Device, arrangement and method for regulating and controlling an actuation system of an aircraft

Abstract

A device for regulating and controlling an actuation system of an aircraft is disclosed. The device has a first input interface configured to receive first input data describing a reference variable; a second input interface configured to receive second input data describing a controlled variable; and a control output configured to output a control signal. The control signal describes a manipulated variable for an actuation system of the aircraft, by means of which the aircraft is controlled. The reference variable describes a target acceleration at a certain point of the aircraft controlled by means of the actuation system, and the controlled variable describes an actual acceleration of the aircraft at this point. The device is configured 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 a control signal corresponding to the manipulated variable via the control output. Furthermore, facilities and methods for regulating and controlling an actuation system of an aircraft are disclosed.

Description

Technical Field

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

[0002] The actuation system of an aircraft is, for example: a tail, such as an elevator tail, a stabilizing tail and an aileron, which has an actuator assigned to an adjustment surface for moving the adjustment surface; as well as a nozzle; a propeller; a lift auxiliary device, such as a flap, a spoiler (spoiler); and a lateral force control unit.

[0003] Conventionally, cascade control is used in the actuation systems of aircraft. In this case, in the outer control loop, i.e. the flight control, a reference variable for regulating an actuator is determined as a function of the state variable of the aircraft with respect to the degree of freedom to be regulated as the manipulated variable of the flight control, which actuator controls the relevant degree of freedom. Thus, in the case of pitch attitude control, the reference variable for the actuator control is the actuator deflection and thus the swing amplitude of the elevator. The elevator is usually moved by means of an actuator, for example a translational, hydraulic or rotary electromagnetic actuator, so that the reference variable corresponds to the target rotational position of the actuator. The actuator control with the reference variable as input variable forms the inner control loop. In the case of a rotary electromagnetic actuator for moving the adjustment surface of the tail, this is in particular a so-called servo control, in which the rotational speed (manipulation speed) of the actuator is used as the controlled variable of the inner control loop, in which the corresponding reference variable (target rotational speed) of the inner control loop is determined by the target rotational position. The actuator current for the operation of the actuator is determined by the control deviation of the inner control loop, i.e. the difference between the target rotational speed of the actuator and the actual rotational speed. In addition, a further inner control loop can be provided, in which the actuator current is the controlled variable and the actuator voltage is the manipulated variable. The regulating acceleration of the actuator can be taken into account in the servo control. The measurement of the acceleration can be provided in the flight control for observation or as a substitute for states that are difficult to measure.

[0004] Typically, servo control has significantly faster dynamics than flight control, so that flight control limits the overall dynamics of the control system. However, against the background of the rapidly increasing popularity and importance of unmanned and / or autonomously operated flying objects, new fields of application for flight control technology are emerging, which require highly dynamic control of flight dynamics. In addition to precise route guidance on complex trajectories, many measurement and observation tasks require improved attitude stability and flight stability. In addition, the trend in aircraft development is emphasized towards more efficient, aerodynamic, high-quality configurations, which are characterized by a high aspect ratio and a long span. Especially in connection with increasingly larger, lighter and therefore more elastic structures and connections, they require a reduction of structural loads in control technology and an active stabilization of relatively low-frequency modes of structural dynamics, the natural frequencies of which may exceed the dynamics of known control systems.

[0005] What is suggested in document DE 102016 117 634A1 is that, compared with the known positioning regulation (actuator swing amplitude as reference variable), a method of force or torque regulation is used instead. For this reason, a force / torque regulator is provided for each actuator, and the actuator uses the force / torque regulator to regulate based on the assigned reference variable, i.e., target force, target force change, target torque or target torque change, and a controlled variable, i.e., the force generated by the actuator or the torque generated by the actuator. Here, the controlled variable is obtained by a sensor device present on or in the actuator or in the transmission system of each actuator. Therefore, the controlled variable, i.e., the driving torque in the tail wing, is measured on the mechanical transmission path between the force generation and the aerodynamic adjustment surface in the actuator. The driving torque causes the rotational acceleration of the adjustment surface relative to the aircraft. At the same time, due to the air force, the adjustment surface is subjected to the aerodynamic control surface hinge torque. In the quasi-steady state, there is a balance between the drive torque and the aerodynamic control surface joint torque, so that in this case the reference variable corresponds to the control surface joint torque. The control surface thus "yields" or "elasticly" reacts to the gust loads applied to the outside of the control surface based on force regulation. According to the method proposed in DE 102016 117 634 A1, the feedback proportional to the swing amplitude of the control surface is omitted.

[0006] This makes the design of the system dynamics limited compared to full state vector feedback.

[0007] Document DE 102016 117638 A1 discloses a system for such force / torque control, wherein the influence of gusts on the aircraft is minimized by corresponding over- or under-compensation of the hinge torque. Here, the force component / torque component caused by the gust is known, and the predetermined target value for the control force / torque is modified depending on the influence of the gust on the hinge torque of the control surface. An additional damping term proportional to speed should reduce the oscillation tendency of the natural rudder angle dynamics (rudder surface flutter). According to one design, the control position (actuator deflection) and the control speed can be fed back, wherein this is not carried out in a cascaded manner, but as direct, mutually independent feedback. Summary of the invention

[0008] The object of the present invention is to specify a new technology for regulating and controlling an aircraft actuation system which allows a rapid and precise regulation, in particular taking into account the various factors that influence the flight behavior.

[0009] In order to achieve this object, a device according to the invention for regulating and controlling an actuation system of an aircraft is provided, as well as a device and a method according to the invention for regulating and controlling an actuation system of an aircraft.

[0010] According to one aspect, a device for regulating and controlling an actuation system of an aircraft is provided. The device is formed with a first input interface, which is set up for receiving first input data describing a reference variable; a second input interface, which is set up for receiving second input data describing a controlled variable; and a control output, which is set up for outputting a control signal, which describes a manipulated variable for an actuation system of the aircraft, by means of which the aircraft is controlled. The reference variable describes a target acceleration at a certain point of the aircraft controlled by means of the actuation system, while the controlled variable describes an actual acceleration of the aircraft at this point. The device is set up to determine the manipulated variable, in particular from the difference between the reference variable and the controlled variable, taking into account the reference variable and the controlled variable, and to output a control signal corresponding to the manipulated variable via the control output.

[0011] According to another aspect, a facility for regulating and controlling an actuation system of an aircraft is provided. The facility is formed by an aircraft, a flight regulation device with an output interface, and a device for regulating and controlling an actuation system of an aircraft according to the disclosure. The aircraft has an actuation system configured to control the aircraft in at least one degree of freedom and an acceleration sensor arranged at a certain point on the aircraft. The flight regulation device is configured to calculate a reference variable (which describes a target acceleration at this point of the aircraft) from the flight state of the aircraft and to send first input data describing the reference variable to a first input interface of the device via the output interface. The acceleration sensor is configured to measure a local acceleration at this point of the aircraft and to send second input data describing a controlled variable to a second input interface of the device, the second input data describing the local acceleration at this point. The actuation system is configured to receive a manipulated variable from a control output of the device and to implement a regulating movement corresponding to the manipulated variable.

[0012] According to another aspect, a method for regulating and controlling an actuation system of an aircraft is provided. The method comprises the following steps: providing a device for regulating and controlling an actuation system of an aircraft; generating first input data describing a reference variable, wherein the reference variable describes a target acceleration at a certain point of the aircraft, and the aircraft is controlled by means of the actuation system; generating second input data describing a controlled variable, wherein the controlled variable describes an actual acceleration of the aircraft at this 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 the actuation 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 describing the manipulated variable via a control output of the device.

[0013] The actuation system can be, for example, a tail and an actuator for operating the tail. Here, the actuator can be a system for generating an adjustment force or an adjustment torque without an associated adjustment system. Alternatively, the actuator can have adjustment, especially servo adjustment. The tail can include a fixed, i.e., a part (stabilizer) that cannot move relative to the aircraft and a movable adjustment surface, which exerts a desired control effect on the aircraft according to its position relative to the fixed part. The actuator can be a rotary electromagnetic actuator, i.e., an electric motor, which rotates the adjustment surface relative to the fixed part of the tail. The tail can be a lift tail, a directional tail or an aileron. Alternatively, the actuation system can be, for example, a flap for controlling the lift of the aircraft, a spoiler for longitudinal control of the aircraft, or a lateral force control for lateral control of the aircraft. Other examples for the actuation system include the nozzle and propeller of the aircraft.

[0014] In the sense of the present disclosure, the acceleration of the aircraft is the acceleration of the aircraft itself and therefore, for example, not the adjustment acceleration of individual actuators arranged in the aircraft. In particular, the acceleration of the aircraft can exclude the acceleration of elements arranged on the aircraft so as to be movable relative to the aircraft, such as, for example, a tailplane, which contains adjustment surfaces relative to the acceleration of the aircraft.

[0015] For example, the acceleration of the aircraft may be the acceleration at the center of gravity of the aircraft, or may be the acceleration at a fixed part of the tail (stabilator), wing or fuselage. For example, the acceleration of the aircraft may be the acceleration at a point near an adjustment surface, in particular the adjustment surface controlled by the actuator. In particular, the acceleration of the aircraft may also be the acceleration at a part of the actuator itself that cannot move relative to the aircraft structure, such as a housing, a base plate, a circuit board of the control electronics or other non-movable parts of the actuator. In such an embodiment, the system integration can be simplified and can be independently produced by system component manufacturers.

