A control allocation method for a distributed propulsion tiltrotor aircraft

Abstract

The application discloses a control distribution method of a distributed propulsion tilt-rotor aircraft, and comprises the following steps: constructing an aircraft control efficiency model; solving the aircraft control efficiency model based on a force vector quadratic distribution method; limiting the solution result of the aircraft propeller based on a multi-stage generalized inverse actuator constraint strategy; introducing a variable relaxation factor, and performing smooth processing on the limited solution result by using a Shur complement theorem; and controlling the distributed propulsion tilt-rotor aircraft according to the smoothed solution result. By using the application, the high-dimensional matrix operation caused by a large number of redundant propellers can be avoided, the aircraft is not prone to instability, and the aircraft can realize "no limit" speed regulation from a rotor mode to a fixed-wing mode. The application can be widely applied to the field of aircraft control technology.

Description

Technical Field

[0001] This invention relates to the field of aircraft control technology, and in particular to a control allocation method for a distributed propulsion tiltrotor aircraft. Background Technology

[0002] UAVs with vertical take-off and landing capabilities can complete reconnaissance and patrol missions in complex terrain environments because they do not require runways or special auxiliary take-off and landing methods. They have broad application prospects. Traditional helicopters rely on rotor power to provide all the lift required by the aircraft, which limits their flight speed and flight distance, making it difficult to meet the requirements of long-distance and long-endurance missions. Tiltrotor vertical take-off and landing UAVs combine the vertical take-off and landing capabilities of rotorcraft with the high-speed and long-endurance flight capabilities of fixed-wing aircraft, enabling vertical take-off and landing, high-speed and long-endurance flight missions.

[0003] Due to its higher propulsion efficiency, higher reliability, and lower operating noise, distributed propulsion technology is gradually being applied to large-payload VTOL aircraft. Distributed electric propulsion decouples the traditional turbine power system and fan thrust system, allowing both to operate at optimal speeds. Currently, control algorithms for UAVs mainly focus on their position and attitude control. The outputs of these controllers are virtual control quantities, which need to be distributed to each thruster through control allocation methods. However, existing technologies for tiltrotor aircraft control are mostly suitable for aircraft with dual-power systems, but not for the control of distributed propulsion VTOL aircraft. While current tiltrotor propulsion systems have achieved full-mode flight of tiltrotor aircraft, they have not addressed the control allocation problems caused by highly redundant thrusters. Summary of the Invention

[0004] To address the aforementioned technical problems, the present invention aims to provide a control allocation method for a distributed propulsion tiltrotor aircraft, which can avoid the instability caused by high-dimensional matrix operations due to a large number of redundant propellers and achieve "stepless" speed regulation of the aircraft from rotor mode to fixed-wing mode.

[0005] The first technical solution adopted in this invention is: a control allocation method for a distributed propulsion tiltrotor aircraft, comprising the following steps:

[0006] Based on the position of the aircraft thrusters relative to the aircraft body coordinate system, a control efficiency model for the aircraft is constructed.

[0007] The control efficiency model of the aircraft is solved based on the force vector quadratic allocation method, and the solution results of the aircraft thruster are obtained.

[0008] The solution results for the aircraft thruster are constrained by the multi-level generalized inverse actuator constraint strategy, and the constrained solution results are obtained.

[0009] By introducing a variable relaxation factor and smoothing the constrained solution results using the Schuler complement theorem, a smoothed solution result is obtained.

[0010] The distributed propulsion tiltrotor aircraft is controlled based on the smoothed solution results.

[0011] Furthermore, the expression for the aircraft control efficiency model is as follows:

[0012]

[0013] In the above formula, M r F represents the desired torque given by the upper-level controller. r This represents the desired thrust vector given by the upper-level controller, s represents the number of tilt servos, and f represents the thrust vector. ij n represents the thrust generated by the j-th thruster on the i-th wing. i This represents the position vector of the i-th wing in the body coordinate system.

