure is then as follows:
In the case where, at the time of the breakdown, the air-i .
craft has not yet reached the speed Vl, 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 ; 20 has exceeded the speed Vl, 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 attitude director indicator.
In this case, the pilot's work consists of bringing the pitch command bar of attitude of the director indicator to zero and of keeping it at zero. The movement of the longitudinal bar above or be]ow the aircraft symbol thus indicating an order to ~ pitch up or an order to pitch down.
,: ' r ~ - 2 -The invention thus relates to a method and system for flying an aircraft by means of an attitude director indicator, in particular as regards take-off and overshoot.
According to the invention, this method consists essen-tially of governing the aerodynamic flight path angle y a of the . aircraft by a desired flight path angle y d which is in turn the potential flight path y t modulated by the variation of the aerodynamic speed V with respect to a reference value V2 and of : displaying the error signal ~ between the aerodynamic flight path 10 angle y a and the desired flight path angle y d by means of the pitch command bar of an attitude director indicator.
The method thus uses the potential flight path y t which represents the flight path angle which the aircraft should assume , . . .
: to maintain its constant speed in rectilinear flight and which is expressed as follows:
yt = g (jx cosa - jz sin~) in which: ~ is the aerodynamic angle of attack of the aircraft, ~:i jx is the acceleration on the axis of bank, jz is the accelera-.. tion on the axis of yaw and g is the acceleration of gravity.
Generally, the measurement of y t may be undertaken in a computer, from the real aerodynamic angle of attack 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 energy variometer may be used to advantage, which provides in particular:
Information relating to the potential flight path y t, and Information relating to the aerodynamic flight path angle y a, (which is equal to the algebraic difference between the 30 pitch attitude ~1 and the aerodynamic angle of attack ~. -Furthermore, it is known that aircraft are equipped more and more frequently for inertial navigation and consequently com-~ 3 -:
.- ~ . .
:
`` 1(~88655 ~ prise 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 potential flight path y t.
The invention therefore relates to the realisation of a circuit making it possible to calculate the potential flight path .` y t 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.
.:: 10 To achieve this result, the invention uses mainly the fol-lowing formulae, assuming zero side-slip of the aircraft and that ; the wings are horizontal, y t = _ ddt + sin y a ` y t ~ gl ddt + r a .~ According to these relationships, the value of y t 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 : i~
~ of the aerodynamic flight path angle y a which, as above-mentioned -1 20 may be obtained by forming in conventional manner, the algebraic ;
difference between the pitch attitude ~1 and the aerodynamic : angle of attack ~.
It should be noted in this respect that in the simplest : case and in the hypothesis where the calculation of the value of y t is not indispensable and does not serve to ensure the detec-tion of an engine failure, the calculation of y a is not neces-: ' sary. In fact, if we substitute the value g dt + Y a for the value y t, the control yd - ya = 0 may be written successively in an equivalent manner:
~`: 30 y t - k(V-V2) - y a = 0 g dt + Y a - k(V-V2) - ya = 0 g dt ~ k(V-V2) = 0 ~ - 4 -.; ~
, 1~88655 which consists of controlling the derivative dVt 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 0 = 18, the error signal ~ is transmitted to a voter which compares it with a difference signal ~1 proportional to (l - 18) and to a reference signal ~2~ corresponding~for example to a constant order to pitch down by 20 and which effects the control of the pitch command bar of the attitude director indi-cator by selecting from the signals ~ 2~ 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 energy, in particular when they are trans-lated by values of the signal ~ causing erroneous orders to pitch down, which lead the pilot to impart a dangerous negative aero-, . ~
dynamic flight path angle to the aircraft.
Incontrast toan activebreakdown whichbecomes apparentand as soonas itappears disturbsthe membercontrolled possiblythe pitch commandbar ofthe attitudedirector indicator adormantbreakdownis intendedto mean a breakdown affecting a circuit which operates andacts on the controlled member only under certain particular conditionswhich 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 ~. .~ ,.
. .
1~88~S5 serious consequences from the point of view of safety.
The ob~ect 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 ~, to prevent the pitch command bar from leading to information less than a minimum safety flight path angle 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 correc-tions.
To achieve these results, the invention proposes a device .~,..
providing protection against abnormal negative flight path angles - of the aircraft due in particular to poor operation of the flying .: .
system providing the signal ~ and which uses the properties of said voter.
