Pose stability control method and system under engine fault of composite wing unmanned aerial vehicle

Abstract

The invention relates to the technical field of unmanned aerial vehicle flight control, and mainly relates to a pose stability control method and system under a composite wing unmanned aerial vehicle engine fault. The method comprises the following steps: establishing an unmanned aerial vehicle deceleration solving equation including a forward inequality, a vertical balance equation and a pitching moment balance equation according to the stress after the engine of the composite wing unmanned aerial vehicle breaks down; according to the real-time airspeed after an engine fault occurs, a weight switching equation and a control quantity distribution matrix which take the airspeed as a limiting condition are established, and control quantity distribution between a rotor wing and a control surface is coordinated. And when the deceleration solving equation of the unmanned aerial vehicle is established, the pose stability is realized under the engine fault of the composite wing unmanned aerial vehicle. According to the invention, force and moment distribution and control among different power sources are coordinated, and the unmanned aerial vehicle can maintain a stable flight state when an engine has a fault.

Description

Technical Field

[0001] The present invention relates to the technical field of unmanned aerial vehicle flight control, and in particular to a method and system for attitude stabilization control of an oil-electric hybrid composite wing unmanned aerial vehicle under engine failure. Background Art

[0002] In recent years, with the rapid development of computer technology, various sensors and communication technology, the application scenarios of drones have become increasingly extensive, and they have gradually become an indispensable intelligent tool in modern society. However, traditional mainstream models, such as fixed-wing aircraft and multi-rotor drones, have problems such as limited load capacity, strong dependence on take-off and landing runways, and limited range, which restrict their flexibility and diversity in practical applications. In response to these limitations, composite wing drones came into being. This type of drone combines the technical advantages of fixed wings and rotors, and can better perform tasks such as surveying and mapping, inspection, and emergency rescue.

[0003] Composite wing UAVs achieve long flight time and flexible maneuverability through the synergy of two power sources: the engine and the rotor. However, in actual operations, problems such as insufficient fuel supply and power system failure may occur, resulting in the risk of engine thrust loss and power interruption. To ensure flight safety, in the event of engine failure, measures must be taken quickly to stabilize the flight attitude of the UAV and switch to another power system for emergency fault handling. Therefore, it is urgent to study an attitude stabilization control method under engine failure that is suitable for composite wing UAVs, so as to ensure that the flight attitude and position can be maintained stable in the event of a failure, and to achieve safe flight of composite wing UAVs under power failure conditions. Summary of the invention

[0004] The purpose of the present invention is to provide a method and system for attitude stabilization control of a composite wing UAV under engine failure, which solves the deceleration equation and introduces a weight switching equation based on airspeed threshold and a control quantity distribution matrix to achieve switching and coordinated control between different power sources, thereby ensuring the stability of the attitude and position of the UAV after engine failure, and significantly improving the safety and reliability of the composite wing UAV in complex flight missions.

[0005] This method takes into account that the composite wing UAV cannot provide the force and torque required to maintain normal flight after the engine fails. In order to maintain the stability of the attitude and position of the UAV, the weight switching equation and control quantity allocation matrix based on the airspeed threshold are designed by solving the deceleration equation, so as to realize the switching and coordinated control between different power sources.

[0006] A first aspect of the present invention is to provide a method for attitude stabilization control of a composite wing UAV under engine failure, the method comprising:

[0007] Step S1, according to the force analysis after the composite wing UAV engine failure occurs, establish the UAV deceleration solution equation including the forward inequality, the vertical balance equation and the pitch moment balance equation; the force includes the aerodynamic lift L, the aerodynamic drag D, and the sum of the pulls generated by the rotors T (that is, the sum of the pulls generated by all rotors);

[0008] Step S2, dynamically adjust the control strategy according to the real-time airspeed after the engine failure occurs, establish a weight switching equation and a control quantity allocation matrix with the real-time airspeed as a constraint condition, and coordinate the control quantity allocation between the rotor and the control surface; the airspeed is the true airspeed V of the UAV in the body coordinate system;

[0009] Step S3: Calculate the real-time aerodynamic lift L, the real-time aerodynamic drag D, and the real-time aerodynamic pitch moment M by distributing the control amount between the rotor and the control surface. y , the sum of the thrust T generated by the real-time rotor, and then substitute it into the deceleration equation of the drone in step S1. When the real-time L, D, M y , T makes the UAV decelerate and solve the equation, then the composite wing UAV can achieve attitude stability when the engine fails.

