(19)
(11)EP 1 814 006 B1

(12)EUROPEAN PATENT SPECIFICATION

(45)Mention of the grant of the patent:
21.09.2016 Bulletin 2016/38

(21)Application number: 06001510.4

(22)Date of filing:  25.01.2006
(51)Int. Cl.: 
G05D 1/00  (2006.01)

(54)

Minimizing dynamic structural loads of an aircraft

Minimierung der dynamischen strukturellen Lasten eines Flugzeuges

Minimisation des charges structurelles dynamiques d'un aéronef


(84)Designated Contracting States:
AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LI LT LU LV MC NL PL PT RO SE SI SK TR

(43)Date of publication of application:
01.08.2007 Bulletin 2007/31

(73)Proprietor: Airbus Opérations SAS
31060 Toulouse (FR)

(72)Inventors:
  • Wildschek, Andreas
    85521 Riemerling (DE)
  • Maier, Rudolf, Dr.
    83714 Miesbach (DE)
  • Hoffmann, Falk
    85625 Glonn (DE)
  • Jeanneau, Matthieu
    31200 Toulouse (FR)
  • Aversa, Nicky
    31530 Montaigut S/Save (FR)

(74)Representative: Willems, Volker 
Airbus Defence and Space GmbH Patentabteilung
81663 München
81663 München (DE)


(56)References cited: : 
EP-A- 1 528 448
US-A- 4 796 192
DE-A1- 19 841 632
US-A1- 2004 079 835
  
      
    Note: Within nine months from the publication of the mention of the grant of the European patent, any person may give notice to the European Patent Office of opposition to the European patent granted. Notice of opposition shall be filed in a written reasoned statement. It shall not be deemed to have been filed until the opposition fee has been paid. (Art. 99(1) European Patent Convention).


    Description


    [0001] The invention relates to minimizing dynamic structural loads of an aircraft, introduced by an external excitation to the same.

    [0002] Dynamic structural loads are introduced to the structure of an aircraft by external excitations, e.g. due to wind, gusts, turbulences or similar influences, and by excitations due to pilot/flight control system demands. As a result of the flexible properties of the aircraft structure, such excitation may result in oscillations or vibrations that may exceed a given magnitude, so that they can be harmful to the aircraft structure, especially of frequencies in the range of natural or characteristic oscillations of the flexible aircraft structure. If it is desired to reduce the mass of the aircraft structure at high aspect ratio of wing and fuselage, it is necessary to take measures against excessive dynamic structural loads.

    [0003] From the document DE 198 41 632 C2 is known a method for compensating structural oscillations of an aircraft, which are introduced by an external excitation, comprising detecting at least one body rate of the aircraft with a sensor arrangement, providing said at least one body rate to a flight controller, and producing movements of control surfaces of the aircraft to minimize the excited oscillations.

    [0004] Further it is known from US 5 515 444 an active noise control system for reducing aircraft engine duct noise of a turbofan engine of which a compressor generates a noisy primary sound field. The active noise control system comprises a blade passage sensor mounted within the engine adjacent to the fan for generating a reference acoustic signal, which is correlated with the radiated sound, and a distributed error sensor positioned to be responsive to the primary sound field for generating an error acoustic signal. Acoustic driver means comprised of an array or piezoelectric driven panels are mounted within the fan inlet of the engine for compensating said noisy primary sound field. A controller is responsive to the reference acoustic signal and the error acoustic signal for driving the acoustic driver means to generate a secondary sound field having approximately equal amplitude but opposite phase as said primary sound field to thereby effectively reduce the engine noise.

    [0005] The US 4 796 192 discloses a system for reducing the wing root bending moment load during aircraft manoeuvres by reducing the lift on the outboard section of the wing in response to commands calculated from aircraft parameters and a predetermined wing root bending moment load limit.

    [0006] The EP 1 528 448A1 discloses an active multi-axis modal suppression system for suppressing vibrations of an aircraft due to excitation of the natural modes of the aircraft body. Dedicated sensors are positioned in the aircraft at different locations for sensing the modal induced vibrations, The signals produced by said sensors are processed through control logic which, in response thereto, and in response to aircraft inertial, velocity and altitude related signals produces output control signals. The control signals effect control surface deployment creating forces to suppress the natural mode induced vibrations on multiple geometric axes so as to damp lateral, longitudinal, vertical and torsional vibrational modes.

    [0007] Finally, the US 2004/0079835 A1 discloses a control system for alleviating gust loads on an aircraft wing, comprising deflector mechanisms arranged adjacent to the wing tips. These are controlled in response to signals measured by vertical motion sensors, indicating the vertical motion of the wing tips. A sensor signal processor generates a deflector control signal in response to the sensor output signals and outputs the same to a deflector controller for regulating the deflector mechanism movement in response to the deflector control signal such that the deflector mechanism is alternately deployed into and out of the airstream for alleviating the gust loads on the aircraft wing.

    [0008] The object of the present invention is to provide for effective minimizing dynamic structural loads of an aircraft.

    [0009] The object is achieved by a method of minimizing dynamic structural loads of an aircraft introduced by an external excitation according to claim 1. Further, the object is achieved by an apparatus for minimizing dynamic structural loads of an aircraft structure introduced by an external excitation according to claim 15.

    [0010] Additional features of the invention are included in the respective dependent claims.

    [0011] The invention provides for a method of minimizing dynamic structural loads of an aircraft, introduced by an external excitation to the aircraft, comprising generating a signal indicative of the external excitation; deriving signals of pre-controlling for actuating control elements of the aircraft from said excitation indicating signal in accordance with a pre-controlling rule, so as to reduce the dynamic structural loads introduced to the aircraft; generating an error signal representing performance of said pre-controlling, wherein said error signal is indicative of structural loads of the aircraft structure being introduced to the same by said external excitation; and optimizing the pre-controlling rule by said error signal and/or said excitation indicating signal so as to minimize the dynamic structural loads as they are indicated by said error signal.

    [0012] Generating said excitation indicating signal may include detecting signals indicating one or more of intensity and direction of turbulence, wind and gusts, angle of attack, angle of yaw, and Euler-angles. Information of pilot/FCS demand may be added to excitation indicating signal in order to minimize pilot-induced loads/vibrations.