According to this device, the input ~2 of the voter is connected to a circuit providing a signal K (ya - yo) in which K
is an adjustment gain, yo is a value representing the minimum safety flight path angle, 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 voter, the presence of a signal ~2 when the latter is selected and on the other hand, to automati-cally disconnect the entire flying system when the signal ~2 remains selected at the end of a predetermined period of time.
A circuit of this type may comprise a subtractor arranged between the input ~2 and the output of the voter and a logic circuit comprising a delay which disconnects the flying system when the differential signal provided by said subtractor is can-celled and remains zero for a period of time greater than a pre-determined period of time ~.
Embodiments of the invention will be described hereafter, as non-limiting examples, with reference to the accompanying , :
` 1~8~6S~
drawings in which:
Fig. 1 is a circuit diagram making it possible to illus-trate the principle of the flying system according to the inven-~ .
tion, in the case where signals representative of the values y a ~ and y t are available on board the aircraft;
- Fig. 2 is a circuit diagram of a preferred embodiment of the system illustrated in Fig. l;
. 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-engined - aircraft with failure of one of the engines at the speed Vl, 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 illus-trate the principle of the flying system according to the inven-tion, 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 sys-tem using total energy, 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 j of the flying system occurs; the curve representative of the sig-: i .
nal ~ 1 ~ 18) 1.5; the curve representative of the signal ~; the curve representative of the signal ~2 = K (ya - yo) and the curve representative of a signal ~'2 corresponding to an order to pitch down by analogy with the signal ~2 mentioned in the main patent.
With reference to Fig. 1, the flying system, in particular '~ .
~ - 7 -.-' .- ' for take-off and overshoot of the aircraft (block 1), acts essentially on the pitching bar 2 (pitch command bar) of a flight control horizon 3 of conventional type. I-t is known that the position of this command bar 2, with respect to the air-craft symbol 4, makes it possible to give the pilot either an order to piteh down or an order to fly tail-down or even, when ,::-.
the pitch command bar 2 is superimposed on the aircraft symbol ,~ 4, that the aircraft 1 is in the desired configuration.
Depending on the indieations of the 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 aireraft 1, and in particular, of the speed V, of the potential flight path ~ t and of the aerodynamic flight path angle ~ a which, as above mentioned, are the three essential parameters used in the flying system using total energy according to the inventlon .
The potential flight path signal ~ t which may be obtained - by a computer, from the actual aerodynamie angle of attaek and two aeeelerometers whose pereeptible axes are respeetively parallel to the axis of bank (direetion Jx) and the axis of yaw (deteetion Jz) is transmitted to an adder 6 whieh receives, at its seeond input 7, a signal eorresponding to the variation of the speed V with respeet to a referenee value, for example V2 +
10 knots or even only V2. In the example illustrated, the speed signal V eoming from an anemometer is transmitted to a subtraetor 8 whieh reeeives, at its seeond input 9, a signal eorresponding to the referenee speed. The signal resulting from this differenee is amplified (unit 10) and is then transmitted to an amplitude limiter 11 whieh effeets a limitation of the minimum and maximum values of the differenee. The amplitude limiter 11 is eonneeted to the adder 6 whose output is eonneeted to a subtraetor 12 whieh ''' ~4 ~
' ' ' .
`` 1~886S~>
receives a signal corresponding to the aerodynamic flight path angle r a at its second input 13.
A signal provided at the output of the subtractor 12 is transmitted to a voter 14 (whose function will be explained here-; after), which controls the pitch command bar 2 of the attitude director indicator 3.
In this respect, it will be noted that one of the consider-able advantages of this system consists in that it combines accelerometric information (calculation of r t) and angular infor-mation (calculation of r a) with anemometric information (calcu-lation V) which react conversely at the time of 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 aneomometric information and consequently, the flying system is not subject to any considerable disturbances.
If the operation of the voter 14 is not taken into consid-eration and the latter is replaced by a simple electrical connec-tion to the member for controlling the pitch command bar 2 of the .
attitude director indicator 3, the indications of the bar 2 may . i-- 20 be interpreted as follows: in the case where the signal ~ pro-vided by the subtractor 12 is positive, the pitch command bar 2 is located below the aircraft symbol 4, which corresponds to an order to pitch down.
