Tandem Chain Control Allocation Method for Transition State of Tiltrotor Aircraft Based on Angle of Attack Dynamics

Abstract

The present application discloses a control allocation method for a tilt-rotor aircraft transition state series connection chain based on angle of attack dynamics, including: constructing a tilt-rotor aircraft transition state model to obtain the real-time angle of attack of the tilt-rotor aircraft; comparing the real-time angle of attack of the tilt-rotor aircraft with the stall angle of attack, dynamically allocating control authority between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result, and obtaining the optimal allocation parameters of the tilt-rotor aircraft according to the allocation result of the control authority, where the optimal allocation parameters at least include aileron deflection, rudder deflection, elevator deflection, rotor tilt angle, and rotor speed; inputting the optimal allocation parameters into the tilt-rotor aircraft transition state model to adjust the tilt-rotor aircraft. The present application can solve the control allocation problem in the transition state of the tilt-rotor aircraft under the flight environment of large angle of attack and low dynamic pressure by comparing the real-time angle of attack with the stall angle of attack, thereby improving the cruise speed and control efficiency of the tilt-rotor aircraft.

Description

Technical Field

[0001] This application relates to the technical field of tiltrotor aircraft flight dynamics and control, and particularly to a control allocation method for the transition state series connection chain of a tiltrotor aircraft based on angle of attack dynamics. Background Art

[0002] A tiltrotor aircraft is a multi-purpose and highly flexible aircraft that combines the high speed and long range of a fixed-wing aircraft with the vertical takeoff and landing and hovering in place of a helicopter. More importantly, the tiltrotor aircraft is one of the important ways to develop high-speed helicopters.

[0003] In the prior art, the transition from the helicopter mode to the fixed-wing mode of a tiltrotor aircraft involves the modeling of transition state flight dynamics and the design of a control system. Most of the current modeling and control only cover the linear model under small-angle aerodynamic disturbances and do not deeply explore the influence of the flight state at large angles of attack and low dynamic pressure on the aerodynamic coefficients and the design of the control system. Summary of the Invention

[0004] The purpose of this application is to provide a control allocation method for the transition state series connection chain of a tiltrotor aircraft based on angle of attack dynamics to solve the technical problem in the prior art that the control authority allocation in the transition state of a tiltrotor aircraft is improper and it is difficult to achieve the expected flight performance. The many technical effects that can be produced by the preferred technical solutions provided in this application are described in detail below.

[0005] To achieve the above purpose, this application provides the following technical solutions:

[0006] In a first aspect, this application provides a control allocation method for the transition state series connection chain of a tiltrotor aircraft based on angle of attack dynamics, including: constructing a transition state model of the tiltrotor aircraft, and obtaining the real-time angle of attack of the tiltrotor aircraft; comparing the real-time angle of attack of the tiltrotor aircraft with the stall angle of attack, dynamically allocating control authority between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result, and obtaining the optimal allocation parameters of the tiltrotor aircraft according to the allocation result of the control authority, where the optimal allocation parameters at least include aileron deflection, rudder deflection, elevator deflection, rotor tilt angle, and rotor speed; inputting the optimal allocation parameters into the transition state model of the tiltrotor aircraft to adjust the tiltrotor aircraft.

[0007] In some embodiments, comparing the real-time angle of attack of the tiltrotor with the stall angle of attack, and dynamically allocating control authority between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result, includes: if the real-time angle of attack is greater than the stall angle of attack, allocating the control authority to the rotor propulsion system, and if the rotor propulsion system can provide the force or moment required for the transition state, ending the control allocation task; if the rotor propulsion system reaches the position saturation or rate saturation state and still does not meet the force or moment required for the transition state, then allocating the remaining force or moment required for the transition state to the fixed-wing aerodynamic control surface system.

[0008] In some embodiments, comparing the real-time angle of attack of the tiltrotor with the stall angle of attack, and dynamically allocating control authority between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result, includes: if the real-time angle of attack is less than the stall angle of attack, allocating the control authority to the fixed-wing aerodynamic control surface system, and if the fixed-wing aerodynamic control surface system can provide the force or moment required for the transition state, ending the control allocation task; if the fixed-wing aerodynamic control surface system reaches the position saturation or rate saturation state and still does not meet the force or moment required for the transition state, then allocating the remaining force or moment required for the transition state to the rotor propulsion system.

[0009] In some embodiments, the method further includes: designing a non-linear washout function for the wind tunnel data in the helicopter mode, and designing a non-linear wash-in function for the wind tunnel data in the fixed-wing mode, to obtain the aerodynamic coefficients of the force and moment required for the transition state.

[0010] In some embodiments, the method further includes: designing a tiltrotor transition flight control law, adopting a multi-loop control structure, and establishing a controller based on non-linear dynamic inversion.

[0011] In some embodiments, constructing the tiltrotor transition state model may include: defining a body coordinate system and an airflow coordinate system; expressing the aerodynamic forces generated by each component of the tiltrotor in the airflow coordinate system, expressing the aerodynamic moments in the body coordinate system, and constructing a transformation matrix for converting from the airflow coordinate system to the body coordinate system; obtaining the resultant force and resultant moment of the tiltrotor transition state in the body coordinate system according to the resultant force and resultant moment acting at the center of gravity of the tiltrotor.

