DEVICE, ARRANGEMENT AND METHOD FOR CONTROLLING AND REGULATING AN AIRCRAFT ACTUATING SYSTEM

Description

[0001] The invention relates to a device, an arrangement and a method for controlling and regulating an actuation system of an aircraft. background

[0002] Examples of control systems in an aircraft include tail units such as the horizontal stabilizer, vertical stabilizer and ailerons, with actuators assigned to the control surfaces for moving the control surfaces, as well as nozzles, propellers, lift aids such as flaps, spoilers and lateral force controls.

[0003] Cascade control is traditionally used in aircraft control systems. In an outer control loop, the flight control, a reference variable for controlling an actuator that controls the respective degree of freedom is determined from the aircraft's state variables relating to a degree of freedom to be controlled, serving as the flight control manipulated variable. For example, in the case of pitch control, the reference variable for actuator control is an actuator deflection and thus a deflection of the elevator. The elevator is usually moved by an actuator, for example a translatory, hydraulic, or rotary electromagnetic actuator, so that the reference variable corresponds to a desired rotational position of the actuator. The actuator control, with the reference variable as the input variable, forms an inner control loop.In the case of a rotary electromagnetic actuator for moving a control surface of a tail unit, this is particularly a so-called servo control, in which the actuator's rotational speed (actuating speed) serves as the controlled variable of an inner control loop, with a corresponding reference variable (setpoint rotational speed) of the inner control loop being determined from the desired rotational position. An actuator current for operating the actuator is determined from the control deviation of the inner control loop, i.e., the difference between the desired rotational speed and the actual rotational speed of the actuator. Additionally, a further inner control loop can be provided, in which the actuator current is the controlled variable and the manipulated variable is an actuator voltage. Servo control can take into account the actuator's actuating acceleration.A measurement of accelerations can be used in flight control for observation or as a substitute for conditions that are difficult to measure.

[0004] Servo control generally exhibits significantly faster dynamic responses than flight control, meaning that flight control limits the overall dynamics of the control system. However, given the rapidly increasing prevalence and importance of unmanned and / or autonomously operating aircraft, new areas of application for flight control technologies are emerging that require highly dynamic control of flight dynamics. In addition to precise tracking on complex trajectories, many measurement and observation tasks require increased attitude stability and smooth flight. Furthermore, there is a trend in aircraft development toward more efficient, aerodynamically superior configurations characterized by high wing aspect ratios and wingspans.Particularly in connection with ever larger, lighter and thus more elastic structures and composites, which require a control-technical reduction of structural loads and an active stabilization of the comparatively low-frequency structural dynamic modes, the natural frequencies can exceed the dynamics of the known control systems.

[0005] Document DE 10 2016 117 634 A1 proposes switching to a force- or torque-controlled approach, in contrast to the known position control (actuator deflection as the reference variable). For this purpose, a force / torque controller is provided for each actuator, with which the actuator is controlled based on the assigned reference variable, namely a target force, a target force change, a target torque, or a target torque change, and a controlled variable, namely a force generated by the actuator or a torque generated by the actuator. The controlled variable is determined by a sensor device located on or in the actuator or in the drive train of the respective actuator. The controlled variable, in the case of a tail unit, the drive torque, is thus measured on the mechanical transmission path between the force generation in the actuator and the aerodynamic control surface.This drive torque causes a rotational acceleration of the control surface relative to the aircraft. At the same time, the control surface is subjected to an aerodynamic rudder hinge moment due to the air forces. In the quasi-steady state, there is an equilibrium between the drive torque and the aerodynamic rudder hinge moment, so that the command variable in this case corresponds to the rudder hinge moment. As a result, the control surfaces react "compliantly" or "elastically" to gust loads applied externally to the control surface due to the force control. According to the approach proposed in document DE 10 2016 117 634 A1, feedback proportional to the control surface deflection is dispensed with.

[0006] This limits the design of the system dynamics compared to a complete state vector feedback.

[0007] Document DE 10 2016 117 638 A1 discloses a force / torque-controlled system of this type, which minimizes the impact of a gust on the aircraft by appropriately over- or undercompensating the hinge moment. The force / torque component caused by a gust is determined, and the specified target value for the control force / torque is modified depending on the gust's influence on the hinge moment of the control surface. An additional speed-proportional damping term is intended to reduce the tendency of the natural rudder angle dynamics to oscillate (rudder flutter). According to one embodiment, the setting position (actuator deflection) and setting speed can be fed back, whereby this occurs not in the form of a cascade, but as direct, mutually independent feedback.

[0008] The document DE 10 2010 007 042 A1 describes an aircraft with wings, each having a main wing and at least one control flap arranged so as to be adjustable relative thereto, and with a high-lift system with at least one adjustment flap for adjusting the lift of the wings for adjusting the lift state of the aircraft, comprising at least one arrangement of flow influencing devices for influencing the fluid flowing over the surface segment, a control device functionally connected to the actuators and the flow influencing devices for commanding the same in order to adjust the adjustment flap and / or to adjust the degree of flow influence brought about by the flow influencing devices, a flow influencing specification device connected to the control device for generating target commands for adjusting the lift state of the aircraft.The control device is designed in such a way that, based on desired commands from the flow influencing specification device, it generates commands for the actuator device and the flow influencing devices in order to adjust the adjustment flap and / or to adjust the degree of flow influence caused by the flow influencing devices in order to adjust the lift state of the aircraft.

[0009] Document DE 10 2010 026 162 A1 relates to an aircraft with wings and a system for minimizing the influence of transient flow conditions. The wings are each formed by a main wing and at least one control flap arranged to be adjustable relative thereto, an actuator for actuating the at least one control flap, and a sensor device for detecting the setting position of the control flap. The system for minimizing the influence of transient flow conditions comprises at least one arrangement of flow-influencing devices located in at least one surface segment of the main wing of each wing extending in the wingspan direction and / or at least one control flap for influencing the fluid flowing over the surface segment, which arrangements are functionally connected to the flight control device.a detection device for detecting unsteady flow conditions acting on the aircraft and a control function functionally connected to the flow-influencing devices for influencing the flow in the various segments of a wing, which is designed such that, based on the unsteady flow conditions detected by the detection device and the setting position of the control flap detected by the sensor device, it controls the flow-influencing devices in such a way that the influence of unsteady flow conditions on the aircraft is minimized. Summary

[0010] The object of the invention is to provide new technologies for the control and regulation of an aircraft control system, which in particular allow fast and precise control while taking into account various factors influencing the flight behavior.

[0011] To achieve the object, a device for controlling and regulating a positioning system of an aircraft according to independent claim 1 as well as an arrangement and a method for controlling and regulating a positioning system of an aircraft according to further claims are provided.

[0012] According to one aspect, a device for controlling and regulating an actuating system of an aircraft is provided. The device is formed with a first input interface configured to receive first input data indicating a command variable, a second input interface configured to receive second input data indicating a controlled variable, and a control output configured to output a control signal indicating a manipulated variable for an actuating system of an aircraft to be controlled by the actuating system. The command variable indicates a desired acceleration at a point on the aircraft to be controlled by the actuating system, and the controlled variable indicates an actual acceleration of the aircraft at that point.The device is designed to determine the manipulated variable taking into account the reference variable and the controlled variable, in particular from the difference between the reference variable and the controlled variable, and to output the control signal corresponding to the manipulated variable via the control output.

[0013] According to a further aspect, an arrangement for the control and regulation of an actuating system of an aircraft is provided. The arrangement is formed with an aircraft, a flight control device with an output interface, and a device according to the disclosure for the control and regulation of an actuating system of an aircraft. The aircraft has an actuating system configured to control the aircraft in at least one degree of freedom, and an acceleration sensor arranged at a point on the aircraft. The flight control device is configured to calculate the reference variable from a flight state of the aircraft, which indicates a desired acceleration at the point on the aircraft, and to send first input data indicating the reference variable to the first input interface of the device via the output interface.The acceleration sensor is configured to measure the local acceleration of the aircraft at the point and to send second input data indicating the controlled variable to the second input interface of the device, which second input data indicates the local acceleration at the point. The actuating system is configured to receive the manipulated variable from the control output of the device and to execute an actuating movement corresponding to the manipulated variable.

[0014] According to another aspect, a method for controlling and regulating an actuation system of an aircraft is provided.The method comprises the steps of providing a device for the control and regulation of an actuating system of an aircraft, generating first input data indicating a reference variable, wherein the reference variable indicates a desired acceleration at a point of the aircraft to be controlled by means of the actuating system, generating second input data indicating a controlled variable, wherein the controlled variable indicates an actual acceleration of the aircraft at the point, receiving the first input data at a first input interface of the device, receiving the second input data at a second input interface of the device, determining a manipulated variable for an actuating system of the aircraft taking into account the reference variable and the controlled variable, in particular from the difference between the reference variable and the controlled variable, and outputting a control signal indicating the manipulated variable via a control output of the device.