The state ~ > o signifies that y a is greater than y d and, consequently, inter alia, may result from two situations of the following types:
a) In the case where V = V reference and where y a > y t, 'r,: the pitch command bar 2 of the flight control device thus indi-` cates an order to pitch down to restore y a to the value of y t in order to prevent deceleration;
b) In the case where V < V reference and y a = y t, the pitch command bar 2 of the attitude director indicator then :' .
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1~88655 :;.
j......... indicates an order to pitch down to momentarily reduce y a in -- order to restore V to V ref.
.. In the case where the signal ~ provided by the subtractor 12 is negative, the pitch command bar 2 is located above the air-craft symbol 4, which corresponds to an order to pitch up.
The state ~ < o signifies that y a is smaller than y d .~ and consequently, inter alia, may result from two situations of the following types:
- c) In the case where V = V reference and where y a < y t, . .
the pitch command bar 2 thus indicates an order to pitch up in . order to restore y a to the value of y t in order to prevent acceleration;
d) In the case where V > V reference and y a = y t, the pitch command bar thus indicates an order to pitch up in order to momentarily increase y a 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 voter 14.
This 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 ~1 proportional to the difference ~1 ~
: 18 (91 being the pitch attitude, whereas the third 17 receives a reference signal ~2 corresponding in the embodiment described to an order to dive by 20.
The function of the voter 14 is to compare the signals applied thereto by its three inputs and to transmit to the member for controlling the pitch command bar 2 of the 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 pitch attitude signal) is selected, if ~1 > 18, the pitch command bar 2 gives an order to pitch down, on the other hand if ~1 < 18, the pitch command bar 2 gives an order to pitch up.
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1~88655 . .
With reference to Fig. 2, which is a circuit diagram of one embodiment of the invention, the signal ~ ~ is transmitted by the intermediary of a low pass filter 21, to a subtractor 22 whose second input 23 receives the signal y a which, as above-mentioned, is equal to (~ ) Consequently, this signal y a is obtained by means of a subtractor 24 which receives, on the one hand, a signal representative of ~1 and, on the other hand, a signal representative of the aerodynamic angle of attack ~
which is filtered by means of a low pass filter 25. In order ~; 10 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 ~ during rotation, the circuit ; providing the signal representative of ~ comprises a commutation device which makes it possible: on the one hand, to transmit : . .
~ to the subtractor 24 during the take off roll with the main under-;:, carriage compressed (for example by means of a relay 26 controlled by a detector associated with the under-carriage), a signal repre-sentative of the pitch attitude ~1 in place of the aerodynamic angle of attack ~ 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 dependlng 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 ,., . . .
: ~ :
~886~5 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 y t 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 filter-ing member 40 followed by a diode 41 connected to the adder 38.
- This derivation is particularly provided in order to increase the ~ 10 sensitivity of the level detector 36 when the derivative of y t 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, when the aircraft is not in clean configura-tion (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 y t drops below the value 6~, the level detector 36 acts on the double commutator to interrupt the 10 knots signal.
When the 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 mernber 42, which serves as an 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 acquiring the reference speed. The selector 53 makes it possible to modify the value of the limitations of .. ~
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",, 88655~
42 depending on the state of the level detector 36, on the posi-tion of the leading edges (and possibly on the condition "air-craft on the ground") In particular, when the aircraft is not in clean configura-tion (case of take-off), the level detector36 imposes wide limits when the output of the adder 38 exceeds 6 or narrow limits when ~ t drops below 6. In this latter case, the limitation of nega-tive sign is adjusted in order to constitute an implicit protec-.
, tion of the system against trajectories having a negative flight path angle subsequent to an engine failure on take-off. In clean ~ aircraft configuration (use during cruising) the level detector : 36 is inoperative and wide limitations are imposed.
i~ 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 voter 45, also by means of a matching amplifier 46.
, . .
At its second input 47, the voter 45 receives a signal proportional to (~1 ~ 18) obtained by means of a subtractor 48 receiving the signal ~1 and the signal corresponding to 18.
20 This subtractor ~8 is connected to the voter 45 by means of a matching amplifier 49.
, The third input 50 of the voter 45 receives a constant ` reference signal corresponding for example to an order of 20.