[0010] In a preferred embodiment, the method further comprises verifying the control strategy by simulation.

[0011] In a preferred embodiment, in step S1, the deceleration equation of the drone is composed of a forward inequality, a vertical balance equation and a pitch moment balance equation:

[0012]

[0013] In the formula, s (·) = sin(·),c (·) =cos(·), · is θ or α; T is the sum of the thrusts generated by the rotor; L and D represent the aerodynamic lift and aerodynamic drag, respectively; θ and α are the pitch angle and angle of attack of the fuselage, respectively (measured in real time by the UAV sensor and fed back to the attitude stability control system of the composite wing UAV under engine failure); M y is the aerodynamic pitch moment; l is the distance between the projection of the rotor on the symmetry plane of the aircraft and the center of gravity of the UAV; t Vc →∞ indicates that the rotor thrust changes from t Vc appears at the moment and exists forever; 0→t Va It indicates that the aerodynamic lift existed before and at t Va Moment disappears; V a V is the airspeed at which the aerodynamic force decreases to negligible level; c The airspeed at which flight controls are added to the rotor. V bis the airspeed when the rotor weight reaches 100%. The disappearance of aerodynamic lift has the same meaning as the reduction of aerodynamic force to negligible. In this paper, it is assumed that aerodynamic lift disappears when the airspeed is 10m / s.

[0014] Furthermore, in step S1, the expressions of aerodynamic lift, drag and pitching moment are:

[0015]

[0016] Where ρ is the air density; S ref is the reference area of ​​the wing; V is the true airspeed in the body coordinate system; C D and C L are the aerodynamic drag coefficient and aerodynamic lift coefficient respectively; C my is the aerodynamic pitch moment coefficient; c A is the aerodynamic chord length; δ A ,δ E ,δ R are the deflections of the aileron, elevator, and rudder after weight distribution; β is the sideslip angle; α and β are measured in real time by the UAV sensor and fed back to the attitude stability control system of the composite wing UAV under engine failure. (δ A ,δ E ,δ R ,α,β) is C D , C L , C my The influencing parameters are obtained by conducting wind tunnel tests or numerical simulation traversal calculations on the UAV in advance to obtain specific (δ A ,δ E ,δ R ,α,β) corresponding to C D , C L , C my The value of .

[0017] In a preferred embodiment, in step S2, the airspeed is used as the limiting condition, and when the airspeed V is greater than V c When the airspeed V is less than or equal to V c and greater than V b When the airspeed V is equal to V b or less than V b When the airspeed V is less than V a When the airspeed is greater than or equal to V a When , the rudder weight is 100%;

[0018] The weight switching equation with airspeed as the constraint is:

[0019]

[0020] Among them, K u (V) is the rotor weight switching equation, K δ (V) is the control surface weight switching equation.

[0021] The weight of the rudder surface is K δ (V) value; rotor weight is K u The value of (V).