    [0013] Generating said error signal may include detecting signals indicating one or more of accelerations, stresses or strains at given locations of the aircraft structure.

    [0014] The generating of said error signal may include subtracting contributions of pilot or flight controller commands to control surfaces of the aircraft, which are included in the structural loads. Said subtracting the influence of pilot or flight controller commands may be carried out on the basis of a rigid body model of the aircraft.

    [0015] Said excitation indicating signal may be generated from a flexible body model of the aircraft, or an observer/Kalman filter.

    [0016] Said optimizing the pre-controlling rule may include frequency separation for subtracting the influence of pilot or flight controller commands. Optimizing the pre-controlling rule may include an iterative algorithm.

    [0017] Said actuating of control elements so as to minimize dynamic structural loads may include actuating one or more of elevator, rudder, aileron or other control surfaces of the aircraft. Said actuating of control elements so as to minimize dynamic structural loads may include actuating one or more of electro-mechanical, electro-magnetic, hydraulic, pneumatic or piezoelectric actuators to introduce load damping forces directly into the aircraft structure.

    [0018] Optimizing said pre-controlling rule may include generating a transfer function or any other mathematical model, of the aircraft.

    [0019] Said generating of the transfer function may be carried out by an online-system-identification comprising filtering band-limited random noise or chirp signal or any other identification signal by a number of n adaptive filters, wherein n corresponds to the number of the aircraft transfer functions to be identified, and subtracting the filtering response in vector form from the error signal. System identification may include estimating the primary control path. (For structural vibration/load reduction of an aircraft the secondary control path can be easily estimated from the primary control path, and the primary control path is easier to identify than the secondary path.)

    [0020] Further the present invention is providing for an apparatus for minimizing dynamic structural loads of an aircraft structure introduced by an external excitation to the aircraft, comprising an excitation signal generating arrangement for generating a signal indicative of an external excitation of the aircraft structure; a regulating circuit for deriving pre-controlling signals from said excitation indicating signal in accordance with a pre-controlling rule, for actuating control elements of the aircraft so as to reduce the dynamic structural loads introduced to the aircraft; an error signal generating arrangement for generating a signal representing performance of said pre-controlling as an error signal, said error signal being indicative of structural loads of the aircraft structure introduced to the same by said external excitation; and an optimizing circuit for optimizing the pre-controlling rule by said error signal and/or excitation indicating signal so as to minimize the dynamic structural loads as they are indicated by said error signal.

    [0021] Said excitation signal generating arrangement may include sensor means for detecting signals indicating one or more of intensity and direction of turbulence, wind and gusts, angle of attack, angle of yaw, and Euler-angles.

    [0022] Said error signal generating arrangement may be provided for detecting signals indicating one or more of accelerations, stresses or strains at given locations of the aircraft structure. Further, the error signal generating arrangement may be provided for subtracting contributions of pilot or flight controller commands to control surfaces of the aircraft, which are included in the structural loads.

    [0023] Said error signal generating arrangement may also be provided for subtracting the influence of pilot or flight controller commands on the basis of a rigid body model of the aircraft.

    [0024] Said excitation signal generating arrangement may be provided for generating the excitation indicating signal from a flexible body model of the aircraft, or observer/Kalman filter.

    [0025] Said optimizing circuit may include a frequency separator for optimizing the pre-controlling rule by frequency separation in order to be able to cancel out frequency ranges, where controller shall not have any influence. Said optimizing circuit may be provided for optimizing the pre-controlling rule by an iterative algorithm.

    [0026] Control elements, which are actuated so as to minimize dynamic structural loads may include one or more of elevator, rudder, aileron or other control surfaces of the aircraft. Further, the control elements, which are actuated so as to minimize dynamic structural loads, may include one or more electro-mechanical, electromagnetic, hydraulic, pneumatic or piezoelectric actuators for introducing load damping forces directly into the aircraft structure.

    [0027] Said optimizing circuit may be provided for generating a transfer function of the aircraft for optimizing said pre-controlling rule. The optimizing circuit may also be provided for generating the transfer function by an online-system-identification which comprises filtering band-limited random noise or a chirp signal by a number of n adaptive filters, wherein n corresponds to the number of the aircraft transfer functions to be identified, and for subtracting the filtering response in vector form from the error signal.

    [0028] In the following are disclosed embodiments of the present invention with reference to the drawings, in which:

    Fig. 1 is a block diagram for disclosing method and apparatus of minimizing dynamic structural loads introduced by an external or pilot/FCS excitation to an aircraft, in accordance with one embodiment of the invention;

    Fig. 2 is a block diagram for describing the generation of an error signal, which is used in the present invention;

    Fig. 3 is a schematic block diagram for describing a further embodiment of method and apparatus of minimizing dynamic structural loads of an aircraft, introduced by an external excitation, in accordance with the present invention;

    Fig. 4 is a schematic block diagram of an adaptive pre-regulation or precontrolling circuit used in one embodiment of the present invention;

    Fig. 5 is another schematic block diagram for describing an adaptive pre-controlling;

    Fig. 6 is a schematic view of an aircraft, showing a sensor arrangement which can be used in accordance with one embodiment of the present invention for minimizing dynamic structural loads of an aircraft; and

    Fig. 7 is a diagram, which shows magnitude of an averaged error signal, which is indicative of structural loads introduced by an external excitation to an aircraft, versus frequency.



    [0029] Fig. 1 shows one principle embodiment of minimizing dynamic structural loads of an aircraft 2, which are introduced by an external excitation 1. The external excitation 1 may e.g. be wind, gusts or turbulences. The excitation 1 causes dynamic structural loads in the structure of the aircraft 2, especially by natural or characteristic oscillations within the structure due to flexible body properties of the same. In Fig. 1, the excitation 1 is introduced to the aircraft 2, which is regarded as a flexible body.

    [0030] The excitation 1 is detected by an excitation or reference sensor 5, which is generating a signal x indicative of the external excitation 1, the reference sensor 5 being in this case one or more sensors or a sensor arrangement. The signal x indicative of the external excitation 1 is input to an adaptive pre-control circuitry 6, 17 for deriving control signals y for actuating control elements of the aircraft 2 in accordance with a pre-controlling rule, which is implemented in said adaptive pre-control circuitry 16, 17. The control signals y for actuating control elements reduce the dynamic structural loads, which are introduced to the aircraft 2 by the external excitation 1. The control signals y are output to the aircraft 2 so as to actuate the control elements, which are provided in the same.