The output of the voter 45 is connected to the member for controlling the pitch command bar of the attitude director indi-cator by means of a matching amplifier 51 and a limiter 52 which makes it possible toauthorise maximum displacement of the 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 .' '~ .
.
1(~88ti5~
:.
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: palm switch 6f of the throttle levers, etc.) The attitude director indicator thus has the configuration illustrated in Fig. 3, in which the pitch command bar gives an order to pitch down, since only the limited term (V - V2) is ' different from zero.
:; 10 During the take-off roll, as soon as the brakes are , :
released and the take-off thrust established, the potential flight path y t increases (approximately 12) and, consequently the level detector 36 causes the selection of V2 + 10 knots.
The command bar which is located approximately in the upper posi-~,.
tion thus indicates an order to pitch up and is maintained inthis position approximately until rotation (Fig. 4).
During rotation, owing to the fact that the potential :.
flight path decreases and the aerodynamic flight path angle increases, the pitch command bar begins to drop and this move-ment is accentuated upon take-off (Fig. 5).
- The pllot must then act on the flying controls in order to maintain the pitch command bar in coincidence with the air-craft symbol (Fig. 6).
Maintaining the position of the tendency bar firstly ,~ ensures the acquisition of an acceleration to V2 + 10 knots then; ~;~
either the holding of V2 + 10 knots (al being less than 18); or the holding of (~1 at 18) (with acceleration).
Once this flying configuration has been achieved, the pilot may proceed to accelerate from V2 + 10 (or from ~1 = 18) with the hyperlift devices retracted.
For this, as soon as the aircraft reaches a predetermined altitude, since it is stablised at V2 + 10 knots (or at l = 18), ~ . ~
v~ - 14 -. ..
1~88 ',' 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 y a to 4 below y t. 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 pltch command bar gives an order to pitch down, i.e. an order to vary the pitch attitude in order to bring y a to 4 below y t (Fig. 7).
After the retraction of the flaps, the pitch command bar gives an order to 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 inven-. tion 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 Vl, or during the stage ,~ , succeeding take-off and up to the stabilisation of the aircraft at a speed of V2 + 10 knots (or ~1 = 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 potential flight path y t is less than the flight path angle y a and gives an order (a) to pitch down, whereas the speed variation signal which passes from v - (V2 + 10) to (V - V2) gives an order (b) to pitch up. In this case, the order (a) is preponderant and consequently the pitch command bar gives an order to pitch down.
The pilot consequently acts to bring the pitch command bar into coincidence with the aircraft symbol in order to achieve and maintain the configuration y t = y a and V = V2.
In the case where the engine failure occurs between the ~t - 15 -~ 1~88ti55 . speed Vl and take-off, at the time of the failure, the drop of ~ t 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 rota-:~, ; tion. At the time of rotation and take-off, the command bar con-tinues to drop.
The pilot must act on his flying controls to bring and : maintain the command bar in coincidence with the 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 refer-ence 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 palm switches, implanted on the throttle .~ 20 lever and provided for this operation. This action returns to ..
the logic of the syfitem, 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 pitch ~ -command bar is virtually zero (during a rotation at average speed) (Fig. 12).
After rotationof the pitch attitude, by means of the sys-tem according to the invention, the pilot is able to acquire and maintain a speed of V ref. + 10 knots (or ~1 = 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 , ': " ' ' ;88~55 .: :
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 over-shoot, 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 ~ t < 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 pi~lot system for example.
.
" The indications of the pitch command bar make it possible to achieve and maintain the reference speed displayed. Cruising is distinguished by the system by the condition of 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 air-craft acts in a similar manner to that previously described, on the pitching bar 62 of the attitude director indicator 63.
As previously mentioned, the aircraft shown diagrammati-cally with its control members, respectively by the units 64 and 65, comprises a central inertial unit providing a signal repre-sentative of the speed Vs. After having been derived with respect ' :
~886~5 :~, to time and multiplied by a coefficient equal to -, 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.
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 voter 72 which controls the pitch command bar 62 of the attitude director indicator 63.
!- The 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 ~1 proportional to the dif-ference l ~ 18 (~1 being the pitch attitude), whereas the third input 77 receives a reference signal ~2 corresponding in the embodiment described to an order of 20.