[0012] In some embodiments, in the non-linear washout function, the weight of the wind tunnel data in the helicopter mode is positively correlated with the degree of the tilt angle; in the non-linear wash-in function, the weight of the wind tunnel data in the fixed-wing mode is negatively correlated with the degree of the tilt angle.

[0013] In some embodiments, when allocating the control authority to the rotor propulsion system, if the rotor propulsion system can provide the forces or torques required for the transition state, the control allocation task is ended, which is represented by the following formula:

[0014] ,

[0015] where, is the system function of the rotor propulsion system, is the force or torque required for the transition state, is the saturation function, is the actual input control quantity of the rotor propulsion system.

[0016] In some embodiments, if the rotor propulsion system reaches the position saturation or rate saturation state and still cannot meet the force or torque required for the transition state, the remaining force or torque required for the transition state is then allocated to the fixed-wing aerodynamic control surface system, which is represented by the following formula:

[0017] ,

[0018] where, is the system function of the fixed-wing aerodynamic control surface system, is the actual input control quantity of the fixed-wing aerodynamic control surface system, is the actual control input quantity of the tiltrotor aircraft.

[0019] In a second aspect, the present application provides a computer program product, which includes a computer-readable storage medium storing a computer program. The computer program is operable to cause a computer to execute the tiltrotor aircraft transition state cascade chain control allocation method based on angle of attack dynamics as described above.

[0020] Implementing one of the above technical solutions of the present application has the following advantages or beneficial effects: In the present application, based on the construction of a non-linear system model for the tiltrotor aircraft transition state, by comparing the magnitude of the real-time angle of attack with the stall angle of attack, the control priority is dynamically allocated between the fixed-wing aerodynamic control surface system and the rotor propulsion system based on the principle of optimal control efficiency. After obtaining the optimal allocation parameters, the tiltrotor aircraft is adjusted. In this case, by comparing the real-time angle of attack with the stall angle of attack, the control allocation problem in the large angle of attack and low dynamic pressure flight environment during the tiltrotor aircraft transition state can be effectively solved, thereby improving the cruise speed and control efficiency of the tiltrotor aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] To more clearly illustrate the technical solutions of the embodiments of the present application, the following will briefly introduce the accompanying drawings required for the description of the embodiments. Obviously, the accompanying drawings in the following description are only some embodiments of the present application. For those of ordinary skill in the art, without creative efforts, other accompanying drawings can be obtained based on these drawings. In the drawings:

[0022] Figure 1 is a schematic flow chart of the control allocation method for the transition state series connection chain of a tilt-rotor aircraft based on the dynamic angle of attack in the embodiments of the present application;

[0023] Figure 2 is a schematic flow chart of the control authority allocation in the embodiments of the present application;

[0024] Figure 3 is a schematic diagram of the non-linear washout function in the embodiments of the present application;

[0025] Figure 4 is a schematic diagram of the non-linear wash-in function in the embodiments of the present application;

[0026] Figure 5 is a structural block diagram of the processing device in the embodiments of the present application.

[0027] In the figure: 1, processing device; 10, memory; 11, processor. Detailed implementation manners

[0028] In order to make the objectives, technical solutions and advantages of the present application more clear and understandable, various exemplary embodiments to be described below will refer to the corresponding accompanying drawings, which form a part of the exemplary embodiments and describe various exemplary embodiments that may be adopted to implement the present application. Unless otherwise indicated, the same numbers in different accompanying drawings represent the same or similar elements. The implementation manners described in the following exemplary embodiments do not represent all implementation manners consistent with the present disclosure. It should be understood that they are only examples of processes, methods, devices, etc. consistent with some aspects of the present application disclosed in detail in the appended claims. Other embodiments may also be used, or structural and functional modifications may be made to the embodiments listed herein without departing from the scope and essence of the present application.

[0029] In the description of the present application, it should be understood that terms such as "center", "longitudinal", "transverse", etc. indicate the orientation or positional relationship based on the orientation shown in the drawings. They are only for the convenience of describing the present application and simplifying the description, rather than indicating or implying that the indicated elements must have a specific orientation, be constructed and operated in a specific orientation. Terms such as "first", "second", etc. are only used for descriptive purposes and cannot be understood as indicating or implying relative importance or implicitly specifying the quantity of the indicated technical features. The meaning of the term "plurality" is two or more. The terms "connected" and "coupled" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, an integral connection, a mechanical connection, an electrical connection, a communication connection, a direct connection, an indirect connection through an intermediate medium, and can be the internal communication of two elements or the interaction relationship between two elements. The term "and / or" includes any and all combinations of one or more of the related listed items. For those of ordinary skill in the art, the specific meanings of the above terms in the present application can be understood according to specific circumstances.

[0030] In order to illustrate the technical solutions described in the present application, the following will be described through specific embodiments, and only the parts related to the embodiments of the present application are shown.