[0022] In step S2, in order to balance the force and torque relationship of the UAV, the control quantity distribution matrix between the rotor and the control surface is established

[0023]

[0024] In the formula, δ=[δ A ,δ E ,δ R ,0] T represents the fixed-wing control quantity after weight distribution, U=[U A ,U E ,U R ,U T ] T represents the rotor control quantity after weight distribution, δ′=[δ′ A ,δ′ E ,δ R ′,0] T represents the fixed-wing control quantity without weight assignment, U′=[U′ A ,U′ E ,U′ R ,U′ T ] T represents the unweighted rotor control quantity, δ A and δ A ' A They represent the weighted and unweighted aileron deflections, δ E and δ E ' E Denote the weighted and unweighted elevator deflections, δ R and δ R ′ represents the rudder deflection of the rudder after weight allocation and the rudder without weight allocation respectively, 0 represents the engine throttle (0 due to engine failure); U A and U′ A are the weighted and unweighted rotor roll control quantities, respectively, U E and U′ E are the weighted and unweighted pitch control quantities, U R and U′ Rare the yaw control quantities after weight assignment and without weight assignment, U T and U′ T are the rotor throttle control quantities after weight allocation and without weight allocation respectively; K δ K δ (V); K u K u (V);

[0025] H and φ, θ, ψ are the real-time altitude and attitude angle information of the UAV respectively; H g and φ g ,θ g ,ψ g are the target height and target attitude angle of the UAV respectively; is the control quantity allocation function, K p , K d , K I They are proportional control parameters, differential control parameters, and integral control parameters (and different Δ may have different K p , K d , K I , K corresponding to each Δ p , K d , K I Both are adjustable and determined according to actual conditions), Δ is the deviation between the real-time status information and the target status information; the real-time status information includes the real-time altitude and real-time attitude angle information; the target status information includes the target altitude and target attitude angle information.

[0026] In step S3, the real-time rotor control amount and control surface control amount are calculated based on the control amount distribution matrix between the rotor and the control surface, and the real-time L, D, M y , whether T satisfies the UAV deceleration solution equation of formula (1).

[0027] Specifically, according to formula (4), δ A ,δ E ,δ R , δ A ,δ E ,δ R Substituting into formula (2), we can calculate L, D, M y According to formula (4), U A ,U E ,U R ,U T The rotor speed ω is obtained by mapping the matrix B. Assuming that the UAV is a six-rotor UAV, ω1, ω2, ω3, ω4, ω5, and ω6 are the speeds of the six rotors, respectively, satisfying the following formula:

[0028]

[0029] The above formula can be used to calculate the rotation speed of each rotor, and then calculate the sum of the thrust generated by the rotor. K t is the tensile coefficient. y Enter it together with T into the equation (1) to solve the UAV deceleration equation.

[0030] If the rotor control quantity and rudder control quantity calculated at a certain moment after the UAV fails, the corresponding real-time L, D, M y and T, and the first time the deceleration equation of the UAV is established, it is determined that the posture has achieved stable control, and the UAV continues to control the UAV with the rotor control amount and the rudder control amount when the deceleration equation of the UAV is established.

[0031] According to simulation, usually within 30 seconds after the UAV fails, the attitude stabilization control method of the composite wing UAV engine failure of the present invention can satisfy the UAV deceleration solution equation of formula (1), that is, achieve attitude stability.

[0032] The second aspect of the present invention is to disclose a composite wing UAV engine failure attitude stabilization control system, comprising:

[0033] The UAV deceleration solution module establishes the UAV deceleration solution equations including the forward inequality, vertical balance equation and pitch moment balance equation based on the force analysis after the composite wing UAV engine failure occurs;

[0034] The weight switching module establishes the rotor weight switching equation and the control surface weight switching equation with airspeed as the constraint condition;

[0035] The control quantity allocation module for the rudder and rotor establishes the control quantity allocation matrix between the rotor and the rudder, and completes the control quantity allocation between the rotor and the rudder according to the deviation between the real-time altitude and real-time attitude angle information and the target altitude and target attitude angle information, as well as the control quantity allocation matrix.

[0036] In the weight switching module, the airspeed is used as the limiting condition. When the airspeed V is greater than V c When the airspeed V is less than or equal to V c and greater than V b When the airspeed V is equal to V b or less than V b When the airspeed V is less than V a When the airspeed is greater than or equal to V a , the rudder weight is 100%.