[0015] The control system can, for example, be a tail unit and an actuator for operating the tail unit. The actuator can be the system for generating a control force or a control torque without an associated control system. Alternatively, the actuator can have a control system, in particular a servo control. The tail unit can comprise a fixed part (lifting surface), i.e., one that is immobile relative to the aircraft, and a movable control surface that exerts a desired control effect on the aircraft depending on its position relative to the fixed part. The actuator can be a rotary electromagnetic actuator, i.e., an electric motor that rotates the control surface relative to the fixed part of the tail unit. The tail unit can be a horizontal stabilizer, a vertical stabilizer, or an aileron.Alternatively, the actuation system can be, for example, a flap for controlling the aircraft's lift, a spoiler for longitudinal control of the aircraft, or a lateral force control for lateral control of the aircraft. Other examples of the actuation system include a nozzle and a propeller of the aircraft.

[0016] An acceleration of the aircraft within the meaning of the disclosure is an acceleration of the aircraft as such and therefore not, for example, a control acceleration of an individual actuator arranged in the aircraft. In particular, an acceleration of the aircraft may exclude an acceleration of an element arranged on the aircraft that is movable relative to the aircraft, for example, a control surface of a tail unit, which includes an acceleration relative to the aircraft.

[0017] For example, the acceleration of the aircraft can be an acceleration at a center of gravity of the aircraft or an acceleration at a stationary part of a tail unit (fin), wing, or fuselage. The acceleration of the aircraft can, for example, be the acceleration at a point near a control surface, in particular the control surface controlled by the actuator. The acceleration of the aircraft can, in particular, also be an acceleration at a part of the actuator itself that is immobile relative to the aircraft structure, for example the housing, a base plate, a circuit board of the control electronics, or another immobile part of the actuator. In such an embodiment, system integration can be simplified and independent production by manufacturers of system components can be enabled.

[0018] The acceleration of the aircraft can, in particular, be an acceleration of the aircraft as such with reference to an inertial system or global coordinate system. The acceleration of the aircraft can include the influence of the Earth's gravitational field and / or a compensation for it. The acceleration of the aircraft can be a translational acceleration, a rotational acceleration, or an acceleration that contains translational and rotational components.

[0019] With the actual acceleration of the aircraft at that point, a controlled variable is provided that depends not only on the actuating system's deflection (e.g., the position of a control surface of a tail unit or the rotational or translational position of an actuator moving the control surface), but also includes other external influences, in particular the influences of gusts of wind and / or various flight state variables such as airspeed, flow angle, and yaw rates, which contribute to the generation of aerodynamic forces. This can enable the suppression of disturbing influences (gusts) in an inner, more dynamic control loop. A control system can be provided that is less sensitive to specific aerodynamic or aeroelastic properties of the aircraft (robustness).Furthermore, a simpler and more standardized structure of the flight control and a more agile and precise trajectory guidance can be made possible.

[0020] The device can be configured with a third input interface configured to receive third input data indicating a control system variable. The device is configured to determine a control system reference variable taking into account the reference variable and the controlled variable, and to determine the manipulated variable taking into account the control system reference variable and the control system variable.

[0021] The actuating system reference variable can be a target actuating speed of the actuating system, and the actuating system controlled variable can be an actual actuating speed of the actuating system. For example, the actuating system reference variable and the actuating system controlled variable can be a target rotational speed (target rotation rate) and an actual rotational speed of an actuator of the actuating system, for example, an actuator that moves a control surface of a control unit. The manipulated variable can be a variable that causes a movement of the actuator, in the case of an electromagnetic actuator, in particular an actuator current.

[0022] Alternatively, an actuator reference variable, for example, a target actuator current, can be determined for an actuator of the actuating system, taking into account the actuating system reference variable and the actuating system controlled variable. Taking into account the actuator reference variable and an actuator controlled variable, for example, an actual actuator current, the manipulated variable can be determined, which can be, in particular, an actuator voltage, for example, a terminal voltage of a DC motor or the quadrature voltage component in the case of field-oriented control of an electronically commutated motor. The device can have a corresponding input interface for receiving the actuator controlled variable. The actuator control or servo control can have further control structures, in particular below a control speed control, which are known as such.

[0023] In general, determining a manipulated variable, possibly a subordinate manipulated variable in a subordinate control structure, taking into account variables that are specified (i.e., serve as a reference variable), and variables that are fed back (i.e., serve as controlled variables), involves determining the manipulated variable according to the understanding of control engineering. In particular, cascade structures can be formed in which a control deviation, the difference between a reference variable and a controlled variable (which may be determined as a composite controlled variable from several controlled variables), is multiplied by a proportionality factor to determine the manipulated variable.Alternatively or additionally, a parallel feedback can be provided in which one or more controlled variables are fed back with a respective adaptation, for example amplification and / or integration, and are offset by addition or subtraction with a reference variable which has been modified by means of a pre-filter in accordance with the fed-back controlled variables, in particular to compensate for a static error between the reference variable and the one or more controlled variables.

[0024] The device can be configured to perform one, several, or all of the following operations to determine the manipulated variable: determining a (target) actuating speed (actuating variable or actuating system reference variable) by multiplying a difference between a target acceleration (reference variable) and an actual acceleration (controlled variable) by a first proportionality factor; determining a (target) actuator current (actuating variable or actuator reference variable) by multiplying a difference between a target actuating speed (actuating system reference variable) and an actual actuating speed (actuating system controlled variable) by a second proportionality factor; and determining a (target) actuator voltage (actuating variable) by multiplying a difference between a target actuator current (actuator reference variable) and an actual actuator current (actuator controlled variable) by a third proportionality factor.

[0025] The actuator voltage can be the control variable of the control system. The actuator voltage can be adjusted according to the control variable, and the response of the control system to this, in particular a movement of the control system and the assumption of a position of the control system, can result from this based on the physical and physical system conditions, in particular the control system in connection with the aircraft system.

[0026] Alternatively, the actuator current can be the control variable of the control system. In this case, a target actuator current corresponding to the control variable can be specified to the actuator, whereby no feedback of an actual actuator current takes place. In this case, the actuator can be configured to convert a current specification into a corresponding voltage in order to achieve the current specification. In this case, the actuator can have an internal control system which, for example, uses the current as a reference and controlled variable and the voltage as the control variable. The fact that the actuator current is the control variable of the control system can be provided in particular if the actuator implements a current specification with sufficient dynamics so that sufficiently dynamic control of the actuating system is possible in accordance with the respective application, in particular the type of aircraft, without current control being carried out by the device.

[0027] The device can be configured for operation without taking into account a positioning acceleration of the positioning system, in particular without taking into account a positioning acceleration of an actuator, for example without taking into account a rotational or angular acceleration of a rotary electromagnetic actuator for moving a positioning surface of a tail unit. In particular, the determination of the manipulated variable can be carried out without taking into account a positioning acceleration. While the positioning acceleration is proportional to the positioning torque, in particular the drive torque of an actuator, the acceleration of the aircraft at the point can be proportional to a positioning position of the positioning system. For example, a local acceleration on the immobile part of a tail unit, i.e. on the lifting surface, can be proportional to the positioning surface angle and a lift force generated by it.Within the controlled system, there can be two integration levels between a control acceleration and the acceleration of the aircraft. By feedbacking the aircraft's acceleration, i.e., using the aircraft's acceleration as a controlled variable, it can be possible to influence the system dynamics in a similar way to feedbacking a control position, for example, a control surface angle.

[0028] The device can be configured to determine the manipulated variable independently of an actual actuating position of the actuating system and independently of determining a desired actuating position of the actuating system. In this case, an actuating position of the actuating system can be, in particular, an actuator actuating position, for example, a rotational position or a translational position of an actuator for moving the actuating surface of a control unit, or the position of the actuating surface of the control unit corresponding to this position.

[0029] Alternatively, it can be provided that the device is configured to determine the manipulated variable independently of determining a desired actuating position of the actuating system, but to take into account an actual actuating position of the actuating system.

[0030] For example, the actual position can be taken into account when limiting the movement range of the actuating system. In this case, the manipulated variable can be modified if a test shows that the manipulated variable would lead to a positioning movement beyond a predetermined movement range of the actuating system, such that the positioning movement ends at the limit of the movement range. In this way, in particular, a positioning movement can be provided which corresponds to the function of a shutdown when a limit switch is reached. In this case, the manipulated variable can also be determined independently of taking into account the actual position of the actuating system and independently of determining a target position of the actuating system.

[0031] Alternatively or additionally, the actual setting position can be fed back, for example, to observe conditions of the unsteady aerodynamics or elasticities and hysteresis in the drive train in order to enable a further increase in the dynamics of the control, whereby the control variable is determined independently of determining a target setting position of the control system.

[0032] The device can have a further input interface configured to receive further input data indicating a further controlled variable. The further controlled variable can indicate an actual acceleration of the aircraft at a further point, and the device can be configured to adapt the controlled variable taking into account the further controlled variable and subsequently determine the manipulated variable taking into account the reference variable and the controlled variable. Adapting the controlled variable using the further controlled variable can, for example, comprise adding the further controlled variable to the controlled variable. Additional controlled variables can be received and used accordingly to adapt the controlled variable.

[0033] This can make it possible to determine the manipulated variable based on a controlled variable which indicates an acceleration at a point on the aircraft at which no acceleration measurement is taken, wherein the acceleration is determined from the accelerations at at least two other points on the aircraft at which an acceleration measurement is taken. This can provide a virtual acceleration measurement at a measuring point which is different from the point and the further point on the aircraft. In particular, it can be made possible to determine accelerations at different points on the aircraft and / or in different degrees of freedom using a limited number of acceleration sensors at different points on an aircraft, which may at least partially already be arranged on or in the aircraft for other purposes.These can be used for the control and regulation of control systems for different degrees of freedom of the aircraft, for example, the horizontal stabilizer, vertical stabilizer, and ailerons, as a controlled variable in several devices according to the disclosure, which are assigned to a respective control system. For an aircraft considered to be rigid, six acceleration measurements can be provided at at least three different points and in at least three different directions to determine accelerations at any point on the aircraft.