[0016] In particular, the acceleration of the aircraft may be the acceleration of the aircraft itself relative to an inertial system or a global coordinate system. The acceleration of the aircraft may include the influence of the earth's gravitational field and / or compensate for it. The acceleration of the aircraft may be a translational acceleration, a rotational acceleration, or an acceleration including a translational and a rotational component.

[0017] Thus, the actual acceleration of the aircraft at this point provides a controlled variable that depends not only on the control amplitude of the actuating system (i.e., for example, the position of the control surface of the tail or the rotational or translational position of the actuator that moves the control surface), but also already includes other external influences, in particular the influence of gusts and / or different flight state variables, such as flight speed, flow angle and rotation rate, which all play a role in the aerodynamic force generation. This makes it possible to suppress interfering influences (gusts) in the inner, more dynamic control loop. A control that is insensitive to the specific aerodynamic or aeroelastic properties of the aircraft (robustness) can be provided. In addition, a simpler and more standardized design of the flight control and a more flexible and more precise flight guidance can be made possible.

[0018] The device may be formed with a third input interface, which is configured to receive third input data describing a controlled variable of the actuation system. Here, the device is configured to determine the actuation system reference variable taking into account the reference variable and the controlled variable, and to determine the manipulated variable taking into account the actuation system reference variable and the actuation system controlled variable.

[0019] The actuation system reference variable may be a target control speed of the actuation system, and the actuation system controlled variable may be an actual control speed of the actuation system. For example, the actuation system reference variable and the actuation system controlled variable may be a target rotation speed (target rotation rate) and an actual rotation speed of an actuator of the actuation system, for example an actuator that moves the adjustment surface of a tail wing. In this case, the manipulated variable may be a variable that causes the actuator to move, in the case of an electromagnetic actuator, in particular the actuator current.

[0020] Alternatively, an actuator reference variable, such as a target actuator current, can be determined for an actuator of the actuating system, taking into account an actuating system reference variable and an actuating system controlled variable. Taking into account the actuator reference variable and the actuator controlled variable, such as the actual actuator current, a manipulated variable can be determined, which can be in particular an actuator voltage, such as the terminal voltage of a direct current motor or the transverse voltage component in the case of a 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 under the control speed control known per se.

[0021] In general, the manipulation variable, optionally a subordinate manipulated variable in a subordinate control structure, is determined taking into account a variable which is predefined (i.e. used as a reference variable) and a variable which is fed back (i.e. used as a controlled variable), including the determination of the manipulated variable according to control technology. In particular, a cascade structure can be formed in which a control deviation as a difference between a reference variable and a controlled variable (which can optionally be determined as a controlled variable composed of a plurality of controlled variables) is multiplied by a proportionality factor in order to determine the manipulated variable. Alternatively or additionally, parallel feedback can be provided in which one or more controlled variables are fed back with respective adaptations, such as gain and / or integration, and are offset by addition or subtraction from a reference variable which has been modified by means of a prefilter according to the fed back controlled variable, in particular for compensating for static errors between the reference variable and one or more controlled variables.

[0022] The device can be set up to perform one, more than one or all of the following operations for determining the manipulated variable: determining the (target) manipulation speed (manipulated variable or actuation system reference variable) by multiplying the difference between the target acceleration (reference variable) and the actual acceleration (controlled variable) by a first proportionality coefficient; determining the (target) actuator current (manipulated variable or actuator reference variable) by multiplying the difference between the target manipulation speed (actuation system reference variable) and the actual manipulation speed (actuation system controlled variable) by a second proportionality coefficient; and determining the (target) actuator voltage (manipulated variable) by multiplying the difference between the target actuator current (actuator reference variable) and the actual actuator current (actuator controlled variable) by a third proportionality coefficient.

[0023] The actuator voltage may be a controlled manipulated variable. The actuator voltage may be adjusted as a function of the controlled variable and may thereby cause the actuation system to react to it, in particular to move the actuation system and to occupy a position of the actuation system, based on physical and physical system known conditions of the system, in particular physical and physical system known conditions of the actuation system associated with the system, i.e. the aircraft.

[0024] Alternatively, the actuator current can be the manipulated variable for regulation. In this case, a target actuator current corresponding to the manipulated variable can be pre-set to the actuator, wherein no feedback of the actual actuator current is performed. In this case, the actuator can be set up to convert the pre-set current into a corresponding voltage in order to achieve the pre-set current. In this case, the actuator can have an internal regulation, which, for example, uses the current as a reference variable and controlled variable and the voltage as a manipulated variable. In particular, it can be provided that the actuator current is the manipulated variable for regulation when the actuator achieves a pre-set current amount with sufficient dynamics, so that, depending on the relevant application, in particular the type of aircraft, a sufficiently dynamic regulation of the actuation system can be achieved without current regulation by the device.

[0025] The device can be set up for operation without taking into account the adjustment acceleration of the actuating system, in particular without taking into account the adjustment acceleration of the actuator, for example without taking into account the rotational or rotational acceleration of a rotary electromagnetic actuator for moving the adjustment surface of the tail. In particular, the determination of the manipulated variable can be carried out without taking into account the adjustment acceleration. The adjustment acceleration is proportional to the adjustment torque, in particular the drive torque of the actuator, while the acceleration of the aircraft at this point can be proportional to the control position of the actuating system. For example, the local acceleration at the immovable part of the tail, that is, at the stabilizer surface, can be proportional to the adjustment surface angle and the lift generated thereby. In particular, within the regulated system, there may be two integration stages between the adjustment acceleration and the acceleration of the aircraft. By feeding back the acceleration of the aircraft, that is, using the acceleration of the aircraft as the controlled variable, it is possible to influence the system dynamics in a manner similar to the feedback of the control position, for example the adjustment surface angle.

[0026] The device can be designed to determine the manipulated variable without taking into account the actual manipulated position of the actuating system and without determining a target manipulated position of the actuating system. The manipulated position of the actuating system can in particular be an actuator manipulated position, for example a rotational position or a translational position of an actuator for moving an adjustment surface of the tail, or a position of the adjustment surface of the tail corresponding to this rotational position or translational position.

[0027] Alternatively, it can be provided that the device is designed to determine the manipulated variable without determining a target actuation position of the actuation system, but taking into account the actual actuation position of the actuation system.

[0028] For example, the actual manipulated position can be taken into account within the limits of the motion space of the actuating system. In this case, the manipulated variable can be modified if it is checked that the manipulated variable would lead to an adjustment movement that exceeds the predefined motion space of the actuating system, so that the adjustment movement ends at the limits of the motion space. In this way, in particular, an adjustment movement can be provided, which corresponds to a function that is switched off when a limit switch is reached. In this case, the manipulated variable can also be determined without taking into account the actual manipulated position of the actuating system and without determining the target manipulated position of the actuating system.

[0029] Alternatively or additionally, the actual control position can be fed back, for example in order to observe non-steady-state aerodynamic or elastic and hysteresis states in the drive train in order to make it possible to further improve the control dynamics, wherein the control variable is determined without determining the target control position of the actuation system.

[0030] The device can have a further input interface, which is set up to receive further input data that describe a further controlled variable. In this case, the further controlled variable can describe the actual acceleration of the aircraft at another point, and the device can be set up to adjust 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. Adjusting the controlled variable with the help of the further controlled variable can, for example, include adding the further controlled variable to the controlled variable. The additional controlled variable can be received and used accordingly to adjust the controlled variable.

[0031] This makes it possible to determine a manipulated variable based on a controlled variable that describes the acceleration at a point on the aircraft where no acceleration measurement was performed, wherein the acceleration is determined by the acceleration at at least two other points on the aircraft where acceleration measurements were performed. This makes it possible to provide a virtual acceleration measurement at a measuring point different from the point and the other point of the aircraft. In particular, it is possible to determine the acceleration at different points of the aircraft and / or the acceleration in different degrees of freedom using a limited number of acceleration sensors at different points of the aircraft, which may have been arranged on or in the aircraft at least partially for other purposes. These accelerations can be used as controlled variables in a plurality of devices according to the present disclosure that are associated with the respective actuation systems for regulating and controlling actuation systems for different degrees of freedom of the aircraft, such as elevators, rudders and ailerons. For an aircraft that is considered to be rigid, six acceleration measurements can be provided at at least three different points and in at least three different directions in order to determine the acceleration at any point of the aircraft.

[0032] In general, the acceleration specified by the controlled variable can be determined from a plurality of measurements. In this case, the individual measurements of the acceleration can determine different measured variables in order to infer the acceleration from the controlled variable. For example, the roll acceleration, in particular the roll acceleration about the center of gravity of the aircraft, can be determined from two vertical movements of the wing surface, wherein the vertical movements are measured by means of corresponding sensors.

[0033] The components of the device can be provided as independent devices. Alternatively, a single or all components of the device can be provided as virtual components of a physical component. The components of the device configured as independent physical devices can be installed as a whole, or can also be constructed separately from each other.