    [0031] In the aircraft 2 is generated a signal indicative of structural loads which are introduced to the same by said external excitation 1, which signal is used as an error signal e, e* and which is describing control performance of the pre-controlling rule of circuitry 6, 17. The error signal e, e* is input to a circuitry 8, 9 for optimizing the pre-controlling rule by said error signal e, e*, so as to minimize the dynamic structural loads, i.e. to reduce the amplitude of the oscillations within the structure of the aircraft 2. The optimization in circuitry 8, 9 is performed in a way that a minimization of the oscillations and of the dynamic structural loads is adapted to actual load and flight conditions of the aircraft 2. If the optimizing rule is not iterative, the signal x is also used for optimizing the adjustable pre-controlling rule of circuitry 6, 17.

    [0032] The reference sensor 5 may comprise a number of sensors, which detect intensity and direction of wind, turbulence and gusts, angle of attack, angle of yaw, etc., e.g. as Euler-angles. The error signal e, e* which is detected by error sensors in the aircraft 2, may include accelerations at different given locations of the aircraft structure, as well as stresses or strains at given locations at the aircraft structure, for example. The control elements of the aircraft 2, which are actuated in order to minimize the dynamic structural loads, may include elevator, rudder, aileron or other control surfaces, and they may include electro-mechanical, electromagnetic, hydraulic, pneumatic or piezoelectric actuators for directly introducing load damping forces into the aircraft structure.

    [0033] Further, as a specific option, also pilot or flight control system commands 3, 4, which may also induce unwanted structural loads and vibrations in the aircraft 2, are added to the excitation indication signal x. Thus, also structural loads and vibrations of the aircraft due to pilot or flight system commands are minimized.

    [0034] The optimization of the pre-controlling rule may be carried out in accordance with an appropriate method, i.e. methods of least mean square root errors or recursive methods of least mean square root errors (LMS, RLS) or methods of minimizing quadratic cost function.

    [0035] Fig. 2 shows the generation of a filtered error signal e. There an internal model 16 of the aircraft 2 is provided for subtracting contributions of pilot or flight controller commands 3, 4, which are output to the control surfaces of the aircraft and causing structural loads by the corresponding flight maneuver. However, these commands shall not result in an activity of the pre-controlling circuitry 6, 17. That means that, of course, pilot or flight controller commands will generally not be compensated by the pre-controlling circuitry 6, 17, but pilot-induced vibrations will be compensated. For this purpose internal model 16 is a rigid body model.

    [0036] The signals output from the internal aircraft model 16 are combined in a combining circuit 25 with the original error signal e* to generate the filtered error signal e which is output to optimizing circuitry 8, 9. Alternatively the control commands y for the control surfaces of the aircraft can be input to the internal model 16, and the signal x indicative of the external excitation can be taken from the combining circuit 25.

    [0037] If it is intended to use the internal model 16 of the aircraft 2 only for subtracting the influence of pilot or flight controller commands from the original error signal e*, the internal model 16 can be a rigid body model. However, if the internal model 16 is intended to generate the external excitation indicating signal x, it has to be a complete flexible body model of the aircraft 2. The internal model may also be used to subtract parasitic feedback from the excitation indicating signal.

    [0038] Fig. 3 shows in a schematic block diagram one embodiment of an apparatus for minimizing dynamic structural loads of an aircraft. The external excitation 1 which may be wind, gusts or turbulences, is introduced to the structure of the aircraft 2. The excitation 1 causes vibrations and oscillations, especially natural or characteristic oscillations within the aircraft structure. The aircraft 2 is controlled by pilot command 3 and by flight controller 4. The wind excitation 1 is detected by reference sensor 5. This can be a laser-optical sensor or other wind sensors (e.g. α-sensor, β-sensor).

    [0039] This excitation or reference signal, which includes in general a three-dimensional information of velocity and direction of wind, is forwarded to pre-controlling circuitry 6 for actuating the control elements of the aircraft 2 so as to reduce and minimize the structural oscillations and the structural loads of the aircraft 2. The control elements can be, as already described above, elevator, rudder, aileron or other, aerodynamically effective control surfaces, and/or they can be mechanical control elements so as to minimize the structural loads directly, electro-mechanical, electromagnetic, hydraulic, pneumatic or piezoelectric actuators to introduce load damping forces directly into the aircraft structure.

    [0040] For adapting and optimizing, said pre-controlling circuitry 6 has input one or more error signals generated by error sensors 7, which error signals are detected in the aircraft 2 and which are to be minimized, and further information concerning transfer functions of the aircraft 2, which is provided by an online-system-identification 9, and information concerning pilot and flight controller commands, which shall not be compensated by the structural load minimizing system.

    [0041] The online-system-identification 9 is connected to a random noise generator 8, which is generating a band-limited random noise for each actuator. None of the random noise signals must be correlated with the excitation signal or reference signal. The transfer functions for different actuators should not be measured at the same time to provide for a good signal-to-noise ratio. The noise signals are forwarded to the actuators or control elements for minimizing the structural loads and to the system-identification 9. In the system-identification 9, the identification noise signals pass through n adaptive filters (e.g. FIR, IIR, Neuronal Networks, etc.), wherein n is the number of the aircraft transfer functions to be identified.

    [0042] The entity of those adaptive filters is providing a filter model of the aircraft 2. The responses 10 of this filter model to the identification noise signals are subtracted in vector form by subtraction means 11 from the error signals generated by error sensors 7 of the aircraft 2. Those parts of the error signals, which correlate with the identification noise signals, are the responses of the aircraft 2 to the identification noise signals. The result of this vector subtraction represents the error between aircraft transfer functions and filter model.

    [0043] Filter coefficients 12 are frequency field transformed and forwarded to the adaptive pre-controlling circuitry 6. The adaptation of the filter model coefficients 12 can be done by a LMS- or RLS-algorithm or by another algorithm. This makes use of the identification noise signals generated by generator 8 to be sure that in the system identification 9 only those parts in the signals output from the error sensors 7 are considered, which are correlating with the identification noise signals. Accelerations, body rates, etc. of the aircraft 2 are also input to the flight controller 4.