[0031] As Figure 1 shown, the present application provides a control allocation method for the transition state series connection chain of a tiltrotor aircraft with respect to the angle of attack dynamic, including (steps S1 to S3):

[0032] S1. Construct a transition state model of the tiltrotor aircraft and obtain the real-time angle of attack of the tiltrotor aircraft. Specifically, the aerodynamic characteristics of the fixed-wing aerodynamic control surface system and the rotor propulsion system of the tiltrotor aircraft can be analyzed to construct a transition state model of the tiltrotor aircraft.

[0033] In some embodiments, the tiltrotor aircraft may include a wing, a fuselage, a tail, and a rotor system. The fixed-wing aerodynamic control surface system of the tiltrotor aircraft may refer to the movable components for controlling the flight direction and attitude. Specifically, the fixed-wing aerodynamic control surface system can change the angle of the control surface to adjust the aerodynamic characteristics, thereby achieving pitch, roll, and yaw control. The fixed-wing aerodynamic control surface system may include an elevator, an aileron, and a rudder. The rotor propulsion system of the tiltrotor aircraft may refer to the device for generating lift and propulsion force, which controls the flight trajectory and attitude by changing the tilt angle and rotor speed of the rotor-nacelle system.

[0034] In some embodiments, the ailerons can be a pair of movable airfoils mounted on the trailing edge of the wing, used to control the roll motion of the tiltrotor aircraft. Specifically, the ailerons can be used to change the angle of attack of the wing, thereby controlling the left-right roll of the fuselage during horizontal flight. The elevator can be located on the horizontal tail of the tiltrotor aircraft and is used to control the pitch motion of the tiltrotor aircraft. Specifically, the elevator can be used to change the angle of attack of the tail to control the tiltrotor aircraft's nose-up or nose-down. The rudder can be located on the vertical tail of the tiltrotor aircraft and is used to control the yaw motion of the tiltrotor aircraft, that is, the left-right turning of the tiltrotor aircraft's nose. Specifically, the rudder can be used to control the yaw of the tiltrotor aircraft to turn it left or right.

[0035] In some embodiments, constructing a tiltrotor transition state model may include: defining an inertial coordinate system , a body coordinate system , a stability coordinate system and an airflow coordinate system ; expressing the aerodynamic forces generated by each component of the tiltrotor aircraft in the airflow coordinate system , expressing the aerodynamic moments in the body coordinate system , and constructing a transformation matrix for converting from the airflow coordinate system to the body coordinate system ; obtaining the resultant force and resultant moment of the tiltrotor transition state in the body coordinate system according to the resultant force and resultant moment acting on the center of gravity of the tiltrotor aircraft.

[0036] In some embodiments, the inertial coordinate system can refer to a reference coordinate system fixed on the earth; the body coordinate system can refer to a coordinate system fixed on the tiltrotor aircraft, corresponding to the pitch, roll, and yaw motions of the tiltrotor aircraft; the airflow coordinate system ( ) can define the angle of attack and sideslip angle of the tiltrotor aircraft.

[0037] In some embodiments, the orthogonal rotation matrix for converting from the airflow coordinate system ( ) to the body coordinate system ( ) can be expressed by the following formula:

[0038] .

[0039] In some embodiments, the tiltrotor transition state model can satisfy the following equation:

[0040]

[0041] where the subscript represents in the airflow coordinate system ( ), Indicates in the inertial coordinate system ( ), is the flight speed of the tiltrotor aircraft, is the angle of attack, is the sideslip angle, is the roll angle in the airflow coordinate system, is the track angle, is the heading of the flight path, is the displacement in the inertial coordinate system, is the angular velocity in the body coordinate system, is the angular velocity in the airflow coordinate system, are the roll angle, pitch angle and yaw angle in the body coordinate system, is the mass of the tiltrotor aircraft, is the thrust generated by the rotor propulsion system in the airflow coordinate system component in the axis direction, is the aerodynamic drag generated by the wing and fuselage in the body coordinate system axis direction, is the moment of inertia of the tiltrotor aircraft, is the product of inertia of the tiltrotor aircraft, , and axis directions.

[0042] In some embodiments, the angular velocity in the airflow coordinate system can be expressed by the following formula:

[0043]

[0044] where, and are respectively the thrust generated by the rotor propulsion system in the airflow coordinate system and axis direction components, is the aerodynamic thrust generated by the wing, fuselage and horizontal tail in the body coordinate system axis direction, is the aerodynamic sideslip force generated by the fuselage in the body coordinate system axis direction.

[0045] In some embodiments, in the body coordinate system, the resultant force acting on the center of gravity of the tiltrotor aircraft may include the aerodynamic forces generated by the wing, fuselage, horizontal tail and rotor propulsion system, and the resultant moment acting on the center of gravity of the tiltrotor aircraft may include the aerodynamic moments generated by the wing, fuselage and rotor propulsion system.