[0034] The reference variable can be determined in an upstream control process, in particular in a flight control device. Here, the upstream control process can operate at a significantly lower clock rate than the downstream actuation system control (servo control).

[0035] With reference to this facility, the flight control device can be set up to calculate the reference variables taking into account the control variables kinematically determined directly from the target trajectory of the aircraft.

[0036] It can be provided that the reference variable is calculated as a function of the actual flight state of the aircraft and the target flight state of the aircraft, in particular from a deviation or difference between the actual flight state and the target flight state. Preferably, the reference variable can additionally contain a precontrol which is calculated directly from the target flight state and independently of the actual flight state.

[0037] The flight state can be given by one or more physical variables or measured values ​​of these variables, which completely or partially characterize the dynamic behavior of the aircraft or allow the determination of these characteristic variables (for example with the aid of an observer). In particular, the flight state can also include variables which are referred to as output variables in terms of control technology.

[0038] For example, the target flight state can be specified in the form of a time curve or a constant value for the physical variables used to describe the flight state. Alternatively or additionally, the target flight state can be specified in the form of a target trajectory for the aircraft. In this case, the determination of the pilot control contained in the reference variable can be realized in a particularly simple manner by determining the acceleration at a certain part of the aircraft from the target trajectory using known kinematic relationships.

[0039] For example, the target trajectory of an aircraft can be a desired positioning trend of the center of gravity of the aircraft in a plane corresponding to the flight altitude or in three-dimensional space. It can additionally contain the assignment of the positioning to time, that is, a line described in four-dimensional space. In addition, the target trajectory can describe the desired temporal or spatial correlation of the aircraft attitude, which can be represented by one or more angles, rotation matrices or quaternions, for example. The target trajectory can involve an arbitrary, preferably at least approximate inertial coordinate system. For example, the coordinate system can be a coordinate system fixed relative to the earth, or it can also be a coordinate system that moves with the air mass around the aircraft. These angles can be absolute angles, or they can also be angles related to the direction of the flight trajectory.

[0040] The actuation system can be formed with an actuator, which can move the adjustment surface of the tail of the aircraft. The tail can include a fixed part (stabilizer) that cannot move relative to the aircraft and a movable adjustment surface, which exerts a desired control effect on the aircraft according to its position relative to the fixed part. In this case, the actuator can be a rotary or translational electromagnetic actuator, that is, for example, an electric motor, which rotates the adjustment surface relative to the fixed part of the tail. Alternatively, other types of actuators can be provided, such as hydraulic or electro-hydraulic actuators, especially with one or more hydraulic cylinders. The tail can be a lifting tail, a directional tail or an aileron. In the case of a rotating actuator, for a design in which the device is set up to determine the difference between the actuation system reference variable and the actuation system controlled variable, the actuation system reference variable can be a target rotation speed (rotation rate) and the actuation system controlled variable can be an actual rotation speed.

[0041] The acceleration sensor can be arranged in particular on the part of the tail that cannot move relative to the aircraft. Thus, the device is set up to provide regulation based on the local acceleration on the tail of the aircraft as an overall system, wherein the adjustment surface of the tail is not detected relative to the acceleration of the aircraft, in particular relative to the acceleration of the non-movable part of the tail, and is therefore not included in the regulation. The non-movable part of the tail can be, for example, a stabilizer of a lifting tail or a directional tail, or a wing in the case of ailerons, flaps or spoilers. The non-movable part of the tail can also be part of the adjustment surface itself, as long as the acceleration measured there basically, that is, for regulation, reflects the acceleration of the aircraft as an overall system with sufficient approximation and the relative acceleration caused by the adjustment movement itself has only a secondary effect. For example, this can be an acceleration measurement on or near the axis of the rudder hinge or the installation position of the actuator. In particular, this can be provided for a pendulum rudder, in which there is no separation between the stabilizer and the flap, but the entire tail is adjusted.

[0042] Alternatively, the device can be formed with different types of actuation systems. For example, the actuation system can be a nozzle of the aircraft, a propeller of the aircraft, a flap, a spoiler, a lateral force control or another actuation system of the aircraft, wherein the actuation system can be adjusted by means of the device in any case and is designed to control the aircraft in at least one degree of freedom of the aircraft.

[0043] The facility can be formed with another acceleration sensor, which is arranged at another point of the aircraft. Here, the device is a device with another input interface, which is set up to receive another input data describing another controlled variable. The other acceleration sensor is set up to measure the local acceleration at another point and send another input data describing another controlled variable to another input interface of the device, which other input data describes the local acceleration at another point. With the help of the acceleration sensor and the other acceleration sensor, a virtual acceleration measurement can be provided at a measuring point different from the point and the other point of the aircraft. Here, the design explained above with reference to the device can be set accordingly with respect to the virtual acceleration measurement. The facility can include a plurality of other acceleration sensors, wherein the device has a corresponding number of input interfaces and with the help of the acceleration sensor, one or more virtual acceleration measurements can be provided at one or more points, which are different from the points at which the acceleration sensors are arranged.

[0044] The device can additionally have sensors that are not acceleration sensors. Alternatively or additionally, the acceleration sensor is formed by a plurality of sensors, which each measure a variable different from the acceleration, wherein the acceleration sensor determines the acceleration from the sensor variables. For example, the acceleration sensor can be formed by at least two sensors for detecting vertical movements on an aircraft wing surface, wherein the acceleration sensor determines the rolling acceleration of the aircraft from the vertical movement of the wing surface determined by means of the at least two sensors. The acceleration sensor can be formed by one or more force sensors, pressure sensors, strain sensors, motion sensors or positioning sensors, as long as their arrangement and the processing of their measured values ​​are set up to determine the local acceleration at a certain part of the aircraft.

[0045] The facility can have a further device according to the disclosure, wherein the aircraft has a further actuation system set up to control the aircraft in at least one degree of freedom or at least one further degree of freedom, and an additional acceleration sensor arranged at an additional point of the aircraft. In this case, the flight control device can be set up to additionally send first input data describing a reference variable to the further device via an output interface. The additional acceleration sensor can be set up to measure the local acceleration of the aircraft at an additional point and send second input data describing an additional controlled variable to the further device, the second input data describing the local acceleration at the additional point. In addition, the further actuation system can be set up to receive a manipulated variable from a control output of the further device and to implement a regulating movement corresponding to the manipulated variable.

[0046] In the corresponding embodiments, further devices and further actuation systems can be provided, wherein the above-mentioned design scheme for the actuation system can be correspondingly provided for the further actuation system. In this way, regulation can be provided for multiple or all degrees of freedom of movement of the aircraft. A plurality of devices according to the present disclosure can be provided as virtual devices in a physical device.

[0047] The reference variables provided by the flight control device can be used to specify the target acceleration of the aircraft in multiple degrees of freedom. The degrees of freedom of the aircraft can include, for example, the positioning in three spatial directions and three attitude angles. In the case of elastic aircraft, the degrees of freedom can also include variables for characterizing the deformation state, such as the amplitude of a mode. Alternatively or additionally, the degrees of freedom of the aircraft can be given by the positioning of the aircraft at different points in space.

[0048] As an alternative to providing the same (especially vectorial) reference variable in multiple devices according to the present disclosure, it can be provided that the flight control device provides each of the devices with a respective reference variable, wherein the respective reference variable indicates a target acceleration of the aircraft in a degree of freedom corresponding to the degree of freedom that is initially, primarily or exclusively influenced by means of an actuation system associated with the respective device.

[0049] In a design with more than one device according to the present disclosure, a single acceleration sensor or a single system of multiple acceleration sensors can be provided to replace the respective acceleration sensors associated with one of the devices, which determine the acceleration of the aircraft in multiple degrees of freedom and / or at multiple points of the aircraft, if necessary as virtual acceleration measurements, and provide the corresponding respective controlled variables for the device according to the present disclosure. The measured values ​​of one or more acceleration sensors can be provided to multiple devices according to the present disclosure. In this case, the number of acceleration sensors, reference variables and devices according to the present disclosure provided does not have to be consistent with each other. However, this can be the case in particularly advantageous embodiments, thereby making it possible to decouple the different degrees of freedom and to predetermine the system dynamics as desired.

[0050] The aircraft can be a highly flexible aircraft. Here, for example, a complete state feedback can be set in the following manner, that is, the acceleration is measured with a plurality of acceleration sensors distributed at positions on the aircraft, or separated from the structural dynamics via rigid body motion, wherein the amplitude is divided into rigid body degrees of freedom and elastic modes, wherein the motion equations of rigid body motion and structural dynamics are inertially decoupled, but coupling is given via external forces (aerodynamics). In the case of complete state feedback, the innermost control loop with feedback of the acceleration measured locally can be set. Here, in particular, when the number of acceleration points corresponds to the number of degrees of freedom considered, a complete influence on all vibration modes can be achieved. Separate regulation of rigid body motion and structural dynamics can be provided in the external control loop. For this purpose, regulation of rigid body motion and structural dynamics can be set in a cascade structure. The target value of the acceleration for the rigid body degrees of freedom and modal degrees of freedom can be converted into a target value for the local acceleration, for example, by means of eigenvectors and kinematic conversion ratios. The external control loop can involve generalized coordinates, while the internal control loop involves local degrees of freedom. The system behavior can be made independent of the description form, where conversion between systems and state representations with different degrees of freedom can be made possible.