[0037] The third aspect of the present invention is to provide a processor, which is used to run a computer program. When the computer program is running, the attitude stabilization control method under engine failure of the composite wing UAV as described above is executed.

[0038] The fourth aspect of the present invention is to disclose an unmanned aerial vehicle, including the aforementioned composite wing unmanned aerial vehicle engine failure attitude stabilization control system.

[0039] A fifth aspect of the present invention is to provide a computer-readable medium having a computer program stored thereon, wherein the computer program is executed by a processor as described above in the method for attitude stabilization control of a composite wing UAV under engine failure.

[0040] Beneficial effects:

[0041] The weight switching equation and control quantity distribution matrix based on airspeed threshold designed by the present invention can effectively realize the dynamic adjustment of the control mechanism (rudder and rotor), coordinate the force and torque distribution and control between different power sources, and ensure that the UAV can maintain a stable flight state when an engine failure (i.e., failure) occurs. This method reduces the risk of accidents and significantly improves the adaptability and reliability of the UAV in abnormal situations, thereby meeting the safety requirements of complex flight missions. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] Figure 1 A schematic diagram of the attitude stabilization control strategy based on airspeed threshold provided by the present invention;

[0043] Figure 2 A schematic diagram of engine failure and airspeed decay of the UAV provided by the present invention;

[0044] Figure 3 The attitude angle change curve provided by the present invention;

[0045] Figure 4 The position tracking curve provided by the present invention;

[0046] Figure 5 This is a schematic diagram of the rotor control amount and rudder control amount provided by the present invention. DETAILED DESCRIPTION

[0047] The present invention will be further described below in conjunction with the accompanying drawings.

[0048] This embodiment provides a method for attitude stabilization control of a composite wing UAV under engine failure. The relevant physical parameters of the composite wing UAV are: UAV mass m = 135.2 kg, wing reference area S ref =3m 2 , mean aerodynamic chord length c A =0.438m, upper and lower limits of rudder deflection δA ∈[-25°,25°],δ E ∈[-20°,20°],δ R ∈[-30°,30°], the distance between the projection of the rotor on the aircraft symmetry plane and the center of gravity l=1.3m, the airspeed threshold V a =10m / s,V b =15m / s,V c =32m / s. Figure 1 As shown, the present invention constructs a deceleration equation according to the stress conditions of the UAV after the engine of the composite wing UAV fails, and designs a weight switching equation and a control quantity distribution matrix based on the airspeed threshold. By gradually introducing rotor control to compensate for the attenuation of the control surface efficiency, a good flight state can be effectively maintained.

[0049] During the flight, the UAV flight control system determines that the engine has failed (based on the engine output, which is a prior art and will not be described in detail). At this time, the UAV's airspeed V is V max ; Start the attitude stabilization control under the engine failure of the composite wing UAV, and input the true airspeed V in the body coordinate system detected by the UAV sensor (such as the airspeed meter) into the attitude stabilization control system under the engine failure of the composite wing UAV. δ T is the engine throttle. When the engine fails, δ T is 0.

[0050] The user presets the target position X at all times after the drone takes off through the drone flight control system. g ,Y g ,H g . Calculate the target attitude angle at all times after the drone takes off based on the target position (this is the prior art). g ,Y g ,H g is the target position of the UAV in the inertial coordinate system, H g Represents the target height, X g ,Y g is the target position in two mutually perpendicular directions in the horizontal plane of the inertial coordinate system (which is the prior art). g is the target roll angle, θ g is the target pitch angle, ψ g is the target yaw angle.