The function of the 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 pitch command bar 62 of the attitude director indicator 63, the signal whose value, at a given time, is comprised between the value of the two other sig-nals, 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 lS identical to that shown in Fig. 2, apart from the fact that instead of using a signal ~ t, provided by an accelero-, . .' ' :
:;
1~886S5 metric box provided for this purpose, this signal y t is calcu-lated from the speed signal Vs supplied by a central inertial unit 80.
,., This signal Vs coming from the central unit 80 is trans-.: .
~: mitted to a derivation circuit 81 also acting as a high pass ~ilter for the transfer function 1 + s (s being the Laplace oper-ator) and whose output is connected to an amplifier 82 for the `~ gain -. ;
~ This amplifier 82 is in turn connected to an adder 83 i`, 10 which at its second input also receives a signal representative of the flight path angle y a. This signal y a is obtained by forming the difference ~ - a in a conventional manner, by means of a subtractor 84, after having filtered the signals ~ and ~ by , means of low pass filters 85 and 85l having a transfer function ''~". 1 i ~ 1 5 -The signal ~ t 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.
` 20 In the system for flying the aircraft illustrated, with its controls, by the blocks 86 and 86' (Fig. 17), the signal of potential flight path y t which may be obtained in conventional manner from the actual 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 varia-tion 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 . -- 19 . ., .
~ 1~88655 ''`
minimum and maximum values of the variation. The amplitude limiter 92 is connected to the adder 88 whose output is con-nected to a subtractor 93, which receives a signal corresponding to the aerodynamic flight path angle y a at its second input 94.
The signal ~, provided at the output of the subtractor 93 is transmitted to a voter 95, which controls the pitch command bar of the attitude director indicator 87.
The voter 95 comprises three inputs, whereof one 96 is .` connected to the subtractor 93, the second 97 is connected to a 10 circuit providing a signal ~1 proportional to the difference 91 ~ 18 (~1 being the longitudinal pitch attitude), whereas the third input 98 receives a signal K (ya - yo) emanating from an ~; amplifier 99 for an adjustment gain K, which receives the differ-; ential signal of a subtractor 100. On the one hand, this sub-tractor 100 receives a signal representative of the aerodynamic : flight path angle ya taken for example from 94 and on the other - hand, a signal yo representative of an angle of predetermined value, for example 0.5.
The signal provided by the amplifier 99 is also trans-mitted to a subtractor 102 which also receives the output signalof the voter 95.
The differential signal provided by the subtractor 102 is transmitted to a logic unit 103 which acts on the attitude direc-tor indicator 87, in order to cause the 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 des-cribed 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 con-trolled by a circuit for detecting an engine failure, which enables the reference speed to assume two values, namely the . . ~
' ' :.
:
l~ss6ss ~- 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 possibl~ to illustrate the action of the afore-described protection device, ,. .
.j~ 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 r,` aircraft, a curve which comprises: a first part 105 correspond-~j 10 ing to the phase during which the pilot seeks to reach the speed ,;:
V2 + 10 knots or the pitch attitude ~1 = 18; a second part 106 ~'?;'`'~ corresponding to holding a speed of V2 + 10 knots; a third part 107 in which are located in broke line the trajectory 108 cor-~ responding to normal operation, in full line the trajectory 104 : i corresponding to the system according to the main patent dis-, turbed by poor operation, in broken line the trajectory 109 cor-' rected by the protection action according to the invention in the case of poor operation.
This diagram also shows, in correlation with the trajec-~;i 20 tories 104, 108 and 109, the curves 111, 112 and 113 representing ; , the signals ~ and ~2~ at the input of the voter and a hori-zontal line 114 corresponding to a constant signal ~'2 indicating a considerable and eonstant order to pitch down.
,~:
` It will thus be seen that after reaching the speed V2 +
10 knots or ~1 = 18, during the phase of holding this speed, j the signals ~ and ~2 are maintained at a substantially con-stant level and in the order ~2>~>~1 and, consequently, it is ` the signal ~ which is selected (curve shown in broken line 115).
: .
; As soon as a failure occurs in the circuit producing the signal ' 30 ~ and which is translated by an order to pitch down and by a ; deerease in the flight path angle of the aircraft, the value of the signal ~2 = (ro - ~a) K decreases and, at the end of a cer-. .