    [0044] One embodiment of an adaptive pre-controlling circuitry 6 is shown in Fig. 4. Both, the error signals of the aircraft 2 generated by error sensors 7, and the reference signals of the reference sensor 5 (Fig. 3) include in general information caused by pilot and flight controller commands. To avoid a not intended minimizing or compensating of pilot and flight controller commands by the dynamic structural load minimizing system, contributions of pilot and flight controller commands are subtracted from the reference and error signals by combining circuits 14 and15, respectively. The contributions or the influence of pilot and flight controller commands 3, 4 to the signals of reference sensors 5 and error sensors 7 is computed by means of internal rigid body model 16 of the aircraft 2. This internal model 16 includes only the flight-mechanical properties of the rigid body aircraft and is usually well known when designing the aircraft.

    [0045] The compensated error and reference signals are input to the adaptive pre-controlling circuitry 17 together with the frequency field coefficients of the filter model. The output of the pre-controlling circuitry 17, usually in vector form, is passed to a plausibility check unit 18 and input to a circuitry 19 for optimal distribution of control commands to the different actuators 20 in order to minimize structural loads/vibrations.

    [0046] One embodiment of an adaptive pre-controlling is given in Fig. 5a. Adaptation of an adaptive filter 24 can be done by means of an iterative disturbance algorithm 23, which is varying filter coefficients by Δw and checking whether the error signal will be smaller or bigger. If the error signal is smaller, the filter coefficient will be varied once more by Δw, else by -Δw, and so on. For such an iterative algorithm no information concerning the transfer functions of the aircraft is necessary, and no online-system-identification.

    [0047] Another embodiment of an adaptive pre-controlling is given in Fig. 5b. A compensated reference signal which includes information concerning the external excitation is input to n adaptive filters 22, wherein n is the number of the structural load damping actuators times the number of the error sensors. The adaptive filters output control commands to plausibility check unit 18 and to circuitry 19 for optimal distribution of the control commands to the actuators 20 or control elements for minimizing the structural loads. The adaptation of the filter coefficient can be done i.e. by a filtered-X-LMS-algorithm 21, which is state of the art. The filtered-X-operation can be carried out with the coefficients of the filter model of the online-system-identification in the frequency field.

    [0048] Fig. 6 shows an embodiment of a filter arrangement of three sensors for detecting accelerations in Z-direction (accelerations perpendicular to the wing surface of the aircraft). The acceleration sensors NzLF and NzRF are at the wing tips, and NzSP is for measuring the Z-acceleration of the center of mass of the aircraft. However, it is also possible to provide acceleration sensors at the front and/or rear part of the fuselage to measure the acceleration in y- and z-directions for damping vertical and horizontal accelerations of the fuselage.

    [0049] An example of minimizing the structural loads introduced to an aircraft by an external excitation by wind is shown in Fig. 7, wherein an averaged error signal is shown as a function of frequency. The dotted line shows the error signal without structural load minimizing, whereas the solid line shows the same with minimizing the structural loads. The maximum peak shows the first vertical bending vibration of the wing, i.e. natural or characteristic oscillation caused by the flexible properties of the same. It can be seen that the bending vibration and, therefore, the structural loads, are significantly reduced.

    List of reference signs



    [0050] 
    1
    excitation
    2
    aircraft
    3
    pilot command
    4
    flight controller commands
    5
    reference sensor
    6
    adaptive pre-controlling circuitry
    7
    error sensor
    8
    band-limited random noise or chirp generator
    9
    system identification
    10
    response of filter model
    11
    subtraction means
    12
    filter coefficients
    14
    combining circuit
    15
    combining circuit
    16
    internal aircraft model
    17
    adaptive pre-control
    18
    plausibility check unit
    19
    control surface driver circuit
    20
    actuators (control surfaces)
    21
    LMS-algorithm
    22
    adaptive filter
    23
    iterative disturbance algorithm
    24
    adaptive filter
    25
    combining circuit
    x
    excitation indication signal
    y
    control signal
    e*
    error signal
    e
    filtered error signal



    Claims

    1. Method of minimizing dynamic structural loads of an aircraft, introduced by an external excitation (1) to the aircraft (2), comprising:

    generating a signal (x) indicative of the external excitation (1);

    deriving signals (y) of pre-controlling for actuating control elements (20) of the aircraft (2) from said excitation indicating signal (x) in accordance with a pre-controlling rule, so as to reduce the dynamic structural loads introduced to the aircraft (2);

    generating an error signal (e, e*) representing performance of said pre-controlling, wherein said error signal (e, e*) is indicative of structural loads of the aircraft structure being introduced to the same by said external excitation (1); and

    optimizing the pre-controlling rule by said error signal (e, e*) and/or by the excitation indicating signal (x) so as to minimize the dynamic structural loads as they are indicated by said error signal (e, e*).


     
    2. Method of claim 1, wherein generating said excitation indicating signal (x) includes detecting signals indicating one or more of intensity and direction of turbulences, wind and gusts, angle of attack, angle of yaw.
     
    3. Method of claim 1 or 2, wherein information about a pilot or flight control system command is added to the excitation indication signal (x) in order to reduce dynamic structural loads caused by such pilot or flight control system command.
     
    4. Method of one of the preceding claims, wherein generating said error signal (e, e*) includes detecting signals indicating one or more of accelerations, stresses or strains at given locations of the aircraft structure.
     
    5. Method of one of the preceding claims, wherein the generating of said error signal (e, e*) includes subtracting contributions of pilot or flight controller commands (3, 4) to control surfaces of the aircraft (2), which are included in the structural loads.
     
    6. Method of claim 5, wherein said subtracting the influence of pilot or flight controller commands (3, 4) is carried out on the basis of a rigid body model (16) of the aircraft (2).
     
    7. Method of one of the preceding claims, wherein said excitation indicating signal (x) is generated from a flexible body model (16) of the aircraft (2), or an observer/Kalman filter.
     
    8. Method of one of the preceding claims, wherein said optimizing the pre-controlling rule includes frequency separation for canceling out certain frequency ranges that shall not be controlled, in particular the frequency range of pilot commands.
     
    9. Method of one of the preceding claims, wherein optimizing the pre-controlling rule includes an iterative algorithm (23).
     
    10. Method of one of the preceding claims, wherein said actuating of control elements (20) so as to minimize dynamic structural loads includes actuating one or more of elevator, rudder, aileron or other control surfaces of the aircraft (2).
     