[0051] Alternatively, for designs of aircraft with high flexibility or elasticity, local acceleration measurements can be provided on the actuation system, in the case of a tail, in particular directly at the adjustment surface on the immovable part of the tail, wherein the local acceleration is used only for the adjustment of the actuation system at which it is measured. The number of adjustment and measurement locations can be selected appropriately and correspond to the number of degrees of freedom taken into account, so that the system dynamics can also be freely predefined in this case. In the case of elastic aircraft, relative movements can occur between the part of the tail that is immovable relative to the aircraft and other parts of the aircraft, for example a part of the fuselage or the center of gravity of the aircraft, which are caused by elastic deformations.

[0052] In general, the aircraft may be any type of aircraft, for example a slightly flexible or slightly elastic, a moderately flexible or moderately elastic or a highly flexible or highly elastic aircraft. In particular, a highly flexible (elastic) aircraft may be an aircraft that can no longer be described sufficiently accurately using linear considerations. In the case of a flexible aircraft, the acceleration of the aircraft may in particular be the acceleration of the elastic aircraft structure at the point at which the aeroelastic oscillation form (mode shape) has an extreme value or a node.

[0053] The method for controlling and regulating an actuation system of an aircraft may include receiving third input data describing a controlled variable of the actuation system at a third input interface of the device, wherein determining the controlled variable taking into account a reference variable and a controlled variable comprises: determining the actuation system reference variable taking into account the reference variable and the controlled variable and determining the manipulated variable taking into account the actuation system reference variable and the actuation system controlled variable.

[0054] The above-explained designs for the device for regulating and controlling an actuation system of an aircraft can be correspondingly provided with respect to the device and / or the method, and vice versa. In particular, the device can be configured to perform the functions of the components explained in conjunction with the installation, and the installation can have components configured to provide the functions explained in cooperation with the device.

[0055] According to the present disclosure, a device for regulating and controlling an actuation system of a vehicle is provided, the device being constructed with: a first input interface configured to receive first input data describing a reference variable; a second input interface configured to receive second input data describing a controlled variable; and a control output terminal configured to output a control signal describing a manipulated variable for the actuation system of the vehicle, the vehicle being controlled by means of the actuation system. The reference variable describes a target acceleration at a certain point of the vehicle, the vehicle being controlled by means of the actuation system, and the controlled variable describes an actual acceleration of the vehicle at that point. The device is configured to determine a manipulated variable from a difference between the reference variable and the controlled variable, and output a control signal corresponding to the manipulated variable via the control output terminal. With respect to the device for regulating and controlling an actuation system of a vehicle, it can be configured accordingly with reference to the implementation scheme of the device for regulating and controlling an actuation system of an aircraft. In particular, according to the present disclosure, a corresponding facility for regulating and controlling an actuation system of a vehicle is provided, the facility having a vehicle comprising an actuation system and an acceleration sensor, a vehicle control device having an output interface, and a device for regulating and controlling the actuation system of the vehicle. In addition, according to the present disclosure, a corresponding method for regulating and controlling the actuation system of a vehicle is provided. In particular, the vehicle can be an aircraft, such as an aircraft, a helicopter or an airship. Alternatively, the vehicle can be, for example, a watercraft or a spacecraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0056] In the following, further embodiments will be explained in more detail with reference to the accompanying drawings, wherein:

[0057] Figure 1 Shows known devices for regulating and controlling the actuation systems of an aircraft;

[0058] Figure 2An apparatus for regulating and controlling an actuation system of an aircraft according to the present disclosure is shown;

[0059] Figure 3 The arrangement structure of the acceleration sensor on the lifting tail of the aircraft is shown;

[0060] Figure 4 Shows additional facilities for regulating and controlling the actuation system of the aircraft;

[0061] Figure 5 A schematic diagram showing a concept for acceleration-based roll attitude adjustment of an aircraft;

[0062] Figure 6 A schematic diagram showing a concept for acceleration-based regulation of a mechanical system;

[0063] Figure 7 A schematic diagram showing a concept for acceleration-based regulation of an elastic aircraft;

[0064] Figures 8A-8E shows overall system dynamics for known and in accordance with the present disclosure adjustments to the actuation system of an aircraft;

[0065] Fig. 9 Bode plots showing embodiments of previously known and presently disclosed actuation system adjustments for an aircraft;

[0066] Fig.10 A schematic diagram showing the arrangement on a flexible aircraft; and

[0067] Fig.11 A schematic diagram showing an alternative arrangement on a flexible aircraft. DETAILED DESCRIPTION

[0068] Figure 1The device for regulating and controlling the actuation system of an aircraft according to known methods is shown. Here, a flight control device 1 of the aircraft as an automatic control system (which can also be referred to as a flight control system) receives a measured variable 2 for describing the motion state of the aircraft. According to the motion state of the aircraft, the flight control device 1 determines a reference variable 3 for regulating the actuation system 4. The actuation system 4 is constructed with an actuation mechanism 4a and a force generator 4b, wherein the actuation system can include other components. The force generator 4b can be, in particular, an adjustment surface of the tail wing and, if necessary, also a stationary component of the actuation mechanism that participates in the force generation. Alternatively, the force generator 4b can be, for example, a nozzle or a propeller. The actuation mechanism 4a is used to influence the force generator 4b so that the desired force is applied to the regulated system, i.e., the aircraft. In the design scheme in which the actuation system 4 is a tail wing, the actuation mechanism 4a can be a servomotor, which pivots the adjustment surface of the tail wing as a part of the force generator 4b or the force generator relative to the immovable part of the tail wing. In an alternative embodiment, the actuator 4 a can be, for example, a valve of a nozzle or a drive motor of a propeller used as the force generator 4 a .

[0069] Reference variable 3 describes the target control position of the actuating mechanism 4a of the actuating system 4, i.e., for example, the rotational position of a servomotor (corresponding to the position of the adjustment surface), the opening state of a valve of a nozzle, or the drive position of a propeller drive motor from which the propeller speed is obtained, or the drive position of a servomotor for blade angle adjustment.

[0070] The regulating device 5 of the installation receives the reference variable 3 via a corresponding input interface. The regulating device also receives a controlled variable 6 via a corresponding further input interface, which controls the actual manipulated position of the actuating system. The regulating device determines a control deviation as the difference between a target value of the manipulated position according to the reference variable 3 and an actual value of the manipulated position according to the manipulated variable 6. In the regulating device 5, an actuating system reference variable is determined from the control deviation by multiplication with a proportionality factor, which is compared with the actuating system controlled variable 7 in order to determine a manipulated variable 8 of the actuating system. The manipulated variable 8 can be, for example, an actuator voltage or an actuator current.

[0071] Based on the manipulated variable 8, the actuator 4a causes the position of the force generator 4b. This results in a force and / or torque action 9 on the mechanical system 10, i.e., the aircraft. Figure 1 and Figure 2 The aircraft is shown as a mechanical system 10 separately from the remaining components, but the actuation system 4 and, in an advantageous embodiment, also the flight control device 1 and the control device 5 form part of the aircraft.

[0072] In addition to the desired force and / or torque effects 9, there are also disturbing forces and / or disturbing moments 11 acting on the mechanical system 10 of the aircraft caused by external influences, such as wind effects, in particular in the form of gusts. Figure 1 It can be seen from the diagram that the disturbing forces and / or disturbing moments 11 acting on the aircraft can only be compensated by the flight control device 1 if the measured variable 2 is taken into account, which includes the influence of the disturbing forces and / or disturbing moments 11. Therefore, this consideration is only carried out within the scope of the flight control, which is usually slow compared to the actuation system control (servo control).

[0073] Figure 2 Now, a device for regulating and controlling an actuation system of an aircraft according to the present disclosure is shown. Figure 1 Compared to the device of the invention, according to the present disclosure, a device 12 for regulating and controlling the actuation system 4 is provided, which is set up to receive at a first input interface a reference variable 13 which describes a target acceleration at a certain point of the aircraft. To this end, the flight control device 1 is set up to determine the reference variable 13 which describes the target acceleration from the measured variables 2 and to send it to the device 12. At a second input interface, the device 12 receives a controlled variable 14 which describes the actual acceleration at this point of the aircraft.

[0074] In particular, the acceleration of the aircraft may be a local acceleration at the actuation system 4 . Figure 3 The arrangement of an acceleration sensor 15 on an immovable part of an aircraft's tail 16 is shown by way of example. The tail 16 is an actuation system 4 of the aircraft, in which a force is exerted on the aircraft by moving an adjustment surface of the tail as a force generator 4a by means of a servomotor as an actuation element 4a, which causes a pitching action and thus causes the aircraft to climb or descend.

[0075] Alternatively, the acceleration may be the acceleration at another point of the aircraft, for example at the center of gravity of the aircraft. The acceleration may be measured directly using an acceleration sensor or may be determined from one or more measurements, which may be the acceleration at one or more other points or may be a variable other than the acceleration, for example a vertical movement (change in positioning) of a wing.