[0046] In some embodiments, the resultant force of the tiltrotor in the transition state in the body coordinate system can satisfy the following formula:

[0047] ,

[0048] wherein, , , and respectively represent the tilt angle of the rotor, aileron, elevator and rudder deflections, is the dynamic pressure, is the reference area of the wing, is the wingspan, is the aerodynamic chord length of the wing, is the body coordinate system the derivative of the aerodynamic force coefficient component in the x-axis direction with respect to the elevator deflection, is the body coordinate system the derivative of the aerodynamic force coefficient component in the y-axis direction with respect to the aileron deflection, is the body coordinate system the derivative of the aerodynamic force coefficient component in the z-axis direction with respect to the rudder deflection, is the derivative of the aerodynamic force coefficient component in the body coordinate system x-axis direction with respect to the elevator deflection, is the combined aerodynamic force coefficient of the wing, fuselage and horizontal tail in the body coordinate system z-axis direction, is the derivative of the aerodynamic force coefficient generated in the body coordinate system z-axis direction with respect to the angular velocity , is the derivative of the aerodynamic force coefficient generated in the body coordinate system z-axis direction with respect to the sideslip angle , is the derivative of the aerodynamic force coefficient generated in the body coordinate system y-axis direction with respect to the angular velocity , is the derivative of the aerodynamic force coefficient generated in the body coordinate system y-axis direction with respect to the angular velocity , is the aerodynamic force coefficient generated by the fuselage in the body coordinate system y-axis direction, is the combined aerodynamic force coefficient of the wing, fuselage and horizontal tail in the body coordinate system y-axis direction, is the derivative of the aerodynamic force coefficient generated in the body coordinate system in the axis direction with respect to the angular velocity . The rotor tilt angle can refer to the angle of the rotor system relative to the wing or fuselage. In a tiltrotor aircraft, the rotor can rotate (tilt) about an axis to change the direction of thrust.

[0049] In some embodiments, the combined aerodynamic force coefficient of the wing, fuselage, and horizontal tail in the body coordinate system in the axis direction may satisfy the following equation:

[0050] ,

[0051] where is the aerodynamic drag coefficient generated by the wing and fuselage in the body coordinate system in the axis direction, is the aerodynamic pull coefficient generated by the wing, fuselage, and horizontal tail in the body coordinate system in the axis direction.

[0052] In some embodiments, the derivative of the aerodynamic force coefficient generated in the body coordinate system in the axis direction with respect to the angular velocity may satisfy the following equation:

[0053] ,

[0054] where is the derivative of the aerodynamic drag coefficient generated in the body coordinate system in the axis direction with respect to the angular velocity , is the derivative of the aerodynamic pull coefficient generated in the body coordinate system in the axis direction with respect to the angular velocity .

[0055] In some embodiments, the derivative of the aerodynamic force coefficient of the component in the body coordinate system in the axis direction with respect to the elevator deflection may satisfy the following equation:

[0056] ,

[0057] where is the derivative of the aerodynamic drag coefficient generated in the body coordinate system in the axis direction with respect to the elevator deflection, is the derivative of the aerodynamic pull coefficient generated in the body coordinate system in the axis direction with respect to the elevator Derivative of deflection.

[0058] In some embodiments, the aerodynamic resultant force coefficients of the wing, fuselage, and horizontal tail in the body coordinate system in the axis direction can satisfy the following formula:

[0059] ,

[0060] where is the aerodynamic drag coefficient generated by the wing and fuselage in the body coordinate system in the axis direction, is the aerodynamic pulling force coefficient generated by the wing, fuselage, and horizontal tail in the body coordinate system in the axis direction.

[0061] In some embodiments, the derivative of the aerodynamic force coefficient generated in the body coordinate system in the axis direction with respect to the angular velocity can satisfy the following formula:

[0062] ,

[0063] where is the derivative of the aerodynamic drag coefficient generated in the body coordinate system in the axis direction with respect to the angular velocity , is the derivative of the aerodynamic pulling force coefficient generated in the body coordinate system in the axis direction with respect to the angular velocity .

[0064] In some embodiments, the derivative of the aerodynamic force coefficient of the component in the body coordinate system in the axis direction with respect to the elevator deflection can satisfy the following formula:

[0065] ,

[0066] where is the derivative of the aerodynamic drag coefficient generated in the body coordinate system in the axis direction with respect to the elevator deflection, is the derivative of the aerodynamic pulling force coefficient generated in the body coordinate system in the axis direction with respect to the elevator deflection.

[0067] In some embodiments, the aerodynamic pulling force coefficient generated by the wing, fuselage, and horizontal tail in the body coordinate system in the axis direction​ and the aerodynamic drag coefficients generated by the wing and the fuselage in the body coordinate system in the axial direction can satisfy the following formula:

[0068] ,

[0069] where the subscripts 、 and respectively represent the wing, the fuselage, and the horizontal tail. is the aerodynamic lift coefficient generated by the wing, is the aerodynamic lift coefficient generated by the horizontal tail, is the aerodynamic lift coefficient generated by the fuselage, is the aerodynamic drag coefficient generated by the wing, is the aerodynamic drag coefficient generated by the fuselage.

[0070] In some embodiments, the aerodynamic lift coefficient generated by the horizontal tail can satisfy the following formula:

[0071] ,

[0072] where is the reference area of the horizontal tail, is the effective angle of attack of the horizontal tail.