[0034] In general, the acceleration specified by the controlled variable can be determined from multiple measurements. The individual measurements of an acceleration can determine different measured variables to determine the acceleration according to the controlled variable. For example, a roll acceleration, particularly around the aircraft's center of gravity, can be determined from two vertical movements of the wings, with the vertical movements being measured using appropriate sensors.

[0035] Components of the device may be provided as separate devices. Alternatively, some or all components of the device may be provided as virtual components of a physical component. Components of the device configured as separate physical devices may be mounted contiguously or configured separately from one another.

[0036] The reference variable can be determined in an upstream control process, particularly in a flight control system. In this case, the upstream control process can operate at a significantly lower clock rate than the downstream control system (servo control).

[0037] With respect to the arrangement, the flight control device can be configured to calculate the reference variable taking into account a control variable determined directly kinematically from a desired trajectory of the aircraft.

[0038] It may be provided that the reference variable is calculated based on the actual flight condition of the aircraft and the desired flight condition, in particular based on a deviation or difference between the actual flight condition and the desired flight condition. Preferably, the reference variable can additionally include a feedforward control, which is calculated directly from the desired flight condition and independently of the actual flight condition.

[0039] A flight condition can be defined by one or more physical quantities or measured values ​​of these quantities that fully or partially characterize the dynamic behavior of the aircraft, or that allow the determination of such characterizing quantities (e.g., with the help of an observer). A flight condition can also include quantities that are referred to as output variables in the context of control engineering.

[0040] A desired flight state can, for example, be given in the form of temporal profiles or constant values ​​for the physical variables used to describe the flight state. Alternatively or additionally, a desired flight state can be given in the form of a desired trajectory of the aircraft. In this case, a feedforward control contained in the reference variable can be determined particularly easily by determining the acceleration at a location on the aircraft using known kinematic relationships from the desired trajectory.

[0041] A target trajectory of the aircraft can, for example, be a line that describes the desired position of the aircraft's center of gravity in the plane corresponding to a flight altitude or in three-dimensional space. It can also contain a temporal assignment of the positions, i.e. describe a line in a four-dimensional space. Furthermore, the target trajectory can describe a desired temporal or spatial dependence of the aircraft attitude, which can be represented, for example, by one or more angles, rotation matrices, or quaternions. The target trajectory can refer to any desired, preferably at least approximately inertial, coordinate system. For example, the coordinate system can be an earth-fixed coordinate system or a coordinate system that moves with the air mass surrounding the aircraft. The angles can be absolute angles or angles that refer to the direction of the flight path.

[0042] The control system can be formed with an actuator that moves a control surface of an aircraft's tail unit. The tail unit can comprise a fixed part (lifting surface), i.e., one that is immobile relative to the aircraft, and a movable control surface that exerts a desired control effect on the aircraft depending on its position relative to the fixed part. The actuator can be a rotary or translatory electromagnetic actuator, for example, an electric motor that rotates the control surface relative to the fixed part of the tail unit. Alternatively, another type of actuator can be provided, for example, a hydraulic or electro-hydraulic actuator, in particular with one or more hydraulic cylinders. The tail unit can be a horizontal stabilizer, a vertical stabilizer, or an aileron.In a rotary actuator, for embodiments in which the device is configured to determine the difference between a control system command variable and a control system controlled variable, the control system command variable may be a desired rotational speed (rotation rate) and the control system controlled variable may be an actual rotational speed.

[0043] The acceleration sensor can be arranged, in particular, on a part of the tail unit that is immobile relative to the aircraft. Thus, the arrangement is configured to provide control based on local acceleration at the tail unit of the aircraft as a whole system, wherein an acceleration of the footprint of the tail unit relative to the aircraft, in particular relative to the immobile part of the tail unit, is not detected and thus does not influence the control. A immobile part of a tail unit can be, for example, the fin of a horizontal or vertical stabilizer or, in the case of ailerons, flaps, or spoilers, the wing.A stationary part of the tail unit can also be part of the control surface itself, provided that the acceleration measured there essentially reflects the acceleration of the aircraft as a whole, i.e., with sufficient approximation for control, and the relative acceleration caused by the control movement itself has only a minor influence. For example, this could be an acceleration measurement on or near the rudder hinge axis or an actuator installation location. This can be particularly useful for a pendulum rudder, where there is no separation between the fin and flap, but the entire tail unit is adjusted.

[0044] Alternatively, the arrangement can be formed with a different type of actuating system. For example, the actuating system can be a nozzle of the aircraft, a propeller of the aircraft, a flap, a spoiler, a lateral force control, or another actuating system of the aircraft, wherein the actuating system is in any case controllable by means of the device and configured to act in at least one degree of freedom of the aircraft to control the aircraft.

[0045] The arrangement can be formed with a further acceleration sensor arranged at a further point of the aircraft. In this case, the device is a device with a further input interface, which is configured to receive further input data indicating a further controlled variable. The further acceleration sensor is configured to measure the local acceleration at the further point and to send further input data indicating the further controlled variable to the further input interface of the device, which input data indicates the local acceleration at the further point. By means of the acceleration sensor and the further acceleration sensor, a virtual acceleration measurement can be provided at a measuring point that is different from the point and the further point of the aircraft.In this case, the configurations explained above with reference to the device can be provided correspondingly with respect to the virtual acceleration measurement. The arrangement can comprise several additional acceleration sensors, wherein the device has a corresponding number of input interfaces, and by means of the acceleration sensors, one or more virtual acceleration measurements can be provided at one or more points, each of which is different from the points at which the acceleration sensors are arranged.

[0046] The arrangement can additionally comprise further sensors that are not acceleration sensors. Alternatively or additionally, an acceleration sensor can be formed with a plurality of sensors, each of which measures a quantity other than acceleration, wherein the acceleration sensor determines the acceleration from the quantities of the sensors. For example, an acceleration sensor can be formed with at least two sensors for detecting a vertical movement on the wings of the aircraft, wherein the acceleration sensor determines a roll acceleration of the aircraft from the vertical movement of the wings determined by means of the at least two sensors. An acceleration sensor can be formed by one or more force, pressure, strain, movement, or position sensors, provided that their arrangement and the processing of their measured values ​​is configured to determine a local acceleration at a location on the aircraft.

[0047] The arrangement can comprise a further device according to the disclosure, wherein the aircraft has a further control system configured to control the aircraft in the at least one degree of freedom or in at least one further degree of freedom, and an additional acceleration sensor arranged at an additional point of the aircraft. The flight control device can be configured to additionally send the first input data indicating the reference variable to the further device via the output interface. The additional acceleration sensor can be configured to measure the local acceleration of the aircraft at the additional point and to send second input data indicating an additional controlled variable to the further device, which second input data indicates the local acceleration at the additional point.Furthermore, the further actuating system can be configured to receive the actuating variable from the control output of the further device and to execute an actuating movement corresponding to this actuating variable.

[0048] In a corresponding embodiment, additional devices and additional control systems can be provided, wherein the above-described configurations regarding the control system can be provided accordingly for the additional control systems. In this way, control can be provided for several or all degrees of freedom of movement of the aircraft. Several devices according to the disclosure can be provided as virtual devices in a physical device.

[0049] The command variable provided by the flight control device can specify a desired acceleration of the aircraft in several degrees of freedom. The aircraft's degrees of freedom can include, for example, positions in three spatial directions and three attitude angles. In an elastic aircraft, degrees of freedom can also include variables characterizing the deformation state, such as modal amplitudes. Alternatively or additionally, the aircraft's degrees of freedom can be determined by the position of various points on the aircraft in space.

[0050] As an alternative to providing the same (in particular vectorial) command variable in several devices according to the disclosure, it can be provided that the flight control device provides a respective command variable for each of the devices, wherein the respective command variable indicates a desired acceleration of the aircraft in a degree of freedom that corresponds to the degree of freedom that is primarily, predominantly or exclusively influenced by means of the actuating system assigned to the respective device.

[0051] In embodiments with more than one device according to the disclosure, instead of a respective acceleration sensor assigned to one of the devices, a single acceleration sensor or a single system of multiple acceleration sensors can be provided, which determine an acceleration of the aircraft in multiple degrees of freedom and / or at multiple points on the aircraft, possibly as a virtual acceleration measurement, and provide corresponding control variables for the devices according to the disclosure. The measured values ​​of one or more acceleration sensors can be provided for multiple devices according to the disclosure. In this case, the number of acceleration sensors, reference variables, and devices according to the disclosure provided need not be identical.In particularly advantageous embodiments, however, this may be the case, which may enable decoupling of different degrees of freedom and any specification of the system dynamics.