[0076] The control device 12 determines a control deviation as a difference between a target value of the acceleration according to the reference variable 13 and an actual value of the acceleration according to the controlled variable 14. In the control device 12, an actuation system reference variable is determined from the control deviation by multiplication with a proportionality factor, which is compared with the actuation system controlled variable 7 in order to determine the actuation system manipulated variable 8. Based on the manipulated variable 8, the actuation mechanism 4a causes the position of the force generator 4b, which results in a force action 9 on the mechanical system 10, i.e. the aircraft.

[0077] In an alternative design, instead of such a cascade structure, a parallel feedback can be provided, in which the controlled variables 7 and 14 are fed back and modified, in particular multiplied and / or integrated, respectively, at this time. Here, the reference variable 13 is modified according to the controlled variables 7 and 14 by means of a prefilter, and then the manipulated variable 8 is determined by adding the controlled variables 7, 14 and the reference variable 13.

[0078] As from Figure 2 It can be seen from FIG. 1 that the controlled variable 14 is a variable of the mechanical system 10, i.e., the aircraft. The disturbing forces and / or disturbing moments 11 influence the acceleration of the aircraft, so that the acceleration of the aircraft described by the controlled variable 14 already at least partially includes these influences. Figure 2 In the adjustment concept shown in Figure 1 Compared to the known concept illustrated in FIG. 1 , the disturbing forces and / or disturbing moments 11 acting on the aircraft are already taken into account in the regulation of the actuation system 4 of the aircraft.

[0079] The manipulated variable 8 can be, in particular, an actuator voltage or an actuator current. For example, the actuation system reference variable can be a target value for the actuation speed of the actuation mechanism 4a, i.e., in particular an actuator. In this case, the actuation system controlled variable 7 can be the actual actuation speed of the actuation mechanism 4a. In an exemplary design, the target value for the actuator current is determined by the difference between the actuation system reference variable and the actuation system controlled variable. The target value for 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 target value for the actuator current.

[0080] Figure 4 The embodiment of such a device for regulating and controlling an aircraft actuation system is shown, wherein an additional internal regulating loop is provided. Here, a pitch regulation using a rotating electromagnetic actuator of the elevator tail is shown by way of example. Compared to the known regulation of the adjustment surface amplitude n with feedback, the local acceleration b at the elevator tail is fed back. zH , and determines the relative predefined acceleration b within the actuator control 17 zH,c The adjustment deviation. When using the coefficient KbzH In the case of The control speed command is set by the internal speed regulation loop. c (actuator reference variable) relative speed error Control speed command (actuation system reference variable) and the actual operating speed as the actuation system controlled variable is proportional to the difference between Finally, the motor terminal voltage U forms the manipulated variable. The manipulated variable and the regulation error I c -I is proportional (coefficient K I Here, the actual current I is the actuator controlled variable.

[0081] Within the scope of the physical processes within the actuator, according to the modeling as a DC shunt, the terminal voltage causes a change in the current flowing in the motor winding, which is inversely proportional to its inductance L. However, the voltage drop ΔU due to the winding resistance R must be taken into account. res =R·I and motor constant K e The reverse voltage induced by the rotational motion is proportional to They reduce the terminal voltage. The through current I is obtained by integrating the current change and contributes to the motor constant K t Proportional to the driving torque M act .

[0082] Regarding the physical effects on the actuation system, in addition to the drive torque M act In addition, the aerodynamic control surface hinge moment M aero Also acting on the adjustment surface, the aerodynamic control surface hinge moment includes the coefficient C n，aero The component proportional to the swing amplitude n and the damping (coefficient ) component. In addition, the aerodynamic control surface hinge moment M aero Affected by the airflow direction (coefficient ). The resulting total torque results in an adjusted acceleration that scales proportionally to the inverse of the rotational inertia 1 / J

[0083] exist Figure 4 In the right part of FIG. 1 , the dynamics of the pitching motion 18 of the aircraft are shown in a simplified manner. With the reverse pitch inertia 1 / l yy Proportional to the pitching moment, which is achieved by utilizing the dynamic pressure Wing area S and wing depth I μThe denormalization is obtained from the pitch moment coefficient. This pitch moment coefficient basically includes the elevator (C mn n), pitch rate and angle of attack (C mα ·α). The angle of attack α has an influence on the longitudinal attitude. The angle of attack also includes the wind angle of attack α W The local acceleration b at the elevator tail zH Pitch acceleration Multiply by the lever arm r H and the vertical acceleration b of the aircraft's center of gravity z Conclude.

[0084] According to the present disclosure, for example, Figure 4 As can be seen, the control position is not considered as a controlled variable. Force measurement or torque measurement is also not used as a controlled variable. In addition, the controlled variable is not measured in the drive train or adjustment surface of the actuator, but in Figure 4 In the embodiment of the invention, the acceleration is measured on the tail wing associated with the adjustment surface (which is a non-movable stabilizer relative to the aircraft). No (rotational) acceleration of the actuator is performed. The acceleration has been measured with the adjustment torque (actuator driving torque, M act ). On the contrary, as in Figure 4 It can be seen that the local acceleration at the stabilizer surface is proportional to the angle of the adjustment surface and the lift generated by it, that is, to the acceleration of the actuator. The two integration steps are proportional to the variables separated. The local acceleration is an output variable that depends largely on the control surface angle n as the system state and therefore the system dynamics can be influenced in a similar way by feeding back the control surface angle. Figure 4 The embodiment shown in , should maintain the conventional servo-regulated speed control loop ( Figure 4 The intermediate cascade in the control system is thus always fed back a number of linearly independent output variables corresponding to the system order. This makes it possible to design the system dynamics in any way. It is desirable to design the rudder angle dynamics as "rigidly" as possible. In particular, it is not necessary to aim at reducing the actuator load or the adjustment effort, but it can be provided that the adjustment surface is moved as quickly as possible to that position which compensates for the influence of the gust of wind on the corresponding tail. This position is usually different from the equilibrium position into which the free rudder is blown when the adjustment torque is kept constant.

[0085] Compared to adjusting the hinge torque of the control surface, local acceleration regulation can achieve advantages. The local acceleration measurement (in contrast to adjusting the surface angle or the hinge torque of the control surface) directly detects the additional angle of attack α W The additional lift caused by gusts. In elastic aircraft, the local acceleration directly reflects the oscillatory state of the structural dynamics. The acceleration feedback to the control speed of the adjustment surface acting in the same direction corresponds to a virtual damping (similar to the so-called ILAF principle). Therefore, it can be particularly suitable for active stabilization of highly elastic configurations. In addition, the local acceleration contains different flight state variables ( γ, q, see Figure 4 ), these flight state variables can be compensated by adjusting them together. bz,H Compared to the highly dynamic feedback branch of the control system, these influences can become insignificant, so that significantly greater robustness can be achieved with respect to the constantly changing aerodynamic characteristics. The local acceleration can be derived from the planned flight path and flight attitude trajectory in a direct, purely kinematic relationship. This makes it possible to deduce simple pre-control procedures that are independent of the characteristics of a specific aircraft. Therefore, the high dynamics of the local acceleration control (which corresponds to the traditional positioning control loop of the servo control) can be used not only for interference suppression, but also for guiding behavior. This makes significantly more flexible flight guidance possible.

[0086] The regulation of actuators (servo control) and the regulation of flight conditions (flight control) are traditionally separate research disciplines that are handled in different circles of experts. Feedback of local accelerations measured on the aircraft structure in an internal control loop that is traditionally part of the servo control creates a bridge between these two areas. This requires a comprehensive view of the entire control path, which considers the aircraft and its control elements as a unit. By using local accelerations as predefined variables, the part of the flight dynamics can be incorporated into the control path of the servo control. This makes it possible to simplify the control path of the flight control and reduce the dependence on special flight characteristics, so that traditional flight control structures are no longer suitable.

[0087] according to Figure 4 , at the local acceleration b zH And the adjustment surface swing n draws the boundary of the actuation system adjustment 17. Other representations are also possible, in which the definition of the subsystem, especially the boundary, is defined in different ways (see, for example, Figure 5 ), without thereby resulting in a change in the regulation principle according to the present disclosure.