[0073] In some embodiments, the effective angle of attack of the horizontal tail can satisfy the following formula:

[0074] ,

[0075] where is the downwash angle of the horizontal tail, is the longitudinal distance between the aerodynamic center of the horizontal tail and the aerodynamic centers of the wing and the fuselage.

[0076] In some embodiments, the aerodynamic lift and the drag coefficient generated by the fuselage can satisfy the following formula:

[0077] ,

[0078] where is the sideslip angle of the fuselage, is the derivative of the aerodynamic lift coefficient generated by the fuselage with respect to the angle of attack, is the derivative of the aerodynamic lift coefficient generated by the fuselage with respect to the sideslip angle, is the aerodynamic lift coefficient generated by the fuselage when both the angle of attack and the sideslip angle are zero.​ is the additional pulling force coefficient of the fuselage, is the derivative of the aerodynamic drag coefficient generated by the fuselage with respect to the angle of attack, is the derivative of the aerodynamic drag coefficient generated by the fuselage with respect to the sideslip angle, is the aerodynamic drag coefficient generated by the fuselage when both the angle of attack and the sideslip angle are zero, is the additional drag coefficient of the fuselage.

[0079] In some embodiments, the resultant moment of the tilt-rotor aircraft in the transition state in the body coordinate system can satisfy the following formula:

[0080]

[0081] , where is the distance from the two tilt rotors to the center of gravity of the tilt-rotor aircraft, is the derivative of the rolling moment coefficient generated by the wing with respect to the sideslip angle, is the rolling moment coefficient with respect to the angular velocity of the derivative, is the rolling moment coefficient with respect to the angular velocity of the derivative, is the pitching moment coefficient generated by the wing, is the derivative of the pitching moment coefficient with respect to the angular velocity of the derivative, is the derivative of the yaw moment coefficient with respect to the sideslip angle, is the yaw moment coefficient with respect to the angular velocity of the derivative, is the yaw moment coefficient with respect to the angular velocity of the derivative, is the derivative of the rolling moment coefficient generated by the wing with respect to the aileron deflection, is the derivative of the rolling moment coefficient with respect to the rudder deflection, is the derivative of the pitching moment coefficient with respect to the elevator deflection, is the derivative of the yaw moment coefficient generated by the wing with respect to the aileron deflection, is the derivative of the yaw moment coefficient with respect to the rudder deflection, is the rolling moment coefficient generated by the fuselage, is the pitching moment coefficient generated by the fuselage, is the yaw moment coefficient generated by the fuselage.

[0082] In some embodiments, the pitching moment coefficient generated by the fuselage can satisfy the following formula:

[0083] ,

[0084] wherein, is the derivative of the pitching moment coefficient generated by the fuselage with respect to the angle of attack, is the derivative of the pitching moment coefficient generated by the fuselage with respect to the sideslip angle, is the pitching moment coefficient generated by the fuselage when both the angle of attack and the sideslip angle are zero.

[0085] S2. Compare the real-time angle of attack of the tiltrotor with the stall angle of attack, dynamically allocate control authority between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result, and obtain the optimal allocation parameters of the tiltrotor according to the allocation result of the control authority, where the optimal allocation parameters at least include aileron deflection, rudder deflection, elevator deflection, rotor tilt angle, and rotor speed.

[0086] In some embodiments, as Figure 2 shown, comparing the real-time angle of attack of the tiltrotor with the stall angle of attack and dynamically allocating control priorities between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result may include: if the real-time angle of attack is greater than the stall angle of attack, allocate the control authority to the rotor propulsion system, and if the rotor propulsion system meets the force or moment required for the transition state, end the control allocation task; if the rotor propulsion system reaches the position saturation or rate saturation state and still does not meet the force or moment required for the transition state, then allocate the remaining force or moment required for the transition state to the fixed-wing aerodynamic control surface system.

[0087] In some embodiments, the force and moment required for the transition state may include lift, thrust, pitching force, and pitching moment, rolling moment, yaw moment, and rotor tilt moment. Allocating the control authority to the rotor propulsion system may refer to allocating control commands generated by the flight control system, such as pitching, rolling, yawing, and elevating commands, to each rotor of the rotor propulsion system. The rotor propulsion system reaching the position saturation state means that its angle or position has reached the limit, such as pitch angle saturation, roll angle saturation, yaw angle saturation, maximum rotational speed saturation, etc.

[0088] In some embodiments, allocating the control authority to the rotor propulsion system and ending the control allocation task if the rotor propulsion system can provide the force or moment required for the transition state can be represented by the following formula:

[0089] ,

[0090] wherein, is the system function of the rotor propulsion system, is the force or moment required for the transition state of the tiltrotor, is the saturation function, is the actual input control quantity of the rotor propulsion system.