[0052] The aircraft can be highly flexible. Complete state feedback can be provided, for example, by measuring accelerations with multiple acceleration sensors at locations distributed throughout the aircraft, or by separating rigid-body motion from structural dynamics, with a division into rigid-body degrees of freedom and amplitudes of the elastic methods. The equations of motion of rigid-body motion and structural dynamics are inertially decoupled, but coupled via external forces (aerodynamics). In the case of complete state feedback, an innermost control loop with feedback of a locally measured acceleration can be provided. This can enable complete control of all eigenmodes, particularly if a number of acceleration points corresponds to a number of considered degrees of freedom.Outer control loops can provide separate control of rigid-body motion and structural dynamics. For this purpose, the control of rigid-body motion and structural dynamics can be provided in a cascade configuration. Setpoints for accelerations of the rigid-body degrees of freedom and the modal degrees of freedom can be converted into setpoints for the local accelerations, for example, using eigenvectors and kinematic translations. The outer control loops can refer to generalized coordinates, while inner control loops refer to the local degrees of freedom. The system behavior can be independent of the description form, whereby a transformation between different degree-of-freedom systems and state representations can be enabled.

[0053] Alternatively, for a design with a highly flexible or elastic aircraft, local acceleration measurement can be provided on the actuating system, in the case of a tail unit in particular directly on the actuating surface on an immovable part of the tail unit, whereby the local acceleration is used exclusively for the control of the actuating system on which the measurement is taken. The number of actuating and measuring positions can be suitably selected and correspond to the number of degrees of freedom taken into account in order to enable system dynamics to be freely specified in this case as well. In an elastic aircraft, relative movements caused by the elastic deformation can occur between a part of the tail unit that is immovable relative to the aircraft and other parts of the aircraft, for example part of the fuselage or the aircraft's center of gravity.

[0054] In general, the aircraft can be any type of aircraft, for example, a slightly flexible or elastic, moderately flexible or elastic, or highly flexible or elastic aircraft. A highly flexible (elastic) aircraft can, in particular, be an aircraft that can no longer be described with sufficient accuracy using a linear approach. In a flexible aircraft, the acceleration of the aircraft can, in particular, be an acceleration of the elastic aircraft structure at a point where aeroelastic vibration modes (eigenmodes) exhibit an extremum or a node.

[0055] The method for controlling and regulating an actuating system of an aircraft may comprise receiving third input data indicating an actuating system controlled variable at a third input interface of the device, wherein determining the manipulated variable taking into account the reference variable and the controlled variable comprises determining an actuating system reference variable taking into account the reference variable and the controlled variable, and determining the manipulated variable taking into account the actuating system reference variable and the actuating system controlled variable.

[0056] The configurations explained above with regard to the device for controlling and regulating an aircraft control system can be provided accordingly in connection with the device and / or the method, and vice versa. In particular, the device can be configured to function with components explained in connection with the arrangement, and the arrangement can have components configured to provide the explained functions in interaction with the device.

[0057] According to the disclosure, a device for controlling and regulating an actuating system of a vehicle is provided, which, unless the vehicle is an aircraft, is not within the scope of protection and which is formed with a first input interface configured to receive first input data indicating a reference variable, a second input interface configured to receive second input data indicating a controlled variable, and a control output configured to output a control signal indicating a manipulated variable for an actuating system of a vehicle to be controlled by the actuating system. The reference variable indicates a desired acceleration at a point on the vehicle to be controlled by the actuating system, and the controlled variable indicates an actual acceleration of the vehicle at that point.The device is configured to determine the manipulated variable from the difference between the reference variable and the controlled variable and to output the control signal corresponding to the manipulated variable via the control output. In connection with the device for the control and regulation of an actuating system of a vehicle, the embodiments provided with reference to the device for the control and regulation of an actuating system of an aircraft can be provided accordingly. In particular, according to the disclosure, a corresponding arrangement is provided for the control and regulation of an actuating system of a vehicle, which arrangement is not within the scope of protection insofar as it is not related to an aircraft, comprising a vehicle, an actuating system and an acceleration sensor, a vehicle control device with an output interface, and a device for the control and regulation of an actuating system of a vehicle.Furthermore, according to the disclosure, a corresponding method for controlling and regulating an actuating system is provided. The vehicle can be, in particular, an aircraft, according to the invention, an airplane, or, in examples not falling within the scope of protection, a helicopter or an airship. Alternatively, in examples not falling within the scope of protection, the vehicle can be, for example, a watercraft or a spacecraft. Description of implementation examples

[0058] Further embodiments are explained in more detail below with reference to the figures of a drawing. Herein: Fig. 1 shows a known arrangement for the control and regulation of an aircraft actuating system; Fig. 2 shows an arrangement according to the disclosure for the control and regulation of an aircraft actuating system; Fig. 3 shows an arrangement of an acceleration sensor on a horizontal stabilizer of an aircraft; Fig. 4 shows a further arrangement for the control and regulation of an aircraft actuating system; Fig. 5 shows a schematic representation of a concept for acceleration-based roll attitude control of an aircraft; Fig. 6 shows a schematic representation of a concept for acceleration-based control of mechanical systems; Fig. 7 shows a schematic representation of a concept for acceleration-based control of elastic aircraft; Fig. 8A-E shows the overall system dynamics for a known and a disclosed control of an aircraft actuating system; Fig.Fig. 9 shows a Bode diagram for a previously known and a disclosed embodiment of a control system for an aircraft actuator; Fig. 10 shows a schematic representation of an arrangement on a flexible aircraft; and Fig. 11 shows a schematic representation of an alternative arrangement on a flexible aircraft.

[0059] The Fig. 1shows an arrangement for the control and regulation of an actuating system of an aircraft according to a known approach. A flight control device 1 of the aircraft, which is an automatic control system that can also be referred to as a flight control system, receives measured variables 2 to describe the state of motion of the aircraft. Based on the state of motion of the aircraft, the flight control device 1 determines a reference variable 3 for the control of the actuating system 4. The actuating system 4 is formed with an actuating element 4a and a force generator 4b, whereby the actuating system can comprise further components. The force generator 4b can be, in particular, a control surface of a tail unit and, if appropriate, fixed components of the actuating system that participate in the force generation. Alternatively, the force generator 4b can be, for example, a nozzle or a propeller.The actuator 4a serves to influence the force generator 4b such that a desired force is exerted on the controlled system, i.e., the aircraft. In embodiments in which the actuating system 4 is a tail unit, the actuator 4a can be a servomotor that pivots a control surface of the tail unit as a force generator 4b or part of the force generator relative to a stationary part of the tail unit. In alternative embodiments, the actuator 4a can be, for example, a valve of a nozzle serving as a force generator 4a or a drive motor of a propeller.

[0060] The reference variable 3 indicates a desired setting position of the actuator 4a of the setting system 4, for example a rotation position of a servo motor (corresponding to a setting surface position), an opening state of a valve of a nozzle or a drive position of a drive motor of a propeller, from which a propeller speed results, or of a servo motor for blade angle adjustment.

[0061] A control device 5 of the arrangement receives the reference variable 3 via a corresponding input interface. Via a corresponding further input interface, the control device also receives a controlled variable 6, which indicates the actual actuating position of the actuating system. The control device determines the control deviation as the difference between the setpoint of the actuating position according to the reference variable 3 and the actual value of the actuating position according to the controlled variable 6. By multiplying this by a proportionality factor, the control device 5 determines a control system reference variable from the control deviation, which is compared with a control system controlled variable 7 to determine a control variable 8 of the actuating system. For example, the control variable 8 can be an actuator voltage or an actuator current.

[0062] Based on the manipulated variable 8, the actuator 4a causes a position of the force generator 4b. This results in a force and / or torque effect 9 on the mechanical system 10 of the aircraft. While the aircraft as a mechanical system 10 is in the Figures 1 and 2 is shown separately from the other components, the control system 4 and, in advantageous embodiments, also the flight control device 1 and the control device 5 form part of the aircraft.

[0063] In addition to the desired force and / or torque effect 9, disturbing forces and / or moments 11 act on the mechanical system 10 of the aircraft, which are caused by external influences, for example wind effects, especially in the form of gusts of wind. As can be seen from the illustration of the Fig. 1As can be seen, the disturbing forces and / or moments 11 acting on the aircraft can only be compensated by the flight control device 1 if measured variables 2 are taken into account, which include the influence of the disturbing forces and / or moments 11. Thus, such consideration takes place exclusively within the framework of the flight control, which is usually slow compared to the control of the positioning system (servo control).

[0064] The Fig. 2 now shows a disclosed arrangement for the control and regulation of an aircraft control system. In comparison to the arrangement according to Fig. 1According to the disclosure, a device 12 is provided for the control and regulation of the positioning system 4, which is configured to receive a reference variable 13 at a first input interface, which indicates a desired acceleration at a point on the aircraft. For this purpose, the flight control device 1 is configured to determine the reference variable 13 indicating the desired acceleration from the measured variables 2 and send it to the device 12. At a second input interface, the device 12 receives a controlled variable 14, which indicates the actual acceleration at that point on the aircraft.

[0065] In particular, the acceleration of the aircraft can be a local acceleration at the positioning system 4. The Fig. 3shows, by way of example, the arrangement of an acceleration sensor 15 on the stationary part of a horizontal stabilizer 16 of an aircraft. The horizontal stabilizer 16 is a control system 4 of the aircraft, in which a force is exerted on the aircraft by a movement of the control surface of the stabilizer, which functions as a force generator 4a, by means of a servo motor as the control element 4a. This force causes the aircraft to pitch and thus subsequently climb or descend.