,' ~
:, 1~88655 :
tain period of time, becomes less than the value of the signal ~.
In other words, there is an intersection of the curves ~ and ~2.
Beyond the point where the value of ~2 becomes less than the value of ~, the value of ~2 is located between that of ~ and ~l and due to this, it is the signal ~2 which is selected (see curve shown in broken line 115).
Consequently, the flying system is governed by K(ya - yo), which makes it possible to maintain an aerodynamic flight path angle ya equal to yo. In practice, the value of yo which consti-tutes the bottom aerodynamic flight path angle, below which one should not fall, is approximately 0.5.
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 pitch command bar of the attitude director indicator 87 to disappear, when the signal ~2 is selected at the end of a predetermined time, for example five seconds.
;~! 20 It is clear that this disappearance of the 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 (22)
PROPERTY OR PRIVILEGE IS CLAIMED ARE DEFINED AS FOLLOWS:
1. A method of deriving a signal indicative of the manner in which an aircraft is to be flown using total energy, comprising the following steps, governing the aerodynamic flight path angle .gamma. a by reference to a desired flight path angle .gamma. d, obtaining the desired flight path angle .gamma. d by modulating the potential flight path .gamma. t by the difference between the aircraft speed V and a reference speed V2, and displaying an error signal .delta. representative of the difference between the aerodynamic flight path angle .gamma. a and the desired flight path angle .gamma. d.
2. A method according to claim 1, wherein said step of displaying includes displaying said error signal .delta. by means of the pitch command bar of an attitude director indicator.
3. A method according to claim 1, and further includ-ing calculating the potential flight path .gamma. t from an accelero-metric box.
4. A method according to claim 1, and further including calculating the potential flight path .gamma. t 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 method according to claim 4, and further including calculating .gamma. t from the formula:
(g being the acceleration due to gravity), or the formula:
(g being the acceleration due to gravity), or the formula:
6. A method according to claim 5 and further including governing the derivative 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 method according to claim 1, and further including including taking into account either the variation .delta., intended to subject the speed V to the control of the reference value, or a variation of the pitch attitude with respect to a reference position, and including transmitting the error signal .delta. to a voter, comparing .delta. in the voter with a signal .delta.1 corresponding to said variation of attitude and with a constant reference signal .delta.2 and controlling the pitch command bar of the attitude director indicator by the intermediate signal of the group (.delta., .delta., .delta.2) i.e.
the signal whose instantaneous value is between those of the other two signals.
the signal whose instantaneous value is between those of the other two signals.
8. A method according to claim 7, and utilizing a refer-ence attitude which corresponds to .theta.1 = 18° and a constant refer-ence signal which corresponds to 20°.
9. A 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 potential flight path .gamma. t is above or below a calibrated value.
10. A method according to claim 9, and utilizing a calibrated value of 6°.
11. A method according to claim 1, and including limit-ing the speed difference signal between positive and negative values.
12. A system for deriving a signal indicative of the manner in which an aircraft is to be flown using total energy comprising the steps of governing the flight path angle .gamma. a by reference to a desired flight path angle .gamma. d, obtaining the desired flight path angle .gamma. d by modulating the potential flight path angle .gamma. c by the difference between the aircraft speed V
and a reference speed V2, and displaying an error signal .delta. repre-sentative of the difference between the aerodynamic flight path angle .gamma. a and the desired flight path angle .gamma. d, and wherein a signal representative of the total potential flight path angle .gamma. c 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 in-put, a signal corresponding to the aerodynamic flight path angle .gamma. a, the subtractor output being transmitted to a voter which is connected to a member for controlling the pitch command bar of the attitude director indicator and the voter also receiving a signal representative of a variation of pitch attitude and a reference signal.
and a reference speed V2, and displaying an error signal .delta. repre-sentative of the difference between the aerodynamic flight path angle .gamma. a and the desired flight path angle .gamma. d, and wherein a signal representative of the total potential flight path angle .gamma. c 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 in-put, a signal corresponding to the aerodynamic flight path angle .gamma. a, the subtractor output being transmitted to a voter which is connected to a member for controlling the pitch command bar of the attitude director indicator and the voter also receiving a signal representative of a variation of 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 pitch atti-tude .theta.1 and the aerodynamic angle of attack .gamma. , this circuit being able to comprise a commutation device controlled by a sig-nal representing the state of the main undercarriage, in order to substitute the pitch attitude .theta.1 for the aerodynamic angle of attack .gamma. when said main undercarriage is compressed, in order to cancel out said difference.