    11. Method of one of the preceding claims, wherein said actuating of control elements so as to minimize dynamic structural loads includes actuating one or more of electro-mechanical, electro-magnetic, hydraulic, pneumatic or piezoelectric actuators to introduce load damping forces directly into the aircraft structure.
     
    12. Method of one of the preceding claims, wherein optimizing said pre-controlling rule includes generating a transfer function of the aircraft (2).
     
    13. Method of claim 12, wherein the generating of the transfer function is carried out by an online-system-identification (9) comprising filtering band-limited random noise or chirp signal by a number of n adaptive filters (22), wherein n corresponds to the number of the aircraft transfer functions to be identified, and subtracting the filtering response in vector form from the error signal.
     
    14. Apparatus for minimizing dynamic structural loads of an aircraft structure introduced by an external excitation (1) to the aircraft (2), comprising:

    an excitation signal generating arrangement (5) for generating a signal (x) indicative of the external excitation (1) of the aircraft structure;

    a regulating circuit (6, 17) for deriving control signals (y) for actuating control elements of the aircraft (2) from said excitation indicating signal (x) in accordance with a pre-controlling rule, so as to reduce the dynamic structural loads introduced to the aircraft (2);

    characterized by further comprising :

    an error signal generating arrangement (7, 25) for generating a signal representing performance of said pre-controlling as an error signal (e, e*), said error signal (e, e*) being indicative of structural loads of the aircraft structure introduced to the same by said external excitation (1); and

    an optimizing circuit (8, 9) for optimizing the pre-controlling rule by said error signal (e, e*) and/or excitation indicating signal (x) so as to minimize the dynamic structural loads as they are indicated by said error signal (e, e*).


     
    15. Apparatus of claim 14, wherein said excitation signal generating arrangement (5) includes sensor means for detecting signals indicating one or more of intensity and direction of turbulence, wind and gusts, angle of attack, angle of yaw.
     
    16. Apparatus of claim 14 or 15, wherein said error signal generating arrangement (7, 25) is provided with means (7) for detecting signals indicating one or more of accelerations, stresses or strains at given locations of the aircraft structure.
     
    17. Apparatus of one of claims 14 through 16, wherein said error signal generating arrangement (7, 25) is provided with means (25) for subtracting contributions of pilot or flight controller commands (3, 4) to control surfaces of the aircraft (2), which are included in the structural loads.
     
    18. Apparatus of one of claims 14 through 17, wherein said error signal generating arrangement (7,25) is provided for subtracting the influence of pilot or flight controller commands (3, 4) on the basis of a rigid body model (16) of the aircraft (2).
     
    19. Apparatus of one of claims 14 through 18, wherein said excitation signal generating arrangement (5) is provided for generating the excitation indicating signal from a flexible body model (16) of the aircraft (2), or an observer/Kalman filter.
     
    20. Apparatus of one of claims 14 through 19, wherein said optimizing circuit includes a frequency separator for optimizing the pre-controlling rule by frequency separation for subtracting the influence of pilot or flight controller commands (3, 4).
     
    21. Apparatus of one of claims 14 through 20, wherein said optimizing circuit (8, 9) is provided for optimizing the pre-controlling rule by an iterative algorithm.
     
    22. Apparatus of one of claims 14 through 21, wherein control elements (20) which are actuated so as to minimize dynamic structural loads include one or more of elevator, rudder, aileron or other control surfaces of the aircraft (2).
     
    23. Apparatus of one of claims 14 through 22, wherein control elements (20) which are actuated so as to minimize dynamic structural loads include one or more electro-mechanical, electro-magnetic, hydraulic, pneumatic or piezoelectric actuators for introducing load damping forces directly into the aircraft structure.
     
    24. Apparatus of one of claims 14 through 23, wherein said optimizing circuit (8, 9) is provided for generating a transfer function of the aircraft for optimizing said pre-controlling rule.
     
    25. Apparatus of claim 24, wherein the optimizing circuit (8, 9) is provided for generating the transfer function by an online-system-identification (9) which comprises filtering band-limited random noise or chirp signal by a number of n adaptive filters, wherein n corresponds to the number of the aircraft transfer functions to be identified, and for subtracting the filtering response in vector form from the error signal.
     
    26. Apparatus of one of claims 14 through 25, wherein information about a pilot or flight control system command (3, 4) is added to the excitation indication signal (x) in order to reduce dynamic structural loads caused by such pilot or flight control system command.
     


    Ansprüche

    1. Verfahren zum Minimieren der dynamischen strukturellen Belastungen eines Flugzeugs, die durch eine äußere Anregung (1) dem Flugzeug (2) zugeführt werden, das die folgenden Schritte aufweist:

    Erzeugen eines Signals (x), das die äußere Anregung (1) anzeigt;

    Ableiten von Signalen (y) einer Vorsteuerung zum Betätigen von Steuerelementen (20) des Flugzeugs (2) aus dem Anregungsanzeigesignal (x) gemäß einer Vorsteuerungsregel, um die dynamischen strukturellen Belastungen zu reduzieren, die dem Flugzeug (2) zugeführt werden;

    Erzeugen eines Fehlersignals (e, e*), das die Leistung der Vorsteuerung repräsentiert, wobei das Fehlersignal (e, e*) die strukturellen Belastungen der Flugzeugstruktur anzeigt, die derselben durch die äußere Anregung (1) zugeführt werden; und

    Optimieren der Vorsteuerungsregel durch das Fehlersignal (e, e*) und/oder durch das Anregungsanzeigesignal (x), um die dynamischen strukturellen Belastungen zu minimieren, wie sie durch das Fehlersignal (e, e*) angezeigt werden.


     
    2. Verfahren nach Anspruch 1, wobei das Erzeugen des Anregungsanzeigesignals (x) das Detektieren von Signalen umfasst, die die Intensität und Richtung von Turbulenzen, Wind und Böen, und/oder den Anströmwinkel und/oder den Gierwinkel anzeigen.
     
    3. Verfahren nach Anspruch 1 oder 2, wobei Informationen über einen Piloten- oder Flugleitsystembefehl zum Anregungsanzeigesignal (x) hinzugefügt werden, um die dynamischen strukturellen Belastungen zu reduzieren, die durch einen solchen Piloten- oder Flugleitsystembefehl verursacht werden.
     