[0066] Alternatively, the acceleration can be an acceleration at another point on the aircraft, for example, at the aircraft's center of gravity. The acceleration can be measured directly with an accelerometer or determined from one or more measurements, which can be accelerations at one or more other points or quantities other than accelerations, for example, vertical movements (changes in position) of wings.

[0067] The control device 12 determines the control deviation as the difference between the desired acceleration value according to the reference variable 13 and the actual acceleration value according to the controlled variable 14. By multiplying the control deviation by a proportionality factor, the control device 12 determines a control system reference variable, which is compared with a control system controlled variable 7 to determine a control variable 8 of the control system. Based on the control variable 8, the actuator 4a causes a position of the force generator 4b, which leads to a force effect 9 on the mechanical system 10 of the aircraft.

[0068] In an alternative embodiment, instead of such a cascade structure, a parallel feedback can be provided, in which controlled variables 7 and 14 are fed back and modified, in particular multiplied by a gain factor and / or integrated. In this case, reference variable 13 is modified by a prefilter according to controlled variables 7 and 14, and manipulated variable 8 is then determined by adding controlled variables 7 and 14 and reference variable 13.

[0069] As in the Fig. 2 As can be seen, the controlled variable 14 is a variable of the mechanical system 10 of the aircraft. The disturbing forces and / or moments 11 influence the acceleration of the aircraft, so that the acceleration of the aircraft specified by the controlled variable 14 already contains these influences, at least partially. Thus, in the case of the Fig. 2 presented control concept compared to that in the Fig. 1The known concept illustrated already takes into account disturbing forces and / or moments 11 acting on the aircraft when controlling the aircraft's control system 4.

[0070] The manipulated variable 8 can, in particular, be an actuator voltage or an actuator current. For example, the actuating system reference variable can be a setpoint value of an actuating speed of the actuating element 4a, i.e., in particular, an actuator. The actuating system controlled variable 7 can, in this case, be an actual actuating speed of the actuating element 4a. In an exemplary embodiment, a setpoint value of an actuator current is determined from the difference between the actuating system reference variable and the actuating system controlled variable. The setpoint value of the actuator current can be the manipulated variable 8. Alternatively, a further inner control loop can be provided in which the manipulated variable 8 is determined using the setpoint value of the actuator current.

[0071] The Fig. 4shows such a design of an arrangement for the control and regulation of an aircraft actuating system, in which an additional inner control loop is provided. This shows, as an example, a pitch attitude control system with a rotary, electromagnetic actuator of a horizontal stabilizer. In contrast to a known control system with feedback of a control surface deflection n, a local acceleration b zH is fed back to the horizontal stabilizer, and a control deviation from a predetermined acceleration b zH,c is determined within the actuating system control 17. From this, a proportional actuating speed command ṅ c is determined using the factor K bzH, which is set by the inner speed control loop.The commanded current flow I c (actuator reference variable) is proportional (factor K ṅ ) to the speed error ṅ c -ṅ, the difference between the actuating speed command ṅ c (actuating system reference variable) and the actual actuating speed ṅ, which represents the actuating system controlled variable. Finally, the motor terminal voltage U forms the manipulated variable. It is set proportional (factor KI ) to the control error I c -I of the current control cascade. The actual current I represents the actuator controlled variable.

[0072] Within the framework of the physical processes within the actuator, according to a model as a DC shunt machine, the terminal voltage causes a change in the current flow in the motor windings, which is inversely proportional to their inductance L. However, the voltage drop ΔU res = R·I due to the winding resistance R, as well as the counter voltage ΔU emf = K e ·ṅ induced by the rotary movement, which is proportional to the motor constant K e, must be taken into account; these reduce the terminal voltage. The current flow I is calculated by integrating the current change and causes a drive torque M act that is proportional to the motor constant K t.

[0073] Regarding the physical effect on the control system, in addition to the drive torque M act , the aerodynamic rudder hinge moment M aero acts on the control surface. This includes both components proportional to the deflection n (factor C n,aero ) and damping components (factor C ṅ,aero ). Furthermore, the aerodynamic rudder hinge moment M aero is influenced by the flow direction (factor C α,aero ). The resulting total moment leads to a control acceleration n̈ that scales with the inverse 1 / J of the rotational inertia.

[0074] In the right part of the Fig. 4the dynamics of the aircraft pitching motion 18 is shown in simplified form. The pitching acceleration q with the inverse pitch inertia 1 / I yy is proportional to the pitching moment, which results from the pitching moment coefficient by denormalization with dynamic pressure q, wing area S and wing chord I µ. This coefficient essentially contains influences from the elevator (C mn n), the pitch rate (C mq I µ 1 / VA q̇) and the angle of attack (C mα ​​α). The angle of attack α is determined not only by the component of the longitudinal attitude Θ but also by the influence γ of the track motion 19. It also contains the essential part of the disturbing influence (gusts) in the form of the wind angle of attack α W . The local acceleration b zH at the horizontal stabilizer results from the pitching acceleration q with the lever arm r H and from the vertical acceleration bz of the aircraft's center of gravity.

[0075] According to the revelation, as for example from the Fig. 4As can be seen, the actuating position is not used as the controlled variable. Also, no force or torque measurement serves as the controlled variable. Furthermore, the controlled variable is not measured in the actuator drive train or on the actuating surface, but in the design of the Fig. 4 on the tail unit assigned to the landing surface (the lifting surface that is immobile relative to the aircraft). There is no measurement of the (rotational) acceleration n̈ of the actuator, which would be proportional to the actuating torque (drive torque of the actuator, M act ). The local acceleration at the lifting surface, on the other hand, is proportional to the landing surface angle and the lift force generated by it, i.e., to a quantity that is separated from the acceleration n̈ of the actuator by two integration stages, as in the Fig. 4The local acceleration is an output quantity that depends to a large extent on the angle n of the system and whose feedback can therefore influence the system dynamics in a similar way. According to the Fig. 4 The design shown is intended to represent the speed control loop of a classic servo control (middle cascade in the Fig. 4) are maintained so that a number of linearly independent output variables corresponding to the system order is still fed back. This can make it possible to design the system dynamics as arbitrarily as desired. In this case, a design of the rudder angle dynamics that is as "stiff" as possible may be desired. In particular, the aim may not be to reduce the actuator load or the control effort, but rather to move the control surface as quickly as possible to the position that compensates for the gust influence on the corresponding tail unit. This position is generally not identical to the rest position to which the free rudder would blow out if the control torque were kept constant.

[0076] Local acceleration control can offer advantages over control of the rudder hinge moment. Local acceleration measurement (unlike the control surface angle or the control surface hinge moment) directly records the additional lift caused by the gust via the additional angle of attack α w . In elastic aircraft, local acceleration directly reflects the structural dynamic vibration state. The feedback of acceleration to the setting speed of a control surface acting at the same location corresponds to virtual damping (similar to the so-called ILAF principle). It is therefore particularly suitable for the active stabilization of highly elastic configurations. Furthermore, local acceleration includes influences from various flight state variables (Θ, γ, q, cf. Fig. 4), which can also be compensated for by the control system. Compared to the highly dynamic feedback path via K bz,H, these influences can be less significant, resulting in considerably greater robustness against changing aerodynamic properties. The local acceleration can result in a direct, purely kinematic relationship from a planned path and attitude trajectory. It can therefore be possible to derive simple feedforward control laws that are independent of the properties of a specific aircraft. The high dynamics of the local acceleration control (which corresponds to the classic position control loop of the servo control) can thus be used not only for disturbance suppression, but also for guidance behavior. This can enable considerably more agile path guidance.

[0077] Actuator control (servo control) and flight state control (flight control) traditionally represent separate research disciplines that are addressed by different expert groups. The feedback of a local acceleration measured on the aircraft structure in an inner control loop, which is traditionally part of servo control, bridges the gap between the two areas. This requires a holistic view of the entire control system, which considers the aircraft and its actuators as a single unit. By using local acceleration as a control input, components of the flight dynamics can be incorporated into the servo control system. This can make it possible to simplify the flight control system and reduce dependencies on specific flight characteristics, thus rendering traditional flight control structures inapplicable.

[0078] The limit of the positioning system control 17 is defined according to Figure 4 at the local acceleration b zH and the positioning surface deflection n. Other representations are possible in which the definitions, especially the limits, of subsystems are specified differently (see, for example, Fig. 5 ), without this resulting in a change in the regulatory principle as disclosed.