[0051] It includes the following steps:

[0052] Step S1: According to the force analysis after the composite wing UAV engine failure occurs, the deceleration solution equation including the forward inequality, the vertical balance equation and the pitch moment balance equation is established:

[0053]

[0054] In the formula, s (·) = sin(·),c (·) = cos(·); T is the sum of the thrusts generated by the rotor; L and D represent the aerodynamic lift and drag, respectively; θ and α are the pitch angle and angle of attack of the aircraft; M y is the aerodynamic pitch moment; l is the distance between the projection of the rotor on the aircraft symmetry plane and the center of gravity; t Vc →∞ indicates that the rotor thrust changes from t Vc appears at the moment and exists forever; 0→t Va Indicates that it existed before and is in t Va Moment disappears; V a V is the airspeed at which the aerodynamic force decreases to negligible level; c The airspeed when adding flight control to the rotor; V b for Figure 1 The airspeed at which mid rotor weight reaches 100%.

[0055] When the rotor weight reaches 100%, the UAV posture is completely controlled by the rotor. Taking the sum of the thrust generated by the rotor as an example, when the UAV has no fault, the rotor does not rotate, that is, T = 0; when a fault occurs, the airspeed drops to V c When the rotor starts to rotate, T does not start to be 0. When the airspeed V = V a The time at this time is t va , when airspeed V=V c The time at this time is t vc . It is the sum of the thrust generated by the rotor at any time after the rotor thrust begins to appear (including when the rotor thrust begins to appear); is the aerodynamic pitch moment at any moment from the initial moment to the moment when the airspeed reaches 10 m / s; similarly, They are respectively the aerodynamic lift and aerodynamic drag at any moment from the initial moment to the moment when the airspeed reaches 10m / s.

[0056] When the engine fails and the airspeed drops to V c , so that the rotors start to rotate and generate pulling force. The equations for solving the deceleration of the drone are used to verify the magnitude of various forces / torques required for balanced flight after an engine failure occurs. If the forces / torques are unbalanced, the drone’s posture will not be stabilized. The forces / torques mentioned include L, D, T and M y .

[0057] Aerodynamic lift, drag and pitching moment are not only related to the deflection of the rudder surface, but also closely related to the angle of attack and sideslip angle. Their mathematical expressions can be expressed as follows:

[0058]

[0059] Where ρ is the air density; S ref is the reference area of ​​the wing; V is the true airspeed in the body coordinate system; C D and C L are the aerodynamic drag coefficient and aerodynamic lift coefficient respectively; C my is the aerodynamic pitch moment coefficient; c A is the aerodynamic chord length; δ A ,δ E ,δ R are the deflections of the aileron, elevator and rudder after weight distribution; β is the sideslip angle.

[0060] Step S2: dynamically adjust the control mechanism according to the flight airspeed after the engine failure occurs. In the early stage of the failure, since the UAV airspeed is high, continue to use the rudder control to maintain the stability of the flight attitude; when the airspeed is less than or equal to V c and greater than V b When the airspeed is equal to or less than V b When the airspeed is less than V a When the airspeed is greater than or equal to V a When , the weight of the rudder surface is 100%. The weight switching equation with airspeed as the constraint condition is:

[0061]

[0062] Among them, K u (V) is the rotor weight switching equation, K δ (V) is the control surface weight switching equation.

[0063] In order to balance the force and torque relationship of the UAV, a control quantity distribution matrix between the rotor and the rudder is established:

[0064]

[0065] In the formula, δ=[δ A ,δ E ,δ R ,0] T represents the fixed-wing control quantity after weight distribution, U=[U A ,U E ,U R ,U T ] T represents the rotor control quantity after weight distribution, δ′=[δ′ A ,δ′ E ,δR ′,0] T represents the fixed-wing control quantity without weight assignment, U′=[U′ A ,U′ E ,U′ R ,U′ T ] T represents the unweighted rotor control quantity, δ A and δ A ' A They represent the weighted and unweighted aileron deflections, δ E and δ E ' E Denote the weighted and unweighted elevator deflections, δ R and δ R ′ represents the rudder deflection of the rudder after weight allocation and the rudder without weight allocation respectively, 0 represents the engine throttle (0 due to engine failure); U A and U′ A are the weighted and unweighted rotor roll control quantities, respectively, U E and U′ E are the weighted and unweighted pitch control quantities, U R and U′ R are the yaw control quantities after weight assignment and without weight assignment, U T and U′ T are the rotor throttle control quantities after weight allocation and without weight allocation respectively; K δ K δ (V); K u K u (V).