[0014] Furthermore, the step of solving the aircraft control efficiency model based on the force vector quadratic allocation method to obtain the solution result of the aircraft thruster specifically includes:

[0015] Set L i Substitute the values ​​into the aircraft control efficiency model and perform transformation processing to obtain the transformed aircraft control efficiency model;

[0016] The control torque command is distributed to each wing, the antisymmetric matrix of the wing position vector is obtained, and a control efficiency model based on each wing is constructed.

[0017] The control efficiency model of each wing is solved to obtain the total thrust of the aircraft propulsion system on each wing;

[0018] The total thrust of the aircraft propulsion units on each wing is distributed to each aircraft propulsion unit to construct a control efficiency model for the wing;

[0019] The control efficiency model of the wing is solved by generalized inverse solving to obtain the solution results of the aircraft thruster. The solution results of the aircraft thruster include the thrust and tilt angle of the aircraft thruster.

[0020] Furthermore, the expression for the solution result of the aircraft thruster is as follows:

[0021]

[0022]

[0023] In the above formula, F iLet α represent the total thrust of the aircraft's propulsion system on the i-th wing. i This represents the tilt angle of the aircraft thruster on the i-th wing. F represents i In the body coordinate system X b Components of the axis, F represents i In the body coordinate system Z b The components of the axis.

[0024] Furthermore, the step of restricting the solution results of the aircraft thruster based on the multi-level generalized inverse actuator constraint strategy to obtain the restricted solution results specifically includes:

[0025] The body coordinate system is rotated to obtain the rotated body coordinate system.

[0026] Based on the rotated body coordinate system, the desired tilt angle of the aircraft thruster is obtained;

[0027] By performing a multi-level generalized inverse solution on the desired tilt angle of the aircraft thruster using the total thrust of the aircraft thruster, the relationship between the total thrust of the aircraft thruster and the desired tilt angle of the aircraft thruster can be obtained.

[0028] Based on the relationship between the total thrust of the aircraft thruster and the desired tilt angle of the aircraft thruster, a control efficiency model of the aircraft in the rotated body coordinate system is constructed.

[0029] The aircraft control efficiency model in the rotated body coordinate system is simplified by linear constraint processing to obtain the simplified aircraft control efficiency model.

[0030] The simplified aircraft control efficiency model is solved by multi-level generalized inverse solution until the number of solutions meets the preset number of solutions, and the solution result of multi-level generalized inverse is obtained.

[0031] The results of solving the multi-level generalized inverse are accumulated to obtain the restricted solution.

[0032] Furthermore, the expression for the aircraft control efficiency model in the rotated body coordinate system is as follows:

[0033]

[0034]

[0035]

[0036] In the above formula, M represents r With F rIn the rotated body coordinate system, the desired torque and the desired triaxial tension, Q f Evidence is provided to demonstrate control efficiency in the rotated machine coordinate system. The weight matrix in the rotated body coordinate system In the k-th iteration, the tension vector of each thruster in the rotated body coordinate system. The tension on the i-th wing along the X-axis in the rotated body coordinate system. f The tension of the shaft, This represents the tension on the i-th wing along the Z-axis in the rotated body coordinate system. f The tension of the shaft, This represents the maximum thrust that the thruster on the i-th wing can provide. This represents the minimum tilt angle of the i-th wing in the rotated fuselage coordinate system. This represents the maximum tilt angle of the i-th wing in the rotated body coordinate system.

[0037] Furthermore, the expression for the solution result after the constraints is as follows:

[0038]

[0039]

[0040]

[0041] In the above formula, This represents the final calculated thrust vector of each thruster in the rotated body coordinate system. Let represent the tension vector of each thruster in the rotated body coordinate system, obtained from the k-th iteration. This represents the thrust vector of the i-th thruster obtained from the final solution in the rotated body coordinate system. f Components of the axis, This represents the thrust vector of the i-th thruster obtained from the final solution in the rotated body coordinate system, Z. f The axis component, α, represents the desired tilt angle.