[0088] Figure 5 , Figure 6 and Figure 7 The symbols used in represent the following variables:

[0089] Scalar:

[0090] C lβ : Sliding rolling moment

[0091] C lξ : Aileron effectiveness

[0092] C lP : Roll Damping

[0093] I: Actuator current

[0094] l yy : Rolling moment of inertia of the aircraft

[0095] J: Moment of inertia of the actuator

[0096] K t : Torque constant of the actuator

[0097] K…: Regulation gain of the regulation loop

[0098] S: wing area

[0099] V A : Flight speed

[0100] Dynamic Pressure

[0101] b: half span

[0102] p: roll rate

[0103] β: sliding angle

[0104] β w : Wind slip angle

[0105] ω: angular velocity of the actuator

[0106] ξ: Aileron swing

[0107] vector:

[0108] η : Modal amplitude (degrees of freedom for structural dynamics)

[0109] R : The orientation vector to the local acceleration measurement point

[0110] q : Generalized coordinates

[0111] u : Manipulated variables

[0112] x : Rigid body degrees of freedom

[0113] z : Interference variable

[0114] Matrices and Tensors:

[0115] Effect of Generalized Force Adjustment on Structural Dynamics Degrees of Freedom

[0116] Effect of Generalized Force Adjustment on Rigid Body Degrees of Freedom

[0117] Adjustment of the general force

[0118] Generalized stiffness matrix

[0119] Generalized Damping Matrix

[0120] Disturbance Effects of Generalized Forces on Structural Dynamics Degrees of Freedom

[0121] Disturbance effects of generalized forces on rigid body degrees of freedom

[0122] Interference effects on generalized forces

[0123] Effects of aerodynamic forces on rigid body motion due to structural deformation

[0124] Gain matrix of the regulation loop

[0125] Kinematic transformation ratio between the generalized rigid body degrees of freedom and the local acceleration measurement point positioning

[0126] M: generalized inertia matrix

[0127] Effects of aerodynamic forces on structural dynamics due to structural deformation

[0128] Effects of aerodynamic forces associated with rigid body motion on structural dynamics

[0129] Δ : Vibration shapes (eigenvectors) of structural dynamics.

[0130] Generalized structural damping coefficient

[0131] Generalized stiffness matrix

[0132] Modal Mass Matrix

[0133] Subscript:

[0134] c: command variable, preset value, target value

[0135] In conventional flight control, the command corresponds to the positioning (angle) of the aerodynamic adjustment surface. The highly dynamic (strict) positioning control of the actuator ensures that the actual adjustment surface positioning accurately follows the adjustment command. The control structure corresponds to a cascade control, which has an inner control loop, i.e., the actuator control (ACL), and an outer control loop, i.e., the flight control (FCL). Feedback is performed on the attitude angle, the rotation rate and the speed. Acceleration measurement is usually only used for observation or as a substitute for states with poor measurability.

[0136] Flight control based on the hinge torque of the control surface is also known. The command of the FCL corresponds to the torque presetting for the actuator, that is to say the direct presetting current. In the equilibrium state, the torque presetting corresponds to the aerodynamic hinge torque of the control surface. This concept is similar to the force-oriented control behavior of the pilot during manual control. Such a regulation should provide advantages for flight stability and load reduction, because the control surface deviates due to the change of the hinge torque of the gust. This should reduce the load on the actuator and avoid the occurrence of force struggle in the case of redundant actuators.

[0137] In the previously known incremental nonlinear inversion (INDI), a local linearization and inversion of the system dynamics are performed. The incremental growth of the adjustment command is calculated. The method is based on the measured and commanded (rotational) acceleration and reduces the influence of the (aerodynamic) model accuracy and the center of gravity position for improving stability. The adjustment procedure is based on the comparison between the planned and actual changes (therefore derivatives) of state variables, which are calculated or observed based on rotational and translational acceleration measurements. Unlike the concept according to the present disclosure, it is unknown to use this manipulated variable change directly in the internal cascade of the servo control or to extend the INDI method to the actuator dynamics. In the proposed method, the actuator current is used as a predetermined variable and the adjustment procedure corrected for this is inferred. Here, unlike the method according to the present disclosure, the quasi-static dependence of the actuator current on the control surface hinge torque is used as the basis, so that the dynamics of the actuation system itself remain unregulated.

[0138] For active vibration control and load reduction, it is known to feed back acceleration measurements or modal degrees of freedom. In this case, the command corresponds to the positioning of the adjustment surface. Alternatively, additional forces are also applied by vibration actuators. This usually does not take place in the sense of control, but rather in a targeted manner to compensate for individual resonance frequencies.

[0139] In known systems, the dynamics of the flight controller (bandwidth) decisively determine the accuracy of the course and attitude keeping (disturbance suppression), the flight stability (disturbance suppression) and the flexibility of the course guidance (guidance behavior). The maximum available bandwidth is limited by the dynamics of the independently designed actuator regulation (internal control loop) and, if necessary, by the dynamics of the mechanical transmission path between the actuator and the adjustment surface, the structural dynamics of the elastic aircraft and the unstable aerodynamics. For an optimal FCL design, an accurate aerodynamic model is required. This is expensive and may lack robustness. The internal control loop, at least the attitude control, must be designed individually for each aircraft type. An accurate aeroelastic model is required in order to exclude excitations from structural dynamics. Flap swing amplitude η k Vertical load multiple (i.e. load acceleration) n z The direct influence of the gust load regulation makes the design difficult. Without the provision of a pre-control coupled with complex angle of attack measurements, the reduction potential is limited.

[0140] Figure 5 A schematic diagram of a concept for acceleration-based roll attitude control of an aircraft is shown. Compared to known systems, the positioning control of the actuator is replaced by feedback from acceleration measurements, which detect the aerodynamic force effects on the adjustment surface (e.g., local acceleration on the adjustment surface or the rotational acceleration of the aircraft). Here, the traditional division between actuator control and flight control is changed. The interface between FCL and actuator control slides inwards by a cascade. The state feedback of the actuator swing amplitude is replaced by output feedback of an acceleration proportional to it, which additionally contains interference effects (gusts). The command of the FCL then corresponds to the adjustment rate (angular velocity) of the aerodynamic adjustment surface. It is only necessary to measure the adjustment surface positioning in order to take into account the adjustment limits.

[0141] In the control path of the aircraft, that is, FCL, the rolling moment coefficient C l C lξ It is proportional to the aileron swing amplitude ξ, which is the manipulated variable in the known system. Rolling moment coefficient C l C lβ is proportional to the sliding angle β, which for pure roll control is considered as a disturbance variable and includes in particular the wind influence β w . Rolling moment coefficient C l C lp and the dimensionless roll rate p* = p·b / V A The rolling moment is proportional to the coefficient C l Multiply by the reference variable ( S, b) shows that the rolling acceleration is the opposite of the rolling inertia (1 / l yy ). The roll rate p and the roll angle Φ are obtained by integrating the roll acceleration. The known flight control includes the complete feedback of the states "roll rate (p)" and "overhang angle (Φ)". Here, the structure is realized in the form of a cascade control, in which the outer control loop includes a reference variable "roll command (Φ)" c )" and the manipulated variable "roll rate command (p c )”, the roll rate command is calculated with a gain K Φ and adjustment error Φ c -Φ. Thus, the inner regulation loop involves a reference variable “roll rate command (p c )” and the roll rate of the manipulated variable “aileron command (ξc)”, which is aileron command with a gain rate K p With the adjustment error p c - Proportional to p.

[0142] In the actuator, i.e., the regulation path of the ACL, ignoring the electrical time constant, the actual current I and the current command I c The torque is expressed as the torque constant K. t Proportional to the current flowing through. Other torque contributions (friction, aerodynamic control surface hinge torque, etc.) are neglected. The angular acceleration (ω) of the driven part is obtained from the conservation of angular momentum of the torque / moment of inertia (J). The control speed (ω) and the driven part angle (which corresponds to the aileron swing amplitude ξ) are obtained by integrating the rotational acceleration. The known actuator control then includes complete feedback on the states of "control speed (ω)" and "control positioning (ξ)". Here, the construction is realized in the form of a cascade control, in which the outer control loop includes a control loop with a reference variable "aileron command (ξ)". c ) and the manipulated variable “adjustment rate command (ω c )”, the adjustment rate instruction is based on the gain K ξ and adjustment error ξ c -ξ. Thus, the inner regulation loop involves a reference variable “regulation rate command (ω c )" and the manipulated variable "current command (I c )”, the current command is based on the gain K ω and adjustment error ω c - is proportional to ω.

[0143] In contrast, Figure 5 The acceleration-controlled concept is shown with a control path 20 of the aircraft, an actuator 21 and a control path 22 of the actuator. The measured roll acceleration The FCL command ω is fed back instead of the proportional aileron swing (ξ).c Corresponding to the adjustment rate ξ. In addition to adjusting the surface positioning ξ, the feedback according to the present disclosure also directly detects the w The disturbances caused by this are thus compensated one control loop further inwards compared to known flight controls. In a high-dynamic design with feedback (for example, the FCL), a fairly good disturbance suppression can be achieved. The pre-set value of the attitude control loop (the traditional "inner loop" of the FCL) corresponds directly to the rate acceleration (the second derivative of the controlled variable). In a high-dynamic design of acceleration control, the aircraft accelerates directly at the pre-set rate. This results in a simple, linear behavior that is independent of the aircraft-specific parameters. The design of the attitude control can be standardized and can be performed independently of the aircraft type and flight conditions. In highly dynamic designs, The acceleration feedback becomes more important than the rest of the aerodynamics (via C lβ and C lp ) is more important, therefore, reducing the aerodynamic parameters (except C lξ ) on the control loop. Therefore, the precise aerodynamic model can be eliminated for the design of FCL. Only the rudder effectiveness C lZi and dynamic pressures remain important. Avoiding having adjustment commands directly affect acceleration measurements simplifies the design and increases the potential for adjustment-based gust load reduction.