[0066] H and φ, θ, ψ are the real-time altitude and attitude angle information of the UAV respectively; H g and φ g ,θ g ,ψ g are the target height and target attitude angle of the UAV respectively; is the control quantity allocation function, K p , K d , K I They are proportional control parameters, differential control parameters, and integral control parameters (and different Δ may have different K p , K d , K I , K corresponding to each Δ p , K d , K IBoth are adjustable and determined according to actual conditions), Δ is the deviation between the real-time status information and the target status information; t is time; the real-time status information includes the real-time altitude and real-time attitude angle information; the target status information includes the target altitude and target attitude angle information.

[0067] The target height and target attitude angle information of the drone are preset, and the real-time status of the drone is measured by the drone sensor. X, Y, H are the real-time position of the drone, φ, θ, ψ are the real-time attitude angles of the drone, where φ is the real-time roll angle, θ is the pitch angle of the aircraft (which is also measured in real time), and ψ is the real-time yaw angle. Among them, X, Y, H are the real-time position of the drone in the inertial coordinate system, H represents the real-time height, and X and Y are the positions in two mutually perpendicular directions in the horizontal plane of the inertial coordinate system. In this article, except for X, Y, H, X g ,Y g ,H g Except for , all other parameters correspond to the body coordinate system. Regarding the position information, the height H is used in formula (4).

[0068] In formula (4), HH g is ΔH (height deviation); φ-φ g is Δφ (roll angle deviation); θ-θ g is Δθ (pitch angle deviation); ψ-ψ g is Δψ (yaw angle deviation); Δ includes ΔH, Δφ, Δθ, and Δψ.

[0069] Step S3: Calculate the real-time aerodynamic lift L, real-time aerodynamic drag D, and real-time aerodynamic pitch moment M by distributing the control amount between the rotor and the control surface y , the sum of the thrust T generated by the real-time rotor, and then substitute it into the deceleration equation of the drone in step S1. When the real-time L, D, M y , T makes the UAV decelerate and solve the equation, then the composite wing UAV can achieve attitude stability when the engine fails.

[0070] According to the real-time rotor control quantity and rudder control quantity obtained by the control quantity distribution matrix between the rotor and the rudder, the real-time L, D, M are calculated. y , whether T satisfies the UAV deceleration solution equation of formula (1).

[0071] Specifically, according to formula (4), δ A ,δ E ,δ R , δ A ,δ E ,δ R Substituting into formula (2), we can calculate L, D, M y According to formula (4), U A ,UE ,U R ,U T The rotor speed ω is obtained by mapping the matrix B. Assuming that the UAV is a six-rotor UAV, ω1, ω2, ω3, ω4, ω5, and ω6 are the speeds of the six rotors, respectively, satisfying the following formula:

[0072]

[0073] The above formula can be used to calculate the rotation speed of each rotor, and then calculate the sum of the pulling forces generated by the rotor. K t is the tensile coefficient. y Enter it together with T into the equation (1) to solve the UAV deceleration equation.

[0074] If the rotor control quantity and rudder control quantity calculated at a certain moment after the UAV fails, the corresponding real-time L, D, M y and T, and the first time the deceleration equation of the UAV is established, it is determined that the posture has achieved stable control, and the UAV continues to control the UAV with the rotor control amount and the rudder control amount when the deceleration equation of the UAV is established.

[0075] According to simulation, usually within 30 seconds after the UAV fails, the attitude stabilization control method of the composite wing UAV engine failure of the present invention can satisfy the UAV deceleration solution equation of formula (1), that is, achieve attitude stability.