[0091] In some embodiments, when the rotor propulsion system reaches the position saturation or rate saturation state and still does not meet the required force or moment in the transition state, the remaining required force or moment in the transition state is then allocated to the fixed-wing aerodynamic control surface system, which can be expressed by the following formula:

[0092] ,

[0093] Wherein, is the system function of the fixed-wing aerodynamic control surface system, is the actual input control quantity of the fixed-wing aerodynamic control surface system, is the actual control input quantity of the tiltrotor aircraft in this case.

[0094] In some embodiments, the real-time angle of attack of the tiltrotor aircraft is compared with the stall angle of attack, and the control priority is dynamically allocated between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result, including: when the real-time angle of attack is less than the stall angle of attack, the control authority is allocated to the fixed-wing aerodynamic control surface system. If the fixed-wing aerodynamic control surface system can provide the required force or moment in the transition state, the control allocation task is ended; if the fixed-wing aerodynamic control surface system reaches the position saturation or rate saturation state and still does not meet the required force or moment in the transition state, the remaining required force or moment in the transition state is then allocated to the rotor propulsion system.

[0095] Specifically, allocating the control authority to the fixed-wing aerodynamic control surface system may refer to allocating the control instructions generated by the flight control system to the ailerons, elevators, and rudders of the fixed-wing aerodynamic control surface system. The fixed-wing aerodynamic control surface system reaching the position saturation or rate saturation state means that the deflection angle and deflection rate of the control surface reach saturation.

[0096] In some embodiments, when the control authority is allocated to the fixed-wing aerodynamic control surface system and the fixed-wing aerodynamic control surface system can provide the required force or moment in the transition state, the control allocation task is ended, which can be expressed by the following formula:

[0097] ,

[0098] Wherein, is the actual input control quantity of the fixed-wing aerodynamic control surface system, is the system function of the fixed-wing aerodynamic control surface system.

[0099] In some embodiments, when the fixed-wing aerodynamic control surface system reaches the position saturation or rate saturation state and still does not meet the required force or moment in the transition state, the remaining required force or moment in the transition state is then allocated to the rotor propulsion system, which can be expressed by the following formula:

[0100] ,

[0101] Among them, is the actual input control quantity of the rotor propulsion system, is the actual control input quantity of the tilt-rotor aircraft in this case.

[0102] In some embodiments, such as Figure 3 and Figure 4 shown, the cascade chain control allocation method for the transition state of the tilt-rotor aircraft based on the angle of attack dynamics may further include: designing a nonlinear washout function for the wind tunnel data in the helicopter mode, designing a nonlinear wash-in function for the wind tunnel data in the fixed-wing mode, and obtaining the aerodynamic coefficients of the required forces and moments in the transition state.

[0103] Specifically, when the tilt-rotor aircraft is in the transition state, the tilt-rotor aircraft will transition from the helicopter mode to the fixed-wing mode, or from the fixed-wing mode to the helicopter mode. In the transition state, the nonlinear washout function can gradually reduce the weight of the wind tunnel data in the helicopter mode, and thus these data can be applied to the construction process of the relevant aerodynamic coefficients and aerodynamic moment coefficients in the above-mentioned tilt-rotor aircraft transition state model. In the transition state, the nonlinear wash-in function can gradually increase the weight of the wind tunnel data in the fixed-wing mode, and thus these data can be applied to the construction process of the relevant aerodynamic coefficients and aerodynamic moment coefficients in the above-mentioned tilt-rotor aircraft transition state model, such as the lift curve slope, drag coefficient, moment coefficient, etc.

[0104] In some embodiments, in the nonlinear washout function, the weight of the wind tunnel data in the helicopter mode may be positively correlated with the degree of the tilt angle. Specifically, when the tilt angle is from 0 degrees to 30 degrees, the weight of the wind tunnel data in the helicopter mode can linearly increase from 0 to 0.3; when the tilt angle is from 30 degrees to 60 degrees, the weight of the wind tunnel data in the helicopter mode can linearly increase from 0.35 to 0.65; when the tilt angle is from 60 degrees to 90 degrees, the weight of the wind tunnel data in the helicopter mode can linearly increase from 0.7 to 1.

[0105] In some embodiments, in the nonlinear wash-in function, the weight of the wind tunnel data in the fixed-wing mode may be negatively correlated with the degree of the tilt angle. Specifically, when the tilt angle is from 0 degrees to 30 degrees, the weight of the wind tunnel data in the fixed-wing mode can linearly decrease from 1 to 0.7; when the tilt angle is from 30 degrees to 60 degrees, the weight of the wind tunnel data in the fixed-wing mode can linearly decrease from 0.65 to 0.35; when the tilt angle is from 60 degrees to 90 degrees, the weight of the wind tunnel data in the fixed-wing mode can linearly decrease from 0.3 to 0.

[0106] In summary, by designing the non-linear washout function and the non-linear wash-in function, the aerodynamic coefficients of the forces and moments required for the transition state of the tiltrotor aircraft can be obtained, which facilitates the subsequent allocation of the control authority priority according to the required forces or moments of the tiltrotor aircraft, making the subsequent allocation process more accurate.

[0107] S3. Input the optimal allocation parameters into the tiltrotor aircraft transition state model to adjust the tiltrotor aircraft.