[0079] The Figures 5 , 6 and 7 The symbols used indicate the following quantities: Scalars: C Iβ : Side roll moment C Iξ : Aileron effectiveness C Ip : Roll damping I: Actuator current I yy : Roll moment of inertia of the aircraft J: Moment of inertia of the actuator K t : Torque constant of the actuator K...: Controller gain of the ...-control loop S: Wing area VA : Airspeed q: Dynamic pressure b: Half span p: Roll rate β: Side shift angle β W : Wind shift angle ω: Angular velocity of the actuator ξ: Aileron deflection Vectors: n : modal amplitudes (structural dynamic degrees of freedom) R : Position vector to the local acceleration measurement point q : generalized coordinates u : Control variables x : Rigid body degrees of freedom z : Disturbance matrices and tensors: B η : control influence on generalized forces of the structural dynamics degrees of freedom B x : Control influence on generalized forces of the rigid body degrees of freedom B : Position influence on generalized forces C : generalized stiffness matrix D : generalized damping matrix E η : Perturbation influence on generalized forces of the structural dynamics degrees of freedom E x: Perturbation influence on generalized forces of the rigid body degrees of freedom E : Disturbing influence on generalized forces F η ext< , F η˙ ext< : Influence of structural deformation-induced aerodynamic forces on rigid body motion K... : Gain matrix of the ... control loop L : Kinematic transmission ratios between generalized rigid body degrees of freedom and position of the local acceleration measurement points M: generalized inertia matrix Q n , Q η̇ : Influence of structural deformation-induced aerodynamic forces on structural dynamics Q x , Q ẋ : Influence of rigid body motion dependent aerodynamic forces on structural dynamics Λ : Eigenforms (eigenvectors) of structural dynamics β : generalized structural damping factors γ : generalized stiffness matrix µ : modal mass matrix indices: c: command size, default value, target value

[0080] In classical flight control, the command corresponds to the position (angle) of the aerodynamic control surface. Highly dynamic (rigid) position control of the actuators ensures that the actual control surface position precisely follows the control command. The control structure corresponds to a cascade control with an inner control loop, the actuator control (ACL), and an outer control loop, the flight control (FCL). Attitude angle, yaw rates, and velocities are fed back. Acceleration measurements are generally used only for observation or as a substitute for poorly measurable conditions.

[0081] Rudder hinge torque-based flight control is also known. The FCL command corresponds to a torque command, i.e., a direct current command, for the actuator. In the equilibrium state, the torque command corresponds to the aerodynamic rudder hinge torque. The concept is similar to the pilot's force-oriented control behavior during manual control. Such control is intended to offer advantages for smooth flight and load reduction, as the control surface evades the gust due to a changed hinge torque. This is intended to reduce actuator load and avoid force fight with redundant actuators.

[0082] Incremental nonlinear inversion (INDI), also known previously, involves local linearization and inversion of the system dynamics. Incremental increases in the control commands are calculated. The method is based on measured and commanded (rotational) accelerations and reduces the influence of (aerodynamic) model accuracy and the center of gravity position for increased robustness. The control law is based on the comparison between planned and actual changes (thus derivatives) of the state variables, which are calculated or observed based on rotational and translational acceleration measurements. In contrast to the concepts disclosed, a direct application of this control variable change in an internal cascade of the servo control or an extension of the INDI approach to actuator dynamics is not known.In a proposed approach, the actuator current serves as a reference variable, and a modified control law is derived. In contrast to the disclosed approach, the quasi-stationary dependence of the actuator current on the rudder hinge torque is assumed, so that the dynamics of the control system itself remain uncontrolled.

[0083] For active flutter control and load reduction, it is known to use acceleration measurements or modal degrees of freedom. In this case, the command corresponds to the position of the platform. Alternatively, additional forces are applied by vibration actuators. This is often not done for the purpose of control, but rather specifically to compensate for individual resonance frequencies.

[0084] In existing systems, the dynamics (bandwidth) of the flight controller significantly determines the precision of path and attitude control (disturbance suppression), the smoothness of the flight (disturbance suppression), and the agility of the path guidance (guidance behavior). The maximum usable bandwidth is limited by the dynamics of the independently designed actuator control (inner control loop) and, if applicable, by the dynamics of the mechanical transmission path between the actuator and the landing surface, the structural dynamics of an elastic aircraft, and the unsteady aerodynamics. For optimal FCL design, a precise aerodynamic model is required. This is complex and can be fraught with insufficient robustness. The inner control loops, at least the attitude control, must be designed individually for each aircraft type. A precise aeroelastic model is required to exclude excitations from structural dynamics.A direct impact of the flap deflection η k on the vertical load factor (i.e., the load acceleration) nz complicates the design of a gust load control system. Without the provision of a pre-control, which requires complex angle of attack measurement, the mitigation potential is limited.

[0085] The Fig. 5shows a schematic representation of a concept for acceleration-based roll attitude control of an aircraft. In comparison to the conventional system, the position control of the actuators is replaced by a feedback of an acceleration measurement, which records the aerodynamic force effect of the landing surface (e.g., local acceleration at the landing surface or rotational acceleration of the aircraft). This changes the classic division between actuator control and flight control. The interface between the FCL and actuator control moves inward one cascade. The state feedback of the actuator deflection is replaced by an output feedback of the acceleration proportional to it, which also includes the disturbance influence (gusts). The FCL command then corresponds to the actuating rate (angular velocity) of the aerodynamic landing surface. A measurement of the landing surface position is only required to take the actuating constraint into account.

[0086] For an aircraft, the controlled system of the FCL, the roll moment coefficient CI, with C Iξ, is proportional to the aileron deflection ξ, which in known systems represents the manipulated variable. The roll moment coefficient CI, with C Iβ, is proportional to the sideslip angle β, which is considered a disturbance variable for pure roll control and in particular includes the wind influence β w . The roll moment coefficient CI, with C Ip, is proportional to the dimensionless roll rate p*=p b / VA . The roll moment is determined from the coefficient CI by multiplying it by the reference variables (q, S, b); the roll acceleration is proportional to it with the inverse roll inertia (1 / I yy ). Roll rate p and roll angle Φ are obtained by integrating the roll acceleration. Known flight control includes full feedback of the states "roll rate (p)" and "hang angle (Φ)".The setup is a cascade control system, where the outer control loop comprises the roll attitude with the reference variable "roll command (Φ c )" and the manipulated variable "roll rate command (pc )", which is proportional to the control error Φ c - Φ with the gain K Φ . The inner control loop then concerns the roll rate with the reference variable "roll rate command (pc )" and the manipulated variable "aileron command (ξ c )", which is proportional to the control error pc - p with the gain K p .

[0087] For the actuator, the controlled system ACL, the actual current flow I corresponds to the current command I c if the electrical time constant is neglected. The torque is proportional to the current flow with the torque constant K t . Other torque contributions (friction, aerodynamic rudder hinge torque, etc.) are neglected. The angular acceleration (ω̇) of the output follows from the conservation of angular momentum to torque / moment of inertia (J). The actuating speed (ω) and the output angle (which corresponds to the aileron deflection ξ) are determined by integrating the angular acceleration. The familiar actuator control then includes complete feedback of the states "actuating speed (ω)" and "actuating position (ξ)".The setup is based on a cascade control system, where the outer control loop comprises the control position with the reference variable "aileron command (ξ c )" and the manipulated variable "actuator rate command (ω c )", which is proportional to the control error ξ c - ξ with the gain K ξ . The inner control loop then concerns the control speed with the reference variable "actuator rate command (ω c )" and the manipulated variable "current command (Ic)", which is proportional to the control error ω c - ω with the gain K ω .

[0088] In comparison, the Fig. 5for an acceleration-controlled concept with the controlled system aircraft 20, actuator 21, and the controlled system actuator 22. The measured roll acceleration (ṗ) is fed back instead of the proportional aileron deflection (ξ). The FCL command ω c corresponds to the control rate ξ̇. The disclosed feedback not only records the control surface position ξ but also directly detects the disturbance caused by β or β w . Thus, the disturbance is compensated one control loop further inside than in the known flight control system. With a highly dynamic design of this acceleration control loop (p-feedback), significantly better disturbance suppression can be achieved. The default value of the attitude control loops (traditional "inner loops" of the FCL) corresponds directly to the rate acceleration (second derivative of the controlled variable). With a highly dynamic design of the acceleration control, the aircraft immediately follows the specified rate acceleration (ṗ c ≈ ṗ).This results in simple, linear behavior independent of aircraft-specific parameters. The design of the attitude control can be standardized and independent of the aircraft type and flight condition. In a highly dynamic design, the acceleration feedback via K ṗ · C Iξ becomes dominant over the other aerodynamic influences (via C Iβ and C Ip ), thus reducing the influence of the aerodynamic parameters (except C Iξ ) on the control loop. This means that a precise aerodynamic model is not necessary for the FCL design. Only the rudder effectiveness C Izi and the dynamic pressure remain relevant. Avoiding the direct impact of the control command on the acceleration measurement simplifies the design and increases the potential for control-based gust load mitigation.

[0089] For rigid aircraft, the principle can be applied to the pitch degree of freedom (measured variable: pitch acceleration, primary manipulated variable: elevator), the yaw degree of freedom (measured variable: yaw acceleration, primary manipulated variable: rudder), the lift degree of freedom (measured variable: vertical acceleration nz, primary manipulated variable: flap), the longitudinal degree of freedom (measured variable: longitudinal acceleration nx, primary manipulated variable: spoiler), and the lateral degree of freedom (only for lateral force control). Ideally, the acceleration component is fed back not only to the primary manipulated variable, but to all manipulated variables that influence the respective degree of freedom, for example, via aileron roll moment, rudder yaw moment, elevator lift, or flap pitch moment. By selecting a suitable gain matrix, the degrees of freedom can be completely decoupled. The descriptive degrees of freedom whose acceleration is measured can be freely selected.An example is the rotation and translation of the center of gravity in aircraft-fixed coordinates. Other coordinate systems and other (possibly multiple) reference points of the rigid body are also conceivable, for example, the vertical position of both wingtips instead of the roll angle. Any combination of independent degrees of freedom that uniquely describes the system is possible. This represents a valid set of generalized coordinates (q) in the sense of the Lagrange formalism.