[0042] Furthermore, the step of introducing a variable relaxation factor and smoothing the constrained solution result using the Schulze theorem to obtain a smoothed solution result specifically includes:

[0043] The aircraft control efficiency model in the rotated body coordinate system is transformed to obtain the transformed aircraft control efficiency model.

[0044] Find the generalized inverse matrix of the transformed aircraft control efficiency model;

[0045] The generalized inverse matrix of the transformed aircraft control efficiency model is solved using the Schur complement theorem to obtain the aircraft control efficiency matrix.

[0046] By introducing a variable relaxation factor, the control efficiency matrix of the aircraft is smoothed to obtain the control allocation matrix of the aircraft.

[0047] The aircraft control allocation matrix is ​​solved to obtain a smoothed solution.

[0048] Furthermore, the expression for the transformed aircraft control efficiency model is as follows:

[0049]

[0050] In the above formula, This represents the control efficiency matrix after matrix partitioning. This represents the control efficiency matrix in the rotated body coordinate system. This represents the weight matrix in the rotated body coordinate system. express The first matrix block of the matrix, express The second matrix block of the matrix.

[0051] Furthermore, the expression for the aircraft control allocation matrix is ​​as follows:

[0052]

[0053] In the above formula, A k This represents the control allocation matrix after introducing the relaxation factor λ. This represents the control efficiency matrix after matrix partitioning, where λ represents the relaxation factor. express The Shure complement matrix, The inverse of the top left matrix block.

[0054] The beneficial effects of the method of this invention are as follows: This invention solves the control efficiency model of the aircraft by using a force vector-based quadratic allocation method. That is, the control torque command of the upper controller is first allocated to each wing, and then the control torque command is secondary allocated from each wing to each thruster. This avoids the additional constraints caused by the high-dimensional matrix operation of a large number of redundant thrusters, which may lead to instability of the aircraft. Furthermore, the idea of ​​the generalized inverse of the servo is applied to the force vector control allocation method, which realizes the processing of the limitations of thruster thrust and tilt angle, thereby achieving better control effect. The generalized inverse of the servo is solved by Schur complement, and a variable relaxation factor is added in the solution process to achieve a smoothing effect, weaken the boundary between rotor mode and fixed wing mode, realize arbitrary transition between the two modes, and thus realize "stepless" speed regulation from rotor mode to fixed wing mode. Attached Figure Description

[0055] Figure 1 This is a flowchart of the steps of a control allocation method for a distributed propulsion tiltrotor aircraft according to the present invention;

[0056] Figure 2 This is a schematic diagram of the distributed propulsion tiltrotor aircraft constructed according to the present invention;

[0057] Figure 3 This is a simulated schematic diagram of the distributed propulsion tiltrotor aircraft constructed according to the present invention;

[0058] Reference numerals: 1. Wing; 2. Propeller; 3. Tilting servo. Detailed Implementation

[0059] The present invention will now be described in further detail with reference to the accompanying drawings and specific embodiments. The step numbers in the following embodiments are only for ease of explanation and do not limit the order of the steps. The execution order of each step in the embodiments can be adapted according to the understanding of those skilled in the art.

[0060] Reference Figures 1 to 3 This invention provides a control allocation method for a distributed propulsion tiltrotor aircraft, the method comprising the following steps:

[0061] S1. Construct an aircraft control efficiency model;

[0062] Specifically, the torque commands generated by the upper-level controller are first distributed to each wing, and then the torque of each wing is further distributed to each thruster. Based on this idea, the positions of each wing thruster from the body coordinate system are as follows:

[0063] n ij =[N i ,l j H i] T j = 1, ..., p i

[0064] In the above formula, n ij p represents the position vector of the j-th thruster on the i-th wing in the body coordinate system. i N represents the total number of thrusters on the i-th wing. i This indicates that the i-th wing is along the X coordinate system of the fuselage. b The distance from the center of gravity of the aircraft along the axial direction, l j This indicates that the j-th thruster on the i-th wing is positioned along the Y-axis in the body coordinate system. b The distance H from the center of gravity along the axial direction. i This represents the height of the i-th wing relative to the center of gravity in the body coordinate system;

[0065] The control efficiency model of the aircraft can then be written as follows:

[0066]

[0067] f ij =[f ij cα i ,0,-f ij sα i ] T

[0068] In the above formula, M r F represents the desired torque given by the upper-level controller. r This represents the desired thrust vector given by the upper-level controller, s represents the number of tilt servos, and f represents the thrust vector. ij n represents the thrust generated by the j-th thruster on the i-th wing. i cα represents the position vector of the i-th wing in the body coordinate system. i cosα i sα i sinα i α i This represents the tilt angle of the i-th wing in the body coordinate system.

[0069] S2. Solve the aircraft control efficiency model based on the force vector quadratic allocation method to obtain the solution results of the aircraft thruster;

[0070] S21, First allocation method based on force vector;

[0071] Specifically, a method based on the quadratic allocation of force vectors is adopted, assuming...

[0072]

[0073] In the above formula, Li The equivalent point of action of the thruster of the i-th wing is located at a distance l from the fuselage plane of symmetry in the body coordinate system. ij This indicates that the j-th thruster on the i-th wing is positioned along the Y-axis in the body coordinate system. b The distance from the axis of rotation to the plane of symmetry of the fuselage, p i This represents the total number of thrusters on the i-th wing;

[0074] set up

[0075] r i =[N i L i H i ] T

[0076] And because

[0077] n ij =[N i ,l ij H i ] T j = 1, ..., p i

[0078] Then there is

[0079]

[0080] Then based on L i The control efficiency model of the aircraft is transformed to obtain the transformed aircraft efficiency model, the expression of which is shown below:

[0081]

[0082] In the above formula, r s F represents the position vector of the equivalent point of action of the s-th wing in the body coordinate system. s R represents the tension vector at the equivalent point of application of the s-th wing. s Indicates r s The antisymmetric matrix spanned by I 3×3 represents a 3D identity matrix, and s represents the number of wings;

[0083] in,

[0084]

[0085] r i =[N i L i H i ] T

[0086] In the above formula, r iThis represents the position vector of the i-th wing thrust equivalent point. and N represents the thrust vector at the i-th equivalent point of action of the wing. i X represents the position vector of the i-th wing's equivalent point of action in the body coordinate system. b Axial component, L i Y represents the position vector of the i-th wing equivalent point of action in the body coordinate system. b Axial component, H i Z represents the position vector of the i-th wing's equivalent point of action in the body coordinate system. b Axial components;

[0087] in For position vector r i The corresponding antisymmetric matrix is ​​used to construct the aircraft control allocation model, and its expression is shown below:

[0088]

[0089] In the above formula, the symbol Represents the generalized inverse of a matrix;

[0090] Among them, let

[0091]

[0092] In the above formula, L i The equivalent point of action of the thruster of the i-th wing is located at a distance l from the fuselage plane of symmetry in the body coordinate system. ij This indicates that the j-th thruster on the i-th wing is positioned along the Y-axis in the body coordinate system. b The distance from the axis of rotation to the plane of symmetry of the fuselage, p i This represents the total number of thrusters on the i-th wing;

[0093] Based on the aircraft control distribution model, the total thrust and tilt angle of the i-th wing thruster can be calculated as follows:

[0094]

[0095]

[0096] In the above formula, F i Let α represent the total thrust of the aircraft's propulsion system on the i-th wing. i This represents the tilt angle of the aircraft thruster on the i-th wing. F represents i In the body coordinate system X b Components of the axis, F represents i In the body coordinate system Z b The components of the axis;

[0097] S22, A second allocation method based on force vectors;