[0108] In some embodiments, the optimal allocation parameters can be used to adjust the attitude, heading, altitude, and speed of the tiltrotor aircraft. Specifically, aileron deflection can be used to control the roll motion of the tiltrotor aircraft, rudder deflection can be used to control the yaw motion of the tiltrotor aircraft, elevator deflection can be used to control the pitch motion of the tiltrotor aircraft, the rotor tilt angle can be used to control the transition of the tiltrotor aircraft from the helicopter mode to the fixed-wing mode, and the rotor speed can be used to control the lift, thrust, and power of the tiltrotor aircraft.

[0109] In some embodiments, the tiltrotor aircraft transition state cascade chain control allocation method based on the angle of attack dynamics may further include: designing the tiltrotor aircraft transition state flight control law, adopting a multi-loop control structure, and establishing a controller based on non-linear dynamic inversion. Specifically, non-linear dynamic inversion may refer to obtaining the input of the control law according to the desired flight state, where the desired flight state may be related to the optimal allocation parameters, and then obtaining the output of the control law, that is, directly applying the obtained control input to the controller to manipulate the tiltrotor aircraft to reach the expected flight state.

[0110] In some embodiments, the formula of the tiltrotor aircraft transition state dynamics equation is:

[0111]

[0112] Wherein, is the slow-loop state, is the fast-loop state, is the virtual control input.

[0113] In some embodiments, the formula of the tracking error dynamics equation is:

[0114]

[0115] Wherein, and are the tracking errors of the slow loop and the fast loop respectively, is the slow-loop reference signal.

[0116] In some embodiments, the control law of the slow loop can satisfy the following formula:

[0117]

[0118] Among them, the control law can satisfy the following formula:

[0119] ,

[0120] Among them, is the error between the current state and the reference signal, , and are the controller gains.

[0121] In some embodiments, the control law of the fast loop can satisfy the following formula:

[0122]

[0123] Among them, the control law can satisfy the following formula:

[0124] ,

[0125] Among them, is the error between the current state and the reference signal, , and are the controller gains.

[0126] In this application, based on the construction of the nonlinear system model of the tilt-rotor aircraft in the transition state, by comparing the magnitude of the real-time angle of attack and the stall angle of attack, the control priority is dynamically allocated between the fixed-wing aerodynamic control surface system and the rotor propulsion system based on the principle of optimal control efficiency. After obtaining the optimal allocation parameters, the tilt-rotor aircraft is adjusted. In this case, by comparing the real-time angle of attack and the stall angle of attack, the control allocation problem in the large angle of attack and low dynamic pressure flight environment during the transition state of the tilt-rotor aircraft can be effectively solved, thereby improving the cruise speed and control efficiency of the tilt-rotor aircraft.

[0127] The present application also provides a computer program product. The computer program product is stored on a data carrier and is designed to execute the tilt-rotor aircraft transition state concatenated chain control allocation method based on angle of attack dynamics as described above. Therefore, the computer program product according to the present application has the same advantages as those described in detail with reference to the device according to the present application. The computer program product can be executed as computer-readable instruction codes in each appropriate programming language such as JAVA, C++, etc. In addition, the computer program product can be provided on a network, such as the Internet, or a network user can download the computer program product from a network, such as the Internet, when needed. The computer program product can be implemented either by means of a computer program, i.e., software, or by means of one or more dedicated electronic circuits, i.e., hardware, or in any hybrid form, i.e., by means of software components and hardware components, or in a hybrid form of software, hardware, or software and hardware.

[0128] Those of ordinary skill in the art can understand that all or part of the features / steps of implementing the above method embodiments can be achieved by a method, a data processing system, or a computer program. These features can be implemented without using hardware, entirely using software, or using a combination of hardware and software. The aforementioned computer program can be stored in one or more computer-readable storage media. When the computer program stored on the storage media is executed (such as by a processor), it executes the steps of the embodiment of the tilt-rotor aircraft transition state concatenated chain control allocation method based on angle of attack dynamics as described above.

[0129] The aforementioned storage media that can store program codes include: a static hard disk, a solid-state drive, a random access memory (SRAM), an electrically erasable programmable read-only memory (EEPROM), an erasable programmable read-only memory (EPROM), a programmable read-only memory (PROM), a read-only memory (ROM), an optical storage device, a magnetic storage device, a flash memory, a magnetic disk or an optical disc, and / or a combination of the above devices, that is, it can be implemented by any type of volatile or non-volatile storage device or a combination thereof.

[0130] As Figure 5 shown, the present application also provides an embodiment of a processing device 1, including one or more processors 11 and a memory 10; wherein, the memory 10 is used to store one or more computer programs, and one or more processors 11 are used to execute the one or more computer programs stored in the memory 10, so that the processor 11 executes the features / steps of the embodiment of the tilt-rotor aircraft transition state concatenated chain control allocation method based on angle of attack dynamics as described above.

[0131] The above are only the preferred embodiments of the present application. Those skilled in the art will understand that various changes or equivalent substitutions can be made to these features and embodiments without departing from the spirit and scope of the present application. Additionally, under the teachings of the present application, these features and embodiments can be modified to adapt to specific circumstances and materials without departing from the spirit and scope of the present application. Therefore, the present application is not limited by the specific embodiments disclosed herein, and all embodiments falling within the scope of the claims of the present application belong to the protection scope of the present application.