    4. Verfahren nach einem der vorhergehenden Ansprüche, wobei das Erzeugen des Fehlersignals (e, e*) das Detektieren von Signalen umfasst, die Beschleunigungen und/oder Spannungen oder Dehnungen an vorgegebenen Stellen der Flugzeugstruktur anzeigen.
     
    5. Verfahren nach einem der vorhergehenden Ansprüche, wobei das Erzeugen des Fehlersignals (e, e*) das Subtrahieren von Beiträgen von Piloten- oder Flugleitsystembefehlen (3, 4) zu Steuerflächen des Flugzeugs (2) umfasst, die in den strukturellen Belastungen enthalten sind.
     
    6. Verfahren nach Anspruch 5, wobei das Subtrahieren des Einflusses von Piloten- oder Flugleitsystembefehlen (3, 4) auf der Grundlage eines Modells (16) eines starren Körpers des Flugzeugs (2) ausgeführt wird.
     
    7. Verfahren nach einem der vorhergehenden Ansprüche, wobei das Anregungsanzeigesignal (x) aus einem Modell (16) eines flexiblen Körpers des Flugzeugs (2) oder einem Beobachter/Kalman-Filter erzeugt wird.
     
    8. Verfahren nach einem der vorhergehenden Ansprüche, wobei das Optimieren der Vorsteuerungsregel die Frequenztrennung zum Auslöschen bestimmter Frequenzbereiche, die nicht gesteuert werden sollen, insbesondere des Frequenzbereichs der Pilotenbefehle umfasst.
     
    9. Verfahren nach einem der vorhergehenden Ansprüche, wobei das Optimieren der Vorsteuerungsregel einen iterativen Algorithmus (23) umfasst.
     
    10. Verfahren nach einem der vorhergehenden Ansprüche, wobei das Betätigen von Steuerelementen (20), um die dynamischen strukturellen Belastungen zu minimieren, das Betätigen des Höhenruders und/oder des Seitenruders und/oder des Querruders und/oder anderer Steuerflächen des Flugzeugs (2) umfasst.
     
    11. Verfahren nach einem der vorhergehenden Ansprüche, wobei das Betätigen von Steuerelementen, um die dynamischen strukturellen Belastungen zu minimieren, das Betätigen elektromechanischer und/oder elektromagnetischer und/oder hydraulischer und/oder pneumatischer und/oder piezoelektrischer Stellglieder umfasst, um direkt in die Flugzeugstruktur belastungsdämpfende Kräfte einzuführen.
     
    12. Verfahren nach einem der vorhergehenden Ansprüche, wobei das Optimieren der Vorsteuerungsregel das Erzeugen einer Übertragungsfunktion des Flugzeugs (2) umfasst.
     
    13. Verfahren nach Anspruch 12, wobei das Erzeugen der Übertragungsfunktion durch eine Online-Systemidentifikation (9) ausgeführt wird, die das Filtern von bandbegrenztem weißem Rauschen oder eines Chirp-Signals durch eine Anzahl von n adaptiven Filtern (22), wobei n der Anzahl der Flugzeugübertragungsfunktionen entspricht, die identifiziert werden sollen, und das Subtrahieren der Filteranwort in Vektorform vom Fehlersignal umfasst.
     
    14. Vorrichtung zum Minimieren der dynamischen strukturellen Belastungen einer Flugzeugstruktur, die durch eine äußere Anregung (1) dem Flugzeug (2) zugeführt werden, die Folgendes aufweist:

    eine Anregungssignal-Erzeugungsanordnung (5) zum Erzeugen eines Signals (x), das die äußere Anregung (1) der Flugzeugstruktur anzeigt;

    eine Regelschaltung (6, 17) zum Ableiten von Steuersignalen (y) zum Betätigen von Steuerelementen des Flugzeugs (2) aus dem Anregungsanzeigesignal (x) gemäß einer Vorsteuerungsregel, um die dynamischen strukturellen Belastungen zu reduzieren, die dem Flugzeug (2) zugeführt werden;

    dadurch gekennzeichnet, dass sie ferner Folgendes aufweist:

    eine Fehlersignal-Erzeugungsanordnung (7, 25) zum Erzeugen eines Signals, das die Leistung der Vorsteuerung repräsentiert, als ein Fehlersignal (e, e*), wobei das Fehlersignal (e, e*) strukturelle Belastungen der Flugzeugstruktur anzeigt, die derselben durch die äußere Anregung (1) zugeführt werden; und

    eine Optimierungsschaltung (8, 9) zum Optimieren der Vorsteuerungsregel durch das Fehlersignal (e, e*) und/oder das Anregungsanzeigesignal (x), um die dynamischen strukturellen Belastungen zu minimieren, wie sie durch das Fehlersignal (e, e*) angezeigt werden.


     
    15. Vorrichtung nach Anspruch 14, wobei die Anregungssignal-Erzeugungsanordnung (5) Sensormittel zum Detektieren von Signalen umfasst, die die Intensität und Richtung von Turbulenzen, Wind und Böen, und/oder den Anströmwinkel und/oder den Gierwinkel anzeigen.
     
    16. Vorrichtung nach Anspruch 14 oder 15, wobei die Fehlersignal-Erzeugungsanordnung (7, 25) mit Mitteln (7) zum Detektieren von Signalen versehen ist, die Beschleunigungen und/oder Spannungen oder Dehnungen an vorgegebenen Stellen der Flugzeugstruktur anzeigen.
     
    17. Vorrichtung nach Anspruch 14 bis 16, wobei die Fehlersignal-Erzeugungsanordnung (7, 25) mit Mitteln (25) zum Subtrahieren von Beiträgen von Piloten- oder Flugleitsystembefehlen (3, 4) an Steuerflächen des Flugzeugs (2) versehen ist, die in den strukturellen Belastungen enthalten sind.
     
    18. Vorrichtung nach Anspruch 14 bis 17, wobei die Fehlersignal-Erzeugungsanordnung (7,25) zum Subtrahieren des Einflusses von Piloten- oder Flugleitsystembefehlen (3, 4) auf der Grundlage eines Modells eines starren Körpers (16) des Flugzeugs (2) vorgesehen ist.
     