[0098] Specifically, a second allocation calculation is performed based on the expression for the total thrust and tilt angle of the i-th wing thruster, and F is allocated... i The control efficiency model for the i-th wing is listed below, with the thrusters further allocated to the i-th wing:

[0099]

[0100] By taking the generalized inverse of the equation, the thrust of each thruster can be obtained as follows:

[0101]

[0102] S3. The solution results of the aircraft thruster are restricted based on the multi-level generalized inverse actuator constraint strategy to obtain the restricted solution results;

[0103] Specifically, in step S2, although the complete control allocation process is completed, if the actuator becomes saturated during the allocation process, the generated force and torque may not meet the expectations, or even lead to instability. Therefore, the constraints on the actuator are taken into consideration during the allocation process.

[0104] S31. Construct an aircraft control efficiency model based on the transformed body coordinate system;

[0105] Specifically, a new coordinate system F is introduced based on the body coordinate system. f :{x f ,y f ,z f}, new coordinate system F f It can be regarded as being based on the body coordinate system F B Around x B The angle obtained by rotating the axis by an angle α, where α is the expected tilt angle, can be obtained from the x-axis in the body coordinate system. b axis and z b The expected thrust in the axial direction is calculated as follows:

[0106] a=atan2(X r Z r )

[0107] In the above formula, α represents the expected tilt angle, and X r Represents X in the body coordinate system b Desired force in the axial direction, Z r Z represents the coordinate system of the machine body. b Desired force in the axial direction;

[0108] In Ff The following applies the multi-level generalized inverse (CGI) method, which is a multi-iteration method where each iteration is equivalent to solving the generalized inverse. For the k-th iteration, the control efficiency model can be transformed into a new coordinate system F. f The following is a representation:

[0109]

[0110]

[0111]

[0112] In the above formula, M represents r With F r In the rotated body coordinate system, the desired torque and the desired triaxial tension, Q f Evidence is provided to demonstrate control efficiency in the rotated machine coordinate system. The weight matrix in the rotated body coordinate system In the k-th iteration, the tension vector of each thruster in the rotated body coordinate system. The tension on the i-th wing along the X-axis in the rotated body coordinate system. f The tension of the shaft, This represents the tension on the i-th wing along the Z-axis in the rotated body coordinate system. f The tension of the shaft, This represents the maximum thrust that the thruster on the i-th wing can provide. This represents the minimum tilt angle of the i-th wing in the rotated fuselage coordinate system. This represents the maximum tilt angle of the i-th wing in the rotated body coordinate system;

[0113] Where, Δα i For the i-th tilt servo relative to F f z below f The deflection angle of the axis, and there exists

[0114]

[0115]

[0116]

[0117] In the above formula, R α This represents the rotation matrix that rotates the machine from the body coordinate system by an angle α. R represents the desired force in the rotated body coordinate system. s Indicates r s For antisymmetric arrays, rs This represents the position vector of the equivalent point of action of the s-th wing in the rotated body coordinate system;

[0118] Thus, the control efficiency model and its constraints in F f It is expressed in the system and is easy to obtain in F. f In the middle, z f The torque in the direction is mainly generated by the deflection angle Δα. i Achieve, and x f With z f The directional torque is mainly achieved by adjusting the thrust f of each propeller. ij This can be achieved, therefore, a priority can be set, and the nonlinear constraints in the formula can be simplified to linear constraints;

[0119] S32. In the force vector-based control allocation method, CGI is applied to allocate the force of each wing. The results of each iteration are accumulated to obtain the final allocated thrust and tilt angle.