[0076] The Simulink simulation environment is used to simulate and verify the stability control after engine failure to ensure the robustness and safety of the proposed method. Figure 2 A schematic diagram of a UAV suddenly experiencing an engine failure during normal flight, resulting in power interruption and airspeed decay is given; Figure 3 and Figure 4 The attitude angle change curve and position tracking curve from engine failure to stable hovering are given; Figure 5 The corresponding schematic diagram of rotor control quantity and rudder control quantity is given.

[0077] The present invention designs an attitude stabilization control strategy based on airspeed threshold in response to sudden engine failures that may occur during normal operation of the UAV. This strategy dynamically adjusts the distribution of force and torque through the synergy of aerodynamic control surfaces and rotors, thereby achieving safe hovering and flight attitude stability of the UAV under fault conditions. The attitude stabilization control strategy based on airspeed threshold proposed in the present invention can effectively coordinate the control force and torque generated by the rotor and deflection control surfaces according to the airspeed before the fault occurs, stabilize the flight attitude and position of the UAV, and enhance the redundant fault tolerance capability when a fault occurs, thereby ensuring the flight safety and controllability of the UAV under engine failure conditions.

[0078] The above is only a preferred embodiment of the present invention, and is not intended to limit the scope of implementation of the present invention. Several improvements and modifications may be made without departing from the principles of the present invention, and these improvements and modifications should also be regarded as the protection scope of the present invention.

Claims

1. A method for attitude stabilization control of a composite wing UAV under engine failure, characterized in that: The method comprises: Step S1, according to the force analysis after the composite wing UAV engine failure occurs, establish the UAV deceleration solution equation including the forward inequality, the vertical balance equation and the pitch moment balance equation; the force includes the aerodynamic lift L, the aerodynamic drag D, and the sum of the pull generated by the rotor T; Step S2: according to the real-time airspeed after the engine failure occurs, a weight switching equation and a control amount distribution matrix with airspeed as a constraint condition are established to coordinate the control amount distribution between the rotor and the control surface; Step S3: Calculate the real-time aerodynamic lift L, the real-time aerodynamic drag D, and the real-time aerodynamic pitch moment M by distributing the control amount between the rotor and the control surface. y , the sum of the thrust T generated by the real-time rotor, and then substitute it into the deceleration equation of the drone in step S1. When the real-time L, D, M y , T makes the UAV decelerate and solve the equation, then the composite wing UAV can achieve attitude stability when the engine fails.

2. The method for attitude stabilization control of a composite wing UAV under engine failure according to claim 1, characterized in that: In step S1, the deceleration equation of the drone is composed of the forward inequality, the vertical balance equation and the pitch moment balance equation: In the formula, s (·) = sin(·),c (·) = cos(·); T is the sum of the thrusts generated by the rotor; L and D represent the aerodynamic lift and aerodynamic drag, respectively; θ and α are the pitch angle and angle of attack of the aircraft, respectively; M y is the aerodynamic pitch moment; l is the distance between the projection of the rotor on the symmetry plane of the aircraft and the center of gravity of the UAV; t Vc →∞ indicates that the rotor thrust changes from t Vc appears at the moment and exists forever; 0→t Va Indicates that it existed before and is in t Va Moment disappears; V a V is the airspeed at which the aerodynamic force decreases to negligible level; c The airspeed at which flight controls are added to the rotor.

3. The method for attitude stabilization control of a composite wing UAV under engine failure according to claim 1 is characterized in that: In step S1, the expressions of aerodynamic lift, drag and pitching moment are: Where ρ is the air density; S ref is the reference area of ​​the wing; V is the true airspeed in the body coordinate system; C D and C L are the aerodynamic drag coefficient and aerodynamic lift coefficient respectively; C my is the aerodynamic pitch moment coefficient; c A is the aerodynamic chord length; δ A ,δ E ,δ R are the deflections of the aileron, elevator and rudder after weight distribution; β is the sideslip angle.