    19. Vorrichtung nach Anspruch 14 bis 18, wobei die Anregungssignal-Erzeugungsanordnung (5) zum Erzeugen des Anregungsanzeigesignals aus einem Modell (16) eines flexiblen Körpers des Flugzeugs (2) oder einem Beobachter/Kalman-Filter vorgesehen ist.
     
    20. Vorrichtung nach Anspruch 14 bis 19, wobei die Optimierungsschaltung eine Frequenztrennvorrichtung zum Optimieren der Vorsteuerungsregel durch Frequenztrennung zum Subtrahieren des Einflusses von Piloten- oder Flugleitsystembefehlen (3, 4) umfasst.
     
    21. Vorrichtung nach Anspruch 14 bis 20, wobei die Optimierungsschaltung (8, 9) zum Optimieren der Vorsteuerungsregel durch einen iterativen Algorithmus vorgesehen ist.
     
    22. Vorrichtung nach Anspruch 14 bis 21, wobei Steuerelemente (20), die betätigt werden, um die dynamischen strukturellen Belastungen zu minimieren, das Höhenruder und/oder das Seitenruder und/oder das Querruder und/oder andere Steuerflächen des Flugzeugs (2) umfassen.
     
    23. Vorrichtung nach Anspruch 14 bis 22, wobei Steuerelemente (20), die betätigt werden, um die dynamischen strukturellen Belastungen zu minimieren, elektromechanische und/oder elektromagnetische und/oder hydraulische und/oder pneumatische und/oder piezoelektrische Stellglieder zum direkten Einführen von belastungsdämpfenden Kräften in die Flugzeugstruktur umfassen.
     
    24. Vorrichtung nach Anspruch 14 bis 23, wobei die Optimierungsschaltung (8, 9) zum Erzeugen einer Übertragungsfunktion des Flugzeugs zum Optimieren der Vorsteuerungsregel vorgesehen ist.
     
    25. Vorrichtung nach Anspruch 24, wobei die Optimierungsschaltung (8, 9) zum Erzeugen der Übertragungsfunktion durch eine Online-Systemidentifikation (9) vorgesehen ist, die das Filtern von bandbegrenztem weißem Rauschen oder eines Chirp-Signals durch eine Anzahl von n adaptiven Filtern, wobei n der Anzahl der Flugzeugübertragungsfunktionen entspricht, die identifiziert werden sollen, und das Subtrahieren der Filteranwort in Vektorform vom Fehlersignal umfasst.
     
    26. Vorrichtung nach Anspruch 14 bis 25, wobei Informationen über einen Piloten- oder Flugleitsystembefehl (3, 4) zum Anregungsanzeigesignal (x) hinzugefügt werden, um die dynamischen strukturellen Belastungen zu reduzieren, die durch einen solchen Piloten- oder Flugleitsystembefehl verursacht werden.
     


    Revendications

    1. Procédé de minimisation des charges structurelles dynamiques d'un aéronef, appliquées par une excitation extérieure (1) à l'aéronef (2), comprenant de :

    générer un signal (x) indiquant l'excitation extérieure (1) ;

    dériver des signaux (y) de pré-commande pour actionner des éléments de commande (20) de l'aéronef (2) à partir dudit signal d'indication d'excitation (x) conformément à une règle de pré-commande, de manière à réduire les charges structurelles dynamiques appliquées à l'aéronef (2) ;

    générer un signal d'erreur (e, e*) représentant la performance de ladite pré-commande, dans lequel ledit signal d'erreur (e, e*) indique les charges structurelles de la structure d'aéronef qui sont appliquées à celle-ci par ladite excitation extérieure (1) ; et

    optimiser la règle de pré-commande par ledit signal d'erreur (e, e*) et/ou par le signal d'indication d'excitation (x) de manière à minimiser les charges structurelles dynamiques telles qu'elles sont indiquées par ledit signal d'erreur (e, e*).


     
    2. Procédé selon la revendication 1, dans lequel la génération dudit signal d'indication d'excitation (x) comprend de détecter des signaux indiquant un ou plusieurs parmi l'intensité et la direction de turbulences, le vent et les rafales, l'angle d'attaque, et l'angle de lacet.
     
    3. Procédé selon la revendication 1 ou 2, dans lequel des informations sur une commande de système de commande de pilotage ou de vol sont ajoutées au signal d'indication d'excitation (x) afin de réduire les charges structurelles dynamiques causées par une telle commande de système de commande de pilotage ou de vol.
     
    4. Procédé selon l'une des revendications précédentes, dans lequel la génération dudit signal d'erreur (e, e*) comprend de détecter des signaux indiquant un ou plusieurs parmi des accélérations, des contraintes ou des déformations à des emplacements donnés de la structure d'aéronef.
     
    5. Procédé selon l'une des revendications précédentes, dans lequel la génération dudit signal d'erreur (e, e*) comprend de soustraire des contributions de commandes de contrôleur de pilotage ou de vol (3, 4) à des gouvernes de l'aéronef (2), qui sont incluses dans les charges structurelles.
     
    6. Procédé selon la revendication 5, dans lequel ladite soustraction de l'influence de commandes de contrôleur de pilotage ou de vol (3, 4) est réalisée sur la base d'un modèle de corps rigide (16) de l'aéronef (2).
     
    7. Procédé selon l'une des revendications précédentes, dans lequel ledit signal d'indication d'excitation (x) est généré à partir d'un modèle de corps souple (16) de l'aéronef (2), ou d'un filtre d'observateur/de Kalman.
     
    8. Procédé selon l'une des revendications précédentes, dans lequel ladite optimisation de la règle de pré-commande comprend une séparation des fréquences pour supprimer certaines plages de fréquences qui ne doivent pas être commandées, en particulier la plage de fréquences de commandes de pilotage.
     
    9. Procédé selon l'une des revendications précédentes, dans lequel l'optimisation de la règle de pré-commande comprend un algorithme itératif (23).
     
    10. Procédé selon l'une des revendications précédentes, dans lequel l'actionnement d'éléments de commande (20) de manière à minimiser les charges structurelles dynamiques comprend d'actionner un ou plusieurs d'une gouverne de profondeur, d'une gouverne de direction, d'un aileron ou d'autres gouvernes de l'aéronef (2).
     