[0090] The Fig. 6Using a schematic representation of a concept for acceleration-based control, it demonstrates the transferability of the disclosed control concept to general mechanical systems. The principle is transferable to any mechanical system 23 with n degrees of freedom that can be uniquely described by generalized coordinates in the sense of the Lagrange formalism, which is controlled by one or more manipulated variables that have a direct force or torque influence on the system 23 and whose manipulated variables are operated by a controlled actuator 24 with at least single-integrating behavior (all mechanical actuators).

[0091] In contrast to the system according to the Fig. 6 , position-controlled approach, the manipulated variable ui has a force influence, which is determined by generalized forces Q i = B ij ( q , q)·ui. The same applies to force or moment disturbances zi with the effect factors E ij ( q , q ). The acceleration q i of the generalized coordinate is with M ij -1< ( q ) proportional to the generalized force Q i . Generalized velocities q and coordinates q follow by integration. Generalized velocities q cause non-conservative "damping forces" in dissipative systems D ( q , q )· q Conservative forces are C ( q ) proportional to generalized coordinates q . In known control concepts, a complete feedback of all states is carried out, represented by the generalized coordinates ( q ) and speeds ( q). The structure is in the form of a cascade control, with the outer control loop dealing with generalized coordinates. The reference variables are the setpoints of the generalized coordinates ( q c ). The control variables are the setpoints of the generalized speeds ( q c ), which is connected to the gain matrix K q proportional to the control error q c- q The inner control loop concerns generalized speeds, where the reference variable is the setpoint of the generalized speeds ( q c ) and the control variable control commands ( u c ) which are connected to the gain matrix K q̇ proportional to the control error q c- q are.

[0092] The actuator has an arbitrary transfer behavior G(s) between commanded and actual change of the manipulated variable u̇, but at least one integration stage. The actuator control includes the feedback of (at least) the manipulated variables u, where the reference variable is the setpoint of the system manipulated variables u c. The actuator control variable is the setpoint of the system control rates u̇ c , which with reinforcements K u proportional to the control error u c- u is.

[0093] In comparison to known control systems, the acceleration-controlled approach is based on the Fig. 6 a feedback of the measured generalized accelerations q instead of the proportional control variables u. The control command corresponds to the control rate u̇ .

[0094] In the Fig. 7 A concept for acceleration-based control of elastic aircraft is illustrated. Fig. 7the general case of complete state feedback. An elastic aircraft is a special case of the Fig. 6 explained concept according to the disclosure, since it can be described by the Lagrange formalism. The choice of generalized coordinates is, in principle, arbitrary. One possibility is individual positions of acceleration sensors distributed throughout the aircraft. In this case, the measured acceleration corresponds directly to the generalized acceleration. q This requires at least six sensors to record the rigid body motion. The number of additional sensors determines how many elastic modes can be recorded. An alternative approach involves separating rigid body motion (mean axes) from structural dynamics. This involves a division into rigid body degrees of freedom (x = [x, y, z, Φ, Θ, Ψ] T< ) and amplitudes ( n = [η 1 , η 2 , ... , η n ] T< ) of the elastic modes, so that q= [xi , η i ] T< . The equations of motion of rigid body motion and structural dynamics are inertially decoupled, but a coupling via external forces (aerodynamics) is given.

[0095] The controlled system includes the rigid body dynamics (in the Fig. 7 below), which is analogous to the system of Fig. 6 is constructed, whereby the correspondences q = x, B = B x , and E = E x apply. Regarding structural dynamics (above in the Fig. 7 ) the manipulated variable ui has a force influence which is determined by generalized forces Q j .= B η,ij ( q , q ). The same applies to force and moment disturbances (zi ) with the effect factors E η,ij ( q , q). The second derivative of the modal amplitude η ï is proportional to the generalized force Q j with the inverse modal mass matrix µ -1< (ij). The structural damping induces damping forces that are expressed by damping factors β proportional to the rate of change of the modal amplitudes η̇ i. The structural elasticity causes spring forces that are related to the generalized stiffness matrix γ are proportional to the modal amplitudes.

[0096] There is a coupling between structural and rigid body motion through external forces. The external forces (aerodynamic forces / moments) depend on rigid body states. ẋ and x and structural dynamics states η̇ and n . External forces influence both the rigid body motion ( ẍ) and the structural dynamics (η̈). The influence of the rigid body motion dependent part of the forces on the rigid body motion is already D , C The influence of the structural deformation-related component of the forces on the structural dynamics ( Q η and Q η̇ ) is (in contrast) not already in β , γ The dependence of the forces on the rigid body motion has an influence on the structural dynamics: Q x , Q ẋ . The outer control loops refer to the generalized coordinates ( q = [xi , η i ] T< ), the inner control loops on the local degrees of freedom ( R f ). The system behavior depends on the form of description (transformation between different degree of freedom systems / state representations).

[0097] A special case of the concept according to the Fig. 7refers to a local acceleration feedback. Here, the acceleration measurement is performed directly at the location of the platform. The feedback matrix KR is only occupied diagonally, meaning that the acceleration only affects the positioning surface at which the measurement is taken. The system dynamics can be freely specified, provided that the number of positioning / measurement positions equals the number of (considered) degrees of freedom (and is suitably chosen, i.e., is linearly independent; this results in controllability and observability). However, a complete eigenstructure specification is not possible. The system remains coupled. Assuming that the positioning rate command is implemented without delay (neglecting actuator dynamics), the feedback of the acceleration to the positioning rate is equivalent to the feedback of the velocity to the positioning position. Assuming that the aerodynamic force generation by the positioning surface occurs without delay (neglecting unsteady aerodynamics), a counterforce proportional to the velocity is generated.The acceleration feedback thus corresponds to a virtual viscous damper acting at the location of the landing surface. By introducing an integral component or feedback from a local velocity measurement, a virtual spring element could be represented analogously, resulting in a more or less rigid clamping of the wing at the location of the landing surface. Since the damping force always counteracts the direction of motion, any energy input and thus a destabilization of the structural dynamic modes is excluded. However, this only applies as long as the assumptions are justified, i.e., for all structural modes that are significantly lower in frequency than the actuator dynamics / unsteady aerodynamics. This limits the maximum achievable dynamics of the acceleration feedback.

[0098] In contrast, there is a risk of excitation when the acceleration measurement is spatially separated from the support surface (e.g., an IMU in the cockpit), since the acceleration signal reacts to the force generated at the support surface only with a delay due to structural dynamics. In other words, it is possible for a mode shape to exist whose antinode at the measurement location has an opposite sign to that at the location of the force generation (support surface). Consequently, a negative, destabilizing damping force results in the frequency range of this mode. This is excluded if the measurement is performed directly at the location of the force generation.

[0099] In classical flight control, the bandwidth limitation results from the frequency separation of structural dynamics. In contrast, in local acceleration feedback, the bandwidth limitation results from the frequency separation of unsteady aerodynamics and actuator dynamics.

[0100] Advantages of the acceleration-controlled concept can include the ability to achieve higher dynamic flight control and thus better disturbance suppression, smoother flight, and greater agility, especially in highly elastic aircraft, where frequency separation for structural dynamics or filtering of elastic modes is not required. Automatic damping of all elastic modes below the actuator dynamics and, in particular, the aerodynamics can be enabled, regardless of the specific elastic properties of the aircraft. Disturbances (local gusts) are compensated directly at the point of attack without exciting the structural dynamics (similar to a bird that only spreads its feathers locally to accommodate a gust). Structural loads are reduced. Acceleration measurement and control can be integrated into an actuator control system (smart actuator). This enables a decentralized system design.The acceleration-controlled concept enables new redundancy concepts and allows for easy adaptation of the control laws in the event of a parking space failure. With sufficient frequency separation between acceleration control (decentralized in the actuator) and attitude control (centralized in the FCC), the attitude control and all higher-level control loops can remain unchanged, as even reduced acceleration control dynamics are still sufficiently fast. This corresponds to the fact that, in conventional flight control, the failure of a redundant actuator for the same parking space generally does not require adaptation of the FCL.

[0101] The Figures 8A to 8E show the overall system dynamics in the complex plane of an exemplary embodiment of the arrangement, which results from the transmission of the Fig.5 structure shown on the pitch degree of freedom. Here, the Fig. 8B an enlarged section of the Fig. 8A , the Fig. 8Can enlarged section of the Fig. 8B , the Fig. 8D an enlarged section of the Fig. 8C and the Fig. 8E an enlarged section of the Fig. 8D . The representation of the Figures 8A to 8Eis based on a disclosed device for controlling the horizontal stabilizer. The longitudinal position θ and the pitch rate q are fed back. Shown is the pole and zero distribution resulting from the illustrated embodiment for different feedback gains (+-shaped markings). For comparison, the pole and zero distribution resulting from a previously known control system in which the elevator pitch angle is controlled by the control system control (x-shaped markings) is also shown. The larger x and + markings indicate the pole positions resulting when both feedback gains (for q and θ) assume the value zero. For the previously known control system (x), this pole distribution corresponds to a pattern known for uncontrolled aircraft, in which one conjugate complex pole pair is assigned to the phygoid motion and one to the angle of attack oscillation.By returning the local acceleration in the disclosed embodiment, the angle of attack oscillation is strongly damped and split into two real poles. The phygoid poles, however, are barely affected.