[0144] For rigid aircraft, this principle can be transferred to the pitch degree of freedom (measured variable: pitch acceleration, main manipulated variable: elevator), the yaw degree of freedom (measured variable: yaw acceleration, main manipulated variable: rudder), the lift degree of freedom (measured variable: vertical acceleration n z , main manipulated variables: flaps), longitudinal degrees of freedom (measured variables: longitudinal acceleration n x , main manipulated variables: spoilers) and lateral degrees of freedom (only in the case of lateral force control). Ideally, the acceleration components are fed back not only to the main manipulated variables, but also to all manipulated variables that influence the respective degrees of freedom, for example via the aileron rolling moment, the rudder yaw moment, the elevator lift or the flap pitch moment. By suitable choice of the gain matrix, the degrees of freedom can be completely decoupled. The choice of the descriptive degrees of freedom for which the acceleration is measured is arbitrary. By way of example, the rotation and translation of the center of gravity in coordinates fixed relative to the aircraft are mentioned. Also conceivable are other coordinate systems and other (if necessary multiple) reference points for rigid bodies, such as the vertical position of the two wingtips instead of the roll angle. It is possible to unambiguously describe the system for each combination of independent degrees of freedom. This represents a set of effective generalized coordinates (q) in the sense of the Lagrangian formalism.

[0145] Figure 6 The transferability of the control concept according to the present disclosure to general mechanical systems is shown by the schematic diagram of the concept for acceleration-based control. The principle can be transferred to any mechanical system 23 with n degrees of freedom, which can be described explicitly in the sense of Lagrangian formalism by generalized coordinates, which is controlled by one or more manipulated variables, which have a direct force or torque influence on the system 23, and the manipulated variables of the system are operated by regulated actuators 24 with at least simple integration behavior (all mechanical actuators).

[0146] In accordance with Figure 6 In different methods of controlling the positioning of the system, the manipulated variable u i Has a force effect, which can be expressed as a generalized force The same applies to force or torque disturbances z i , and its action coefficient is Acceleration in generalized coordinates With the generalized force Q i M ij -1 ( q ) and is proportional. Generalized speed and coordinates q By integration, we can get the generalized speed Leading to non-conservative "damping forces" in dissipative systems Conservative forces and generalized coordinates q by ( q ) is proportional to. In the known control concept, the generalized coordinates (q) and the speed The structure is implemented in the form of cascade regulation, in which the outer regulation loop involves generalized coordinates. Its reference variable is the target value of the generalized coordinates ( q c ). The manipulated variable is the target value of the generalized speed The target value is expressed as a gain matrix With adjustment error q c - q The inner regulation loop involves the generalized speed, where the reference variable is the target value of the generalized speed and the manipulated variable is the adjustment instruction ( u c ), adjust the instruction to gain the matrix With adjustment error Directly proportional.

[0147] The actuator manipulates the variable The commanded and actual changes of have an arbitrary transfer behavior G(s), but have at least one integration level. The actuator regulation includes (at least) the manipulated variable u The reference variable is the target value of the system manipulated variable. u c The actuator manipulated variable is the target value of the system regulation rate It gains With adjustment error u c - u Directly proportional.

[0148] Compared with the known regulation, Figure 6 The acceleration-adjusted method is used to measure the generalized acceleration feedback, rather than a manipulated variable to which feedback is proportional u . Adjustment command and adjustment rate corresponding.

[0149] exist Figure 7 The middle figure shows the concept of acceleration-based regulation for elastic aircraft. Here, Figure 7 shows the general case of full state feedback. The elastic vehicle is a combination of Figure 6 The explanation of the concept according to the present disclosure is a special case, because it can be described by Lagrangian formalism. The choice of generalized coordinates is in principle arbitrary. One possibility is that the acceleration sensors distributed on the aircraft are positioned individually. In this case, the measured acceleration is directly related to the generalized acceleration Accordingly. This requires at least six sensors for detecting the rigid body motion. The number of additional sensors determines how many elastic modes can be detected. In an alternative possibility, the rigid body motion (mean axis) and the structural dynamics are separated. Here, the division into rigid body degrees of freedom and the amplitude of the elastic mode ( η =[η1,η2,…,η n ] T ),thereby q =[x i ,η1] T The equations of motion for rigid body motion and structural dynamics are inertially decoupled, but coupled given the external forces (aerodynamics).

[0150] The regulation path includes rigid body dynamics ( Figure 7 ), which is Figure 6 The system is constructed similarly, and the corresponding relationship is and About Structural Dynamics ( Figure 7 The upper part of the manipulated variable u iHas a force effect, which can be achieved through a generalized force The same applies to force or torque disturbances (z i ), and its action coefficient is Modal Amplitude The second-order derivative of the inverse modal mass matrix μ -1 (ij) and the generalized force Q j Structural damping induces damping force, which is proportional to the damping coefficient The speed of change of modal amplitude The structural elasticity induces the spring force, which is expressed as a generalized stiffness matrix Proportional to the modal amplitude.

[0151] The coupling between the structural motion and the rigid body motion is provided by external forces. Here, the external forces (aerodynamic forces / torques) depend on the rigid body states and x and structural dynamics and External forces not only affect the motion of rigid bodies It also affects the structural dynamics The influence of some forces on the rigid body motion has been determined by The influence of some forces caused by structural deformation on structural dynamics ( and )(Unlike it) has not yet been included in The impact on the structural dynamics is provided via the correlation of forces and rigid body motions: The outer regulation loop involves the generalized coordinates ( q =[x i , η i ] T ), the internal regulation loop involves local degrees of freedom ( R f ). The system behavior depends on the description form (conversion between system / state representations with different degrees of freedom).

[0152] according to Figure 7 A special case of the concept involves local acceleration feedback. Here, the acceleration measurement is carried out directly at the location of the adjustment surface. Feedback matrix K ROnly the diagonal is occupied, that is, the acceleration affects only the control surface where the measurement is carried out. As long as the number of control / measurement positions is equal to the number of (considered) degrees of freedom (and is appropriately selected, that is, linearly independent; thus resulting in controllability and observability), the system dynamics can be freely predetermined. However, a complete self-structural predetermined is not possible. The system remains coupled. Assuming that the control rate command is implemented without delay (ignoring the actuator dynamics), the feedback of acceleration to the control rate is equivalent to the feedback of velocity to the control position. Assuming that the aerodynamic force generated by the control surface is realized without delay (ignoring unsteady aerodynamics), a reaction force proportional to the velocity is generated. Therefore, the acceleration feedback corresponds to a virtual viscous damper acting at the location of the control surface. By introducing the integral part or feedback of the local velocity measurement, a virtual spring element can be similarly embodied, resulting in a more or less rigid clamping of the airfoil at the location of the control surface. The damping force always acts opposite to the direction of movement, which excludes modal instabilities of the energy transfer and thus the structural dynamics. However, this only applies if the assumption is reasonable, i.e. for all structural modes with frequencies significantly lower than the actuator dynamics / unsteady aerodynamics. This limits the maximum achievable dynamics of the acceleration feedback.

[0153] In contrast, in the case of acceleration measurements that are locally separated from the control surface (e.g., an IMU in the cockpit), there is a risk of excitation, since the acceleration signal reacts to the forces generated on the control surface only with a delay due to the structural dynamics. In other words, there can be vibration modes whose antinodes of oscillation at the location of the measurement have an opposite sign to the location where the forces are generated (control surface). In the frequency range of this vibration mode, negative, destabilizing damping forces are thus obtained. This is excluded if the measurement is carried out directly at the location where the forces are generated.

[0154] In conventional flight control, the bandwidth is limited by the frequency separation for structural dynamics, whereas in the case of local acceleration feedback, the bandwidth is limited by the frequency separation for unsteady aerodynamics and actuator dynamics.

[0155] The advantages of the acceleration-controlled concept can be especially that higher dynamics of flight control can be achieved and thus better interference suppression, improved flight stability and flexibility can be achieved, especially in the case of highly elastic aircraft, where there is no need to perform frequency separation of structural dynamics or filtering of elastic modes. It is possible to make all elastic modes automatically damped below the actuator dynamics and especially the aerodynamics, regardless of the specific elastic properties of the aircraft. Compensation for interfering influences (local gusts) is carried out directly at the point of action without exciting structural dynamics (similar to birds, which only spread their feathers locally to succumb to gusts). Structural loads are reduced. Acceleration measurement and regulation can be integrated into the actuator control unit (intelligent actuator). This enables decentralized system construction. The acceleration-controlled concept enables novel redundancy concepts and allows simple adaptation of the control procedures in the case of failure of the adjustment surface. In the case of sufficient frequency separation between acceleration regulation (distributed in the actuator) and attitude regulation (centralized in the FCC), attitude regulation and all upper-level control loops can remain unchanged, because even if the dynamics of acceleration regulation are reduced, it is still fast enough. This corresponds to the fact that in conventional flight control, redundant actuator failures for the same adjustment surface generally do not require adaptation of the FCL.