[0120] Specifically, the nonlinear constraints in the aircraft control efficiency model based on the transformed body coordinate system are simplified into linear constraints, and the transformation formula is shown below:

[0121]

[0122]

[0123]

[0124]

[0125] In the force vector-based control allocation method, CGI is applied. For the k-th iteration, the allocated forces for each wing are as follows:

[0126]

[0127]

[0128]

[0129] In the above formula, express The value after saturation treatment, Q f This represents the control allocation matrix in the rotated body coordinate system;

[0130] in,

[0131]

[0132]

[0133] In the above formula, This represents the weight matrix for the k-th iteration in the rotated body coordinate system;

[0134] Calculate the number of iterations. When the number of iterations is greater than 8 or when... When the value falls below a certain threshold, the iteration ends, and the final allocation value is obtained. The sum of the results of each iteration is the final allocated thrust and tilt angle, as shown below:

[0135]

[0136]

[0137]

[0138] In the above formula, N represents the total number of iterations. This represents the thrust vector of the i-th wing in the rotated body coordinate system at x. f Components of the axis, This represents the thrust vector of the i-th wing in the rotated body coordinate system at z. f The components of the axis;

[0139] Finally, F in the formula i By continuing to distribute the solution to each thruster according to the formula, the constrained solution can be obtained.

[0140] S4. Introduce a variable relaxation factor and smooth the constrained solution results using the Schuler complement theorem to obtain a smoothed solution result.

[0141] Specifically, in some distributed propulsion vertical takeoff and landing (VTOL) UAVs, the thrusters are embedded in the trailing edge of the wing. The tilt angle of the thrusters is often strictly limited, generally not exceeding 90° to ensure the wing's takeoff performance. Therefore, in fixed-wing mode, it is often impossible to simultaneously satisfy the desired torque and z-force. b The expected value of the directional force, since in fixed-wing modes, generally does not pass through z. b Altitude is controlled by directional forces; therefore, once the aircraft enters fixed-wing mode, the x-axis can be deactivated. f The expectation of directional force, if directly canceled during the tilting process, can easily cause jumps in the propeller and tilting mechanism. Repeated jumps can easily damage these actuators.

[0142] S41. Swap the last two lines in the aircraft control efficiency model to obtain the transformed aircraft control efficiency model.

[0143] Specifically, firstly, the last two lines of the control efficiency model are swapped so that f zThe directional constraint is located in the last row of the control efficiency model, i.e., the last row of the control efficiency matrix. The transformed control efficiency matrix is ​​denoted as follows:

[0144]

[0145] in and They are respectively Given the first four rows and the last row, the generalized inverse of the aircraft control efficiency matrix can be expressed in the following form:

[0146]

[0147] S42. Solve for the generalized inverse of the aircraft control efficiency matrix in step S41 using the form of Schur complement;

[0148] Specifically, the solution steps based on Schur complement are as follows:

[0149]

[0150] In the above formula, The main matrix representing the inverse of the Schur complement process, Represents the Schul complement matrix;

[0151] in,

[0152]

[0153]

[0154]

[0155] The above expression, except for the last row of the control efficiency matrix (i.e., for x), f The impact of the expected force in the direction on the overall control efficiency model, if x is cancelled. f If the expected force in the direction is given, then the control allocation matrix is: If x is added f The expected force in the direction, the control efficiency matrix is

[0156] S43. Introduce a variable relaxation factor for smooth transition.

[0157] Furthermore, it is possible to do so in x f A variable relaxation factor is added to the expected force in the direction, such that during the transition, x f The expected force in the direction can be gradually and smoothly cancelled, and the control allocation matrix is ​​represented as follows:

[0158]

[0159] In the above formula, A k This represents the control allocation matrix after introducing the relaxation factor λ. This represents the control efficiency matrix after matrix partitioning, where λ represents the relaxation factor. express The Shure complement matrix, The inverse of the top-left matrix block;

[0160] make

[0161]

[0162] Therefore, a variable relaxation factor is introduced, and the solution results after constraints are smoothed using the Schuler complement theorem, resulting in a smoothed solution.

[0163] The above is a detailed description of the preferred embodiments of the present invention. However, the present invention is not limited to the embodiments described. Those skilled in the art can make various equivalent modifications or substitutions without departing from the spirit of the present invention. All such equivalent modifications or substitutions are included within the scope defined by the claims of this application.