[0102] The smaller markings connected by lines indicate pole positions that can be achieved by simultaneously increasing the feedback gains for pitch rate and longitudinal attitude in a constant ratio. The star-shaped markings indicate pole positions that can be achieved by selecting the feedback gains advantageously when a high control loop bandwidth is desired with a damping factor not less than 0.7.

[0103] The Fig. 9 shows a Bode diagram for a previously known 25 and for the disclosed 26 embodiment according to the Figures 8A to 8Ewhich can result in a particularly advantageous choice of feedback gains. Shown is the frequency response of a disturbance transfer function from the vertical wind speed w wg (gust) to the longitudinal position θ. In a frequency range below the dynamics of the positioning system control, the disclosed embodiment 26 (solid line) exhibits improved disturbance behavior, since the transfer function of the previously known arrangement 25 (broken line) has an additional zero that almost compensates for the pole associated with the position control loop of the servo control. This zero is eliminated by the acceleration feedback.

[0104] The Fig. 10 illustrates an exemplary embodiment for a flexible aircraft. Without limiting the generality, the representation of the Fig. 11This is limited to the flexibility of the main wing with respect to bending around the longitudinal axis, torsion around the transverse axis, and the resulting local vertical movements. In addition to the six degrees of freedom of a rigid body, a flexible aircraft has further degrees of freedom that describe the deformation state. A common representation includes, in addition to the rigid body degrees of freedom of an undeformed reference configuration 27, the amplitudes of superimposed eigenmodes of the elastic modes, which describe characteristic deformation patterns compared to the reference configuration. One such eigenmode is exemplified in the Fig. 10shown. The eigenform can have local extreme points 28, at which the deviations from the reference configuration are greatest, and nodal points 29, at which the deviations from the reference configuration disappear. A flexible aircraft can have several control systems (tailplanes) for controlling and stabilizing the deformation degrees of freedom, which can be formed by control surfaces 30 and optionally fins 31. A device according to the disclosure can use one or more actual accelerations at arbitrary points 32 of the aircraft for control. In an advantageous embodiment, these can be, for example, extreme points or nodal points of one or more eigenforms, or points at which none of the eigenforms relevant for control has a nodal point.In particular, the number of accelerations used can correspond to a number of eigenvalues ​​that are to be stabilized or influenced by the control system. For example, all accelerations can be received by each of the provided devices. Alternatively, only those accelerations that can be influenced by adjusting the control system controlled by the device can be received by a device.

[0105] In the Fig. 11An embodiment of an arrangement for a flexible aircraft is shown, in which the points 32 of the aircraft at which the local acceleration is used to control the positioning systems are located near the positioning surfaces 30. For example, all accelerations can be received by all devices according to the disclosure. Alternatively, each device can receive only the acceleration that is present in the vicinity of the positioning system controlled by that device.

[0106] The features disclosed in the above description, the claims and the drawings may be important for the realization of the various embodiments both individually and in any combination.

Claims

1. Apparatus (12) for the control and closed-loop control of an actuation system (4) of an aircraft, with - a first input interface, which is configured to receive first input data specifying a reference variable (13); - a second input interface, which is configured to receive second input data specifying a controlled variable (14); and - a control output, which is configured to output a control signal, characterised in that, the control signal specifies a manipulated variable (8) for an actuation system (4) of an aircraft, which is to be controlled by means of the actuation system (4), wherein - the reference variable (13) specifies a target acceleration at a point (32) of the aircraft, which is to be controlled by means of the actuation system (4); - the controlled variable (14) specifies an actual acceleration of the aircraft at the point (32); and - the apparatus (12) is configured to determine the manipulated variable (8), taking into account the reference variable (13) and the controlled variable (14), in particular from the difference between the reference variable (13) and the controlled variable (14), and to output the control signal corresponding to the manipulated variable (8) via the control output.

2. Apparatus (12) in accordance with Claim 1, with a third input interface that is configured to receive third input data specifying an actuation system-controlled variable (7), wherein the apparatus (12) is configured: - to determine an actuation system-reference variable, taking into account the reference variable (13) and the controlled variable (14), and - to determine the manipulated variable (8), taking into account the actuation system-reference variable and the actuation system-controlled variable (7).

3. Apparatus (12) in accordance with Claim 2, wherein the actuation system-reference variable is a target positioning speed of the actuation system (4), and the actuation system-controlled variable (7) is an actual positioning speed of the actuation system (4).

4. Apparatus (12) in accordance with at least one of the preceding claims, wherein the apparatus (12) is configured to determine the manipulated variable (8) without considering an actual actuation position of the actuation system (4), and without determining a target actuation position of the actuation system (4).

5. Apparatus (12) in accordance with at least one of the preceding claims, with a further input interface, which is configured to receive further input data specifying a further controlled variable, wherein - the further controlled variable specifies an actual acceleration of the aircraft at a further point (32); and - the apparatus (12) is configured to adjust the controlled variable (14), taking into account the further controlled variable, and subsequently to determine the manipulated variable (8), taking into account the reference variable (13) and the controlled variable (14).

6. Arrangement for the control and closed-loop control of an actuation system (4) of an aircraft, with - an aircraft, comprising - an actuation system (4), which is configured to control the aircraft in at least one degree of freedom, and - an acceleration sensor (15), which is arranged at a point (32) of the aircraft; - a flight control apparatus (1) with an output interface; and - an apparatus (12) in accordance with at least one of the preceding claims, wherein - the flight control apparatus (1) is configured to calculate, from a flight state of the aircraft, the reference variable (13), which specifies a target acceleration at the point (32) of the aircraft, and, via the output interface, to transmit first input data, specifying the reference variable (13), to the first input interface of the apparatus (12); - the acceleration sensor (15) is configured to measure the local acceleration of the aircraft at the point (32), and to transmit second input data, specifying the controlled variable (14), to the second input interface of the apparatus (12), which data specify the local acceleration at the point (32); and - the actuation system (4) is configured to receive the manipulated variable (8) from the control output of the apparatus (12), and to execute a positioning movement corresponding to the manipulated variable (8).

7. Arrangement in accordance with Claim 6, wherein the flight control apparatus (1) is configured to calculate the reference variable (13) taking into account an actuating variable determined in a directly kinematic manner from a target trajectory of the aircraft.

8. Arrangement in accordance with Claim 6 or 7, wherein the actuation system (4) is formed with an actuator, which moves a flight control surface (30) of a flight control surface assembly (16) of the aircraft.

9. Arrangement in accordance with Claim 8, wherein the acceleration sensor (15) is arranged on a part of the flight control surface assembly (16) that is immovable relative to the aircraft.

10. Arrangement in accordance with at least one of the Claims 6 to 9, with a further acceleration sensor (15), which is arranged at a further point (32) of the aircraft, wherein - the apparatus (12) is an apparatus (12) in accordance with Claim 5; and - the further acceleration sensor (15) is configured to measure the local acceleration at the further point (32), and to transmit further input data specifying the further controlled variable (14) to the further input interface of the apparatus (12), which data specify the local acceleration at the further point (32).

11. Arrangement in accordance with at least one of the Claims 6 to 10, with a further apparatus (12) in accordance with at least one of the Claims 1 to 5, wherein the aircraft has a further actuation system (4), which is configured to control the aircraft in the at least one degree of freedom, or in at least one further degree of freedom, and an additional acceleration sensor (15), which is arranged at an additional point (32) of the aircraft, wherein - the flight control apparatus (1) is configured to additionally transmit the first input data specifying the reference variable (13) to the further apparatus (12) via the output interface; - the additional acceleration sensor (15) is configured to measure the local acceleration of the aircraft at the additional point (32), and to transmit second input data specifying an additional controlled variable (14) to the further apparatus (12), which data specify the local acceleration at the additional point (32); and - the further actuation system (4) is configured to receive the manipulated variable (8) from the control output of the further apparatus (12), and to execute a positioning movement corresponding to this manipulated variable (8).

12. Method for the control and closed-loop control of an actuation system (4) of an aircraft, with the steps: - Provision of an apparatus (12) for the control and closed-loop control of an actuation system (4) of an aircraft; characterised by: - Generation of first input data specifying a reference variable (13), wherein the reference variable (13) specifies a target acceleration at a point (32) of the aircraft, which is to be controlled by means of the actuation system (4); - Generation of second input data specifying a controlled variable (14), wherein the controlled variable (14) specifies an actual acceleration of the aircraft at the point (32); - Reception of the first input data at a first input interface of the apparatus (12); - Reception of the second input data at a second input interface of the apparatus (12); - Determination of a manipulated variable (8) for an actuation system (4) of the aircraft, taking into account the reference variable (13) and the controlled variable (14), in particular from the difference between the reference variable (13) and the controlled variable (14); and - Output of a control signal specifying the manipulated variable (8) via a control output of the apparatus (12).

13. Method in accordance with Claim 12, comprising the reception of third input data specifying an actuation system-controlled variable (7) at a third input interface of the apparatus (12), wherein the determination of the manipulated variable (8), taking into account the reference variable (13) and the controlled variable (14), comprises the determination of an actuation system-reference variable, taking into account the reference variable (13) and the controlled variable (14), and the determination of the manipulated variable (8), taking into account the actuation system-reference variable and the actuation system-controlled variable (7).