4. The method for attitude stabilization control of a composite wing UAV under engine failure according to claim 1, characterized in that: In step S2, the airspeed is used as the limiting condition. When the airspeed V is greater than V c When the airspeed V is less than or equal to V c and greater than V b When the airspeed V is equal to V b or less than V b When the airspeed V is less than V a When the airspeed is greater than or equal to V a When , the rudder weight is 100%; The weight switching equation with airspeed as the constraint is: Among them, K u (V) is the rotor weight switching equation, K δ (V) is the control surface weight switching equation.

5. The method for attitude stabilization control of a composite wing UAV under engine failure according to claim 4 is characterized in that: In step S2, in order to balance the force and torque relationship of the UAV, the control quantity distribution matrix between the rotor and the control surface is established: In the formula, δ=[δ A ,δ E ,δ R ,0] T represents the fixed-wing control quantity after weight distribution, U=[U A ,U E ,U R ,U T ] T represents the rotor control quantity after weight distribution, δ′=[δ′ A ,δ′ E ,δ R ′,0] T represents the fixed-wing control quantity without weight assignment, U′=[U′ A ,U′ E ,U′ R ,U′ T ] T represents the unweighted rotor control quantity, δ A and δ A ' A They represent the weighted and unweighted aileron deflections, δ E and δ E ' E Denote the weighted and unweighted elevator deflections, δ R and δ R ′ respectively represent the rudder deflection after weight allocation and the rudder without weight allocation, 0 represents the engine throttle; U A and U′ A are the weighted and unweighted rotor roll control quantities, respectively, U E and U′ E are the weighted and unweighted pitch control quantities, U R and U′ R are the yaw control quantities after weight assignment and without weight assignment, U T and U′ T are the rotor throttle control quantities after weight allocation and without weight allocation respectively; K δ K δ (V); K u K u (V); H and φ, θ, ψ are the real-time altitude and attitude angle information of the UAV respectively; H g and φ g ,θ g ,ψ g are the target height and target attitude angle of the UAV respectively; is the control quantity allocation function, K p , K d , K I are proportional control parameter, differential control parameter and integral control parameter respectively, Δ is the deviation between real-time state information and target state information; real-time state information includes real-time height and real-time attitude angle information; target state information includes target height and target attitude angle information.

6. A composite wing UAV engine failure attitude stabilization control system, characterized in that: include: The UAV deceleration solution module establishes the UAV deceleration solution equations including the forward inequality, vertical balance equation and pitch moment balance equation based on the force analysis after the composite wing UAV engine failure occurs; The weight switching module establishes the rotor weight switching equation and the control surface weight switching equation with airspeed as the constraint condition; The control quantity allocation module for the rudder and rotor establishes the control quantity allocation matrix between the rotor and the rudder, and completes the control quantity allocation between the rotor and the rudder according to the deviation between the real-time altitude and real-time attitude angle information and the target altitude and target attitude angle information, as well as the control quantity allocation matrix.

7. The attitude stabilization control system for composite wing UAV under engine failure according to claim 6, characterized in that: In the weight switching module, the airspeed is used as the limiting condition. When the airspeed V is greater than V c When the airspeed V is less than or equal to V c and greater than V b When the airspeed V is equal to V b or less than V b When the airspeed V is less than V a When the airspeed is greater than or equal to V a , the rudder weight is 100%.

8. A processor for running a computer program, characterized in that: When the computer program is running, the method for attitude stabilization control under engine failure of a composite wing UAV as described in any one of claims 1 to 5 is executed.

9. A drone, characterized in that: It includes the attitude stabilization control system under engine failure of the composite wing UAV as described in claim 6 or 7.

10. A computer readable medium having a computer program stored thereon, characterized in that: The computer program is executed by the processor to implement the method for attitude stabilization control under engine failure of a composite wing UAV as described in any one of claims 1 to 5.

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