Claims

1. A control allocation method for a tilt-rotor aircraft transition state series connection chain based on angle of attack dynamics, characterized in that Including: Construct a transition state model of a tiltrotor aircraft to obtain the real-time angle of attack of the tiltrotor aircraft; Compare the real-time angle of attack of the tiltrotor aircraft with the stall angle of attack, dynamically allocate control authority between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result, and obtain the optimal allocation parameters of the tiltrotor aircraft according to the allocation result of the control authority, where the optimal allocation parameters at least include aileron deflection, rudder deflection, elevator deflection, rotor tilt angle, and rotor speed; Input the optimal allocation parameters into the transition state model of the tiltrotor aircraft to adjust the tiltrotor aircraft; The comparing the real-time angle of attack of the tiltrotor aircraft with the stall angle of attack and dynamically allocating control authority between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result includes: if the real-time angle of attack is greater than the stall angle of attack, allocate the control authority to the rotor propulsion system, and if the rotor propulsion system can provide the force or moment required for the transition state, end the control allocation task; if the rotor propulsion system reaches the position saturation or rate saturation state and still does not meet the force or moment required for the transition state, then allocate the remaining force or moment required for the transition state to the fixed-wing aerodynamic control surface system; The allocating the control authority to the rotor propulsion system and ending the control allocation task if the rotor propulsion system can provide the force or moment required for the transition state is represented by the following formula: ， Among them, is the system function of the rotor propulsion system, is the force or torque required for the transition state, is the saturation function, is the actual input control quantity of the rotor propulsion system.

2. The control allocation method for the transition state series connection chain of a tilt-rotor aircraft based on angle of attack dynamics according to claim 1, wherein The comparing the real-time angle of attack of the tiltrotor aircraft with the stall angle of attack and dynamically allocating control authority between the fixed-wing aerodynamic control surface system and the rotor propulsion system according to the comparison result includes: if the real-time angle of attack is less than the stall angle of attack, allocate the control authority to the fixed-wing aerodynamic control surface system, and if the fixed-wing aerodynamic control surface system can provide the force or moment required for the transition state, end the control allocation task; if the fixed-wing aerodynamic control surface system reaches the position saturation or rate saturation state and still does not meet the force or moment required for the transition state, then allocate the remaining force or moment required for the transition state to the rotor propulsion system.

3. The control allocation method for the transition state series connection chain of a tilt-rotor aircraft based on angle of attack dynamics according to claim 1, wherein, The method further includes: designing a non-linear washout function for the wind tunnel data in the helicopter mode and a non-linear wash-in function for the wind tunnel data in the fixed-wing mode to obtain the aerodynamic coefficients of the force and moment required for the transition state.

4. The control allocation method for the transition state series connection chain of a tilt-rotor aircraft based on the dynamic angle of attack according to claim 1, wherein The method further includes: designing a transition state flight control law for the tiltrotor aircraft, adopting a multi-loop control structure, and establishing a controller based on non-linear dynamic inversion.

5. The control allocation method for the transition state series connection chain of a tilt-rotor aircraft based on the dynamic angle of attack according to claim 1, wherein The constructing the transition state model of the tiltrotor aircraft includes: defining a body coordinate system and an airflow coordinate system; expressing the aerodynamic forces generated by each component of the tiltrotor aircraft in the airflow coordinate system and expressing the aerodynamic moments in the body coordinate system, and constructing a transformation matrix for converting from the airflow coordinate system to the body coordinate system; obtaining the resultant force and resultant moment of the transition state of the tiltrotor aircraft in the body coordinate system according to the resultant force and resultant moment acting on the center of gravity of the tiltrotor aircraft.

6. The control allocation method for the transition state series connection chain of a tilt-rotor aircraft based on the dynamic angle of attack according to claim 3, wherein In the non-linear washout function, the weight of the wind tunnel data in the helicopter mode is positively correlated with the degree of the tilt angle; in the non-linear wash-in function, the weight of the wind tunnel data in the fixed-wing mode is negatively correlated with the degree of the tilt angle.

7. The control allocation method for the transition state series connection chain of a tilt-rotor aircraft based on the dynamic angle of attack according to claim 1, wherein If the rotor propulsion system reaches the position saturation or rate saturation state and still does not meet the force or moment required for the transition state, the remaining force or moment required for the transition state is then allocated to the fixed-wing aerodynamic control surface system, which is expressed by the following formula: ， Among them, is the system function of the fixed-wing aerodynamic control surface system, is the actual input control quantity of the fixed-wing aerodynamic control surface system, is the actual control input quantity of the tilt-rotor aircraft.

8. A computer program product, characterized in that, The computer program product includes a computer-readable storage medium storing a computer program, and the computer program is operable to cause a computer to execute the control allocation method for the transition state series connection chain of a tiltrotor based on the angle of attack dynamics according to any one of claims 1 to 7.

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