[0156] Figures 8A to 8E The overall system dynamics in a complex plane of an exemplary embodiment of the facility are shown, which can be represented by Figure 5 The structure shown in FIG is transferred to the pitch degree of freedom. Here, Figure 8B Shows Fig. 8A The enlarged section of Figure 8C Shows Figure 8B The enlarged section of Fig.8D yes Figure 8C The enlarged section of Fig. 8E yes Fig.8D A magnified section of . Figures 8A to 8EThe diagram is based on a device for adjusting a tailplane according to the present disclosure. Longitudinal attitude θ and pitch rate q are fed back. Shown is a distribution of pole positions and zero positions, which are obtained for the different feedback gain embodiments shown (+-shaped marks). For comparison, the pole positions and zero position distributions obtained for a previously known adjustment are additionally shown, wherein the adjustment angle of the tailplane is adjusted by the actuator system adjustment (x-shaped marks). The larger x- or + marks shown represent the pole positions obtained when the values ​​of the two feedback gains (for q and θ) are assumed to be zero. For the previously known adjustment (x), this pole distribution corresponds to the style known for an unregulated aircraft, wherein a complex conjugate pole pair is assigned to the longitudinal oscillation movement and a complex conjugate pole pair is assigned to the angle of attack oscillation. By feeding back the local acceleration within the scope of the embodiment according to the present disclosure, the angle of attack oscillation is strongly damped and split into two real poles. The longitudinal oscillation pole is hardly affected by this.

[0157] The smaller marks connected by lines indicate the pole positions which can be obtained with a simultaneous increase in the feedback gains for pitch rate and longitudinal attitude at a constant ratio. The star marks indicate the pole positions which can be obtained for a favorable choice of feedback gain when striving to adjust a high bandwidth of the loop with a damping degree not falling below a value of 0.7.

[0158] Fig. 9 The feedback gain can be advantageously selected for the previously known embodiment 25 and for the embodiment according to the present disclosure. Figures 8A to 8E Bode diagram for embodiment 26. Shown is the vertical wind speed w Wg Frequency response of the disturbance transfer function of (gust) to longitudinal attitude θ. In the frequency range below the dynamics of the actuation system regulation, the embodiment 26 (solid line) according to the present disclosure shows improved disturbance behavior, since the transfer function of the previously known device 25 (dashed line) has an additional zero position that almost compensates for the pole associated with the positioning regulation loop of the servo regulation. This zero position is eliminated by the acceleration feedback.

[0159] Fig.10 An exemplary embodiment of a flexible aircraft is shown. Without limiting the generality, Fig.11The illustration of here limits the flexibility of the main wing with respect to bending about the longitudinal axis, torsion about 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, in addition to the rigid body degrees of freedom of the undeformed reference configuration 27, includes the amplitude of the superimposed vibration mode of the elastic mode, which describes the characteristic deformation pattern compared to the reference configuration. Such vibration modes are exemplarily shown in Fig.10 . Here, the vibration mode can have a local extreme point 28 with the maximum deviation from the reference configuration and a node 29 with zero deviation from the reference configuration. In order to control and stabilize the deformation degrees of freedom, the flexible aircraft can have multiple actuation systems (tail), which can be formed by adjustment surfaces 30 and, if necessary, stabilizers 31. The device according to the present disclosure can be used to adjust one or more actual accelerations at any point 32 of the aircraft. In an advantageous embodiment, this can be, for example, an extreme point or node of one or more vibration modes, but alternatively it can also be a point at which no vibration mode relevant to the adjustment has a node. In particular, the number of accelerations used can correspond to the number of eigenvalues ​​that should be stabilized or influenced by the adjustment. Here, for example, all accelerations can be received by each of the provided devices, for example. Alternatively, a device can only receive those accelerations that can be affected by the adjustment of the actuation system adjusted by the device.

[0160] exist Fig.11 , an embodiment of a device for a flexible aircraft is shown, in which local accelerations are used to adjust the actuation system at a point 32 of the aircraft located near the adjustment surface 30. Here, all accelerations can again be received by all devices according to the present disclosure, for example. Alternatively, each device can respectively only receive the accelerations in the vicinity of the actuation system adjusted by the device.

[0161] The features disclosed in the above description and in the drawings can be essential for the realization of the various embodiments both individually and in any combination.

Claims

1. A device for regulating and controlling an actuation system of an aircraft, the device comprising - a first input interface, which is configured to receive first input data describing a reference variable; a second input interface configured to receive second input data describing a controlled variable; and a control output which is designed to output a control signal which indicates a manipulated variable for an actuation system of the aircraft, by means of which the aircraft is controlled, in, - the reference variable specifies the target acceleration at the point of the aircraft controlled by means of the actuation system; - said controlled variable describes the actual acceleration of said aircraft at said point; and The device is set up to determine the manipulated variable taking into account the reference variable and the controlled variable and to output a control signal corresponding to the manipulated variable via the control output.

2. The device according to claim 1, wherein the device has a third input interface, the third input interface being configured to receive third input data describing a controlled variable of the actuation system, wherein: The device is set up as - determining an actuation system reference variable taking into account said reference variable and said controlled variable, and - determining the manipulated variable taking into account the actuation system reference variable and the actuation system controlled variable.

3. The device according to claim 2, wherein: The actuation system reference variable is a target manipulation speed of the actuation system, and the actuation system controlled variable is an actual manipulation speed of the actuation system.

4. The device according to any one of claims 1 to 3, wherein: The device is set up to determine the manipulated variable without taking into account an actual manipulated position of the actuation system and without determining a target manipulated position of the actuation system.

5. The device according to any one of the preceding claims 1 to 3, comprising a further input interface which is configured to receive further input data describing further controlled variables, wherein: - the further controlled variable describes the actual acceleration of the aircraft at a further point; and The device is set up to adjust the controlled variable taking into account the further controlled variable and subsequently to determine the manipulated variable taking into account the reference variable and the controlled variable.

6. The device according to any one of the preceding claims 1 to 3, wherein: The device is set up to determine the manipulated variable from a difference between the reference variable and the manipulated variable taking into account the reference variable and the manipulated variable.

7. A device for regulating and controlling the actuation system of an aircraft, the device having - an aircraft having - an actuation system configured to control said aircraft in at least one degree of freedom, and - acceleration sensors arranged at points of said aircraft; -Flight control equipment with output interface; and - an apparatus according to any one of the preceding claims, in, the flight control device is configured to calculate the reference variable describing the target acceleration at the point of the aircraft from the flight state of the aircraft and to send first input data describing the reference variable to the first input interface of the device via the output interface; - the acceleration sensor is configured to measure a local acceleration of the aircraft at the point and to send second input data describing a controlled variable to a second input interface of the device, the second input data describing the local acceleration at the point; and The actuation system is designed to receive a manipulated variable from a control output of the device and to carry out a regulating movement corresponding to the manipulated variable.

8. The facility of claim 7, wherein: The flight control device is designed to calculate the reference variable taking into account control variables that are kinematically determined directly from the target trajectory of the aircraft.

9. A facility according to claim 7 or 8, wherein: The actuation system is formed with an actuator which moves an adjustment surface of a tail wing of the aircraft.

10. The facility of claim 9, wherein: The acceleration sensor is arranged on a part of the tail wing that is immovable relative to the aircraft.

11. The device according to any one of claims 7 to 8, comprising a further acceleration sensor, which is arranged at a further point of the aircraft, wherein: - the device is a device according to claim 5; and The further acceleration sensor is designed to measure the local acceleration at the further point and to send further input data describing a further controlled variable to a further input interface of the device, the further input data describing the local acceleration at the further point.

12. The plant according to any one of claims 7 to 8, comprising a further device according to any one of claims 1 to 5, wherein: the aircraft has a further actuation system which is set up to control the aircraft in at least one degree of freedom or in at least one further degree of freedom, and the aircraft has an additional acceleration sensor which is arranged at an additional point of the aircraft, in, The flight control device is configured to additionally send first input data specifying the reference variable to the further device via the output interface; - the additional acceleration sensor is configured to measure a local acceleration of the aircraft at the additional point and to send second input data describing an additional controlled variable to the further device, the second input data describing the local acceleration at the additional point; and The further actuation system is designed to receive the manipulated variable from a control output of the further device and to carry out a control movement corresponding to the manipulated variable.

13. A method for regulating and controlling an actuation system of an aircraft, the method comprising the following steps: - providing equipment for regulating and controlling the actuation system of said aircraft; - generating first input data describing the reference variable, wherein, The reference variable describes a target acceleration at a point on the aircraft, which is controlled by means of the actuation system; - generating second input data describing a controlled variable, wherein the controlled variable describes an actual acceleration of the aircraft at the point; - receiving said first input data at a first input interface of said device; - receiving said second input data at a second input interface of said device; - determining a manipulated variable for an actuation system of the aircraft taking into account the reference variable and the controlled variable; and - outputting a control signal indicating the manipulated variable via a control output of the device.

14. The method according to claim 13, comprising receiving third input data describing a controlled variable of the actuation system at a third input interface of the device, wherein: Determining the manipulated variable taking into account the reference variable and the controlled variable includes determining an actuation system reference variable taking into account the reference variable and the controlled variable and determining the manipulated variable taking into account the actuation system reference variable and the actuation system controlled variable.

15. The method according to claim 13, in, A manipulated variable for an actuation system of the aircraft is determined from a difference between the reference variable and the controlled variable taking into account the reference variable and the controlled variable.

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