14. A system according to claim 12, wherein the signal corresponding to the variation between the speed V and the refer-ence 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 potential flight path .gamma. t, the second 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 .gamma. t 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 voter is connected to a member for controlling the pitch command bar of the attitude director indicator by means of a limiter.
18. A system according to claim 12, wherein, in order to ensure protection against abnormal negative flight path angle of the aircraft, due in particular to poor operation of the flying system circuit providing the signal .delta., the input of the voter corresponding to the signal .delta.2 is connected to a circuit providing a signal K ( .gamma.a - .gamma.o), in which K is an adjustment gain and .gamma. o 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 voter the presence of the signal .delta.2 when the latter is selected, and to disconnect the entire flying system when the signal .delta.2 remains selected at the end of a predetermined period of time.
20. A system according to claim 12, wherein the input .delta.2 of the voter is connected to an amplifier for an adjust-ment gain K, which receives the differential signal of a sub-tractor, which in turn receives a signal representative of the aerodynamic flight path angle .gamma. a and, a signal .gamma. o representa-tive of a flight path angle of a predetermined value.
21. A system according to claim 20, wherein the pre-determined 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 voter, and the differential signal provided by said subtractor is transmitted to a logic unit which acts on the attitude director indicator in order to cause the pitch command bar to disappear when said differential signal remains zero for a period of time greater than a predeter-mined period.
Applications Claiming Priority (6)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| FR7521137A FR2316647A1 (en) | 1975-07-04 | 1975-07-04 | Aircraft total energy control system - has aerodynamic and selected climb or inclination modulated with respect to deviations in velocity |
| FR7521.137 | 1975-07-04 | ||
| FR7537862A FR2334993A2 (en) | 1975-12-10 | 1975-12-10 | Aircraft total energy control system - has aerodynamic and selected climb or inclination modulated with respect to deviations in velocity |
| FR7537,862 | 1975-12-10 | ||
| FR7615.929 | 1976-05-26 | ||
| FR7615929A FR2353090A2 (en) | 1976-05-26 | 1976-05-26 | Aircraft total energy control system - has aerodynamic and selected climb or inclination modulated with respect to deviations in velocity |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1088655A true CA1088655A (en) | 1980-10-28 |
Family
ID=27250494
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA256,222A Expired CA1088655A (en) | 1975-07-04 | 1976-07-02 | Flying method and system using total power for an aircraft |
Country Status (4)
| Country | Link |
|---|---|
| CA (1) | CA1088655A (en) |
| DE (1) | DE2630651A1 (en) |
| GB (1) | GB1553407A (en) |
| IT (1) | IT1067300B (en) |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4609987A (en) * | 1983-03-09 | 1986-09-02 | Safe Flight Instrument Corporation | Aircraft guidance system for take off or go-around during severe wind shear |
| US10302451B1 (en) * | 2018-02-20 | 2019-05-28 | The Boeing Company | Optimizing climb performance during takeoff using variable initial pitch angle target |
| CN112947520B (en) * | 2021-02-08 | 2023-02-28 | 北京电子工程总体研究所 | Attitude control method and device for improving stability of low-speed aircraft under stall |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1102781A (en) * | 1965-01-20 | 1968-02-07 | Smiths Industries Ltd | Improvements in or relating to aircraft instruments |
| US3748980A (en) * | 1971-12-06 | 1973-07-31 | Polaroid Corp | Flash socket assembly |
-
1976
- 1976-06-28 GB GB26757/76A patent/GB1553407A/en not_active Expired
- 1976-07-02 IT IT24989/76A patent/IT1067300B/en active
- 1976-07-02 CA CA256,222A patent/CA1088655A/en not_active Expired
- 1976-07-05 DE DE19762630651 patent/DE2630651A1/en not_active Withdrawn
Also Published As
| Publication number | Publication date |
|---|---|
| IT1067300B (en) | 1985-03-16 |
| DE2630651A1 (en) | 1977-01-20 |
| GB1553407A (en) | 1979-09-26 |
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