    11. Procédé selon l'une des revendications précédentes, dans lequel l'actionnement d'éléments de commande de manière à minimiser les charges structurelles dynamiques comprend d'actionner un ou plusieurs parmi des actionneurs électromécaniques, électromagnétiques, hydrauliques, pneumatiques ou piézoélectriques pour appliquer des forces d'amortissement de charge directement à la structure d'aéronef.
     
    12. Procédé selon l'une des revendications précédentes, dans lequel l'optimisation de ladite règle de pré-commande comprend de générer une fonction de transfert de l'aéronef (2).
     
    13. Procédé selon la revendication 12, dans lequel la génération de la fonction de transfert est réalisée par une identification de système en ligne (9), comprenant de filtrer un signal de bruit aléatoire ou de fluctuation à bande limitée par un certain nombre de n filtres adaptatifs (22), dans lequel n correspond au nombre des fonctions de transfert d'aéronef à identifier, et de soustraire la réponse de filtrage sous forme de vecteur du signal d'erreur.
     
    14. Appareil pour minimiser les charges structurelles dynamiques d'une structure d'aéronef appliquées par une excitation extérieure (1) à l'aéronef (2), comprenant :

    un agencement de génération signal d'excitation (5) pour générer un signal (x) indiquant l'excitation extérieure (1) de la structure d'aéronef;

    un circuit de régulation (6, 17) pour dériver des signaux de commande (y) pour actionner des éléments de commande de l'aéronef (2) à partir dudit signal d'indication d'excitation (x) conformément à une règle de pré-commande, de manière à réduire les charges structurelles dynamiques appliquées à l'aéronef (2);

    caractérisé en ce qu'il comprend en outre :

    un agencement de génération de signal d'erreur (7, 25) pour générer un signal représentant la performance de ladite pré-commande en tant que signal d'erreur (e, e*), ledit signal d'erreur (e, e*) indiquant des charges structurelles de la structure d'aéronef appliquées à celle-ci par ladite excitation extérieure (1) ; et

    un circuit d'optimisation (8, 9) pour optimiser la règle de pré-commande par ledit signal d'erreur (e, e*) et/ou ledit signal d'indication d'excitation (x) de manière à minimiser les charges structurelles dynamiques telles qu'elles sont indiquées par ledit signal d'erreur (e, e*).


     
    15. Appareil selon la revendication 14, dans lequel ledit agencement de génération de signal d'excitation (5) comprend des moyens de capteurs pour détecter des signaux indiquant un ou plusieurs parmi l'intensité et la direction de turbulence, le vent et les rafales, l'angle d'attaque, et l'angle de lacet.
     
    16. Appareil selon la revendication 14 ou 15, dans lequel ledit agencement de génération de signal d'erreur (7, 25) est muni d'un moyen (7) pour détecter des signaux indiquant un ou plusieurs parmi des accélérations, des contraintes ou des déformations à des emplacements donnés de la structure d'aéronef.
     
    17. Appareil selon l'une des revendications 14 à 16, dans lequel ledit agencement de génération de signal d'erreur (7, 25) est muni d'un moyen (25) pour soustraire des contributions de commandes de contrôleur de pilotage ou de vol (3, 4) à des gouvernes de l'aéronef (2) qui sont incluses dans les charges structurelles.
     
    18. Appareil selon l'une des revendications 14 à 17, dans lequel ledit agencement de génération de signal d'erreur (7,25) est prévu pour soustraire l'influence de commandes de contrôleur de pilotage ou de vol (3, 4) sur la base d'un modèle de corps rigide (16) de l'aéronef (2).
     
    19. Appareil selon l'une des revendications 14 à 18, dans lequel ledit agencement de génération de signal d'excitation (5) est prévu pour générer le signal d'indication d'excitation à partir d'un modèle de corps souple (16) de l'aéronef (2), ou d'un filtre d'observateur/de Kalman.
     
    20. Appareil selon l'une des revendications 14 à 19, dans lequel ledit circuit d'optimisation comprend un séparateur de fréquences pour optimiser la règle de pré-commande par une séparation de fréquences pour soustraire l'influence de commandes de contrôleur de pilotage ou de vol (3, 4).
     
    21. Appareil selon l'une des revendications 14 à 20, dans lequel ledit circuit d'optimisation (8, 9) est prévu pour optimiser la règle de pré-commande par un algorithme itératif.
     
    22. Appareil selon l'une des revendications 14 à 21, dans lequel des éléments de commande (20) qui sont actionnés de manière à minimiser les charges structurelles dynamiques comprennent un ou plusieurs parmi une gouverne de profondeur, une gouverne de direction, un aileron ou d'autres gouvernes de l'aéronef (2).
     
    23. Appareil selon l'une des revendications 14 à 22, dans lequel des éléments de commande (20) qui sont actionnés de manière à minimiser les charges structurelles dynamiques comprennent un ou plusieurs parmi des actionneurs électromécaniques, électromagnétiques, hydrauliques, pneumatiques ou piézoélectriques pour appliquer des forces d'amortissement de charge directement à la structure d'aéronef.
     
    24. Appareil selon l'une des revendications 14 à 23, dans lequel ledit circuit d'optimisation (8, 9) est prévu pour générer une fonction de transfert de l'aéronef pour optimiser ladite règle de pré-commande.
     
    25. Appareil selon la revendication 24, dans lequel le circuit d'optimisation (8, 9) est prévu pour générer la fonction de transfert par une identification de système en ligne (9) qui comprend de filtrer un signal de bruit aléatoire ou de fluctuation à bande limitée par un nombre de n filtres adaptatifs, où n correspond au nombre des fonctions de transfert d'aéronef à identifier, et pour soustraire la réponse de filtrage sous forme de vecteur du signal d'erreur.
     
    26. Appareil selon l'une des revendications 14 à 25, dans lequel des informations sur une commande de système commande de pilotage ou de vol (3, 4) sont ajoutées au signal d'indication d'excitation (x) afin de réduire les charges structurelles dynamiques causées par une telle commande de système de commande de pilotage ou de vol.
     




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    REFERENCES CITED IN THE DESCRIPTION



    This list of references cited by the applicant is for the reader's convenience only. It does not form part of the European patent document. Even though great care has been taken in compiling the references, errors or omissions cannot be excluded and the EPO disclaims all liability in this regard.

    Patent documents cited in the description