Claims

1. A control allocation method for a distributed propulsion tiltrotor aircraft, characterized in that, Includes the following steps: Based on the position of the aircraft thrusters relative to the aircraft body coordinate system, a control efficiency model for the aircraft is constructed. The expression for the aircraft control efficiency model is as follows: In the above formula, This represents the desired torque given by the upper-level controller. This represents the desired thrust vector given by the upper-level controller. Indicates the number of tilt servos. Indicates the first The first on the wing The magnitude of thrust generated by each thruster Indicates the first The position vector of each wing in the body coordinate system; based on The aircraft control efficiency model is transformed to obtain the transformed aircraft control efficiency model. For the first The equivalent point of action of the thruster of each wing is the distance from the fuselage plane of symmetry in the body coordinate system; the control torque command is distributed to each wing, the antisymmetric matrix of the wing position vector is obtained, and a control efficiency model based on each wing is constructed. The control efficiency model of each wing is solved to obtain the total thrust of the aircraft propulsion system on each wing; The total thrust of the aircraft propulsion units on each wing is distributed to each aircraft propulsion unit to construct a control efficiency model for the wing; The control efficiency model of the wing is solved by generalized inverse solving to obtain the solution results of the aircraft thruster. The solution results of the aircraft thruster include the thrust of the aircraft thruster and the tilt angle of the aircraft thruster. The solution results for the aircraft thruster are constrained by the multi-level generalized inverse actuator constraint strategy, and the constrained solution results are obtained. By swapping the last two rows of the aircraft control efficiency model in the rotated body coordinate system, we obtain the transformed aircraft control efficiency model. The expression for the transformed aircraft control efficiency model is shown below: In the above formula, This represents the control efficiency matrix after matrix partitioning. This represents the control efficiency matrix in the rotated body coordinate system. This represents the weight matrix in the rotated body coordinate system. express The first matrix block of the matrix, express The second matrix block in the matrix division; Find the generalized inverse matrix of the transformed aircraft control efficiency model; The generalized inverse matrix of the transformed aircraft control efficiency model is solved using the Schur complement theorem to obtain the aircraft control efficiency matrix. By introducing a variable relaxation factor, the control efficiency matrix of the aircraft is smoothed to obtain the control allocation matrix of the aircraft. The expression for the aircraft control allocation matrix is ​​as follows: In the above formula, Indicates the introduction of relaxation factors The subsequent control allocation matrix, This represents the control efficiency matrix after matrix partitioning. Indicates the relaxation factor. express The Shure complement matrix, Inverse of the top left matrix block The aircraft control allocation matrix is ​​solved to obtain a smoothed solution. The distributed propulsion tiltrotor aircraft is controlled based on the smoothed solution results.

2. The control allocation method for a distributed propulsion tiltrotor aircraft according to claim 1, characterized in that, The expression for the solution result of the aircraft thruster is as follows: In the above formula, Indicates the first The total thrust of the aircraft's propulsion units on each wing, Indicates the first The tilt angle of the aircraft's propulsion system on each wing express In the body coordinate system Components of the axis, express In the body coordinate system The components of the axis.

3. The control allocation method for a distributed propulsion tiltrotor aircraft according to claim 1, characterized in that, The step of restricting the solution results of the aircraft thruster based on the multi-level generalized inverse actuator constraint strategy to obtain the restricted solution results specifically includes: The body coordinate system is rotated to obtain the rotated body coordinate system. Based on the rotated body coordinate system, the desired tilt angle of the aircraft thruster is obtained; By performing a multi-level generalized inverse solution on the desired tilt angle of the aircraft thruster using the total thrust of the aircraft thruster, the relationship between the total thrust of the aircraft thruster and the desired tilt angle of the aircraft thruster can be obtained. Based on the relationship between the total thrust of the aircraft thruster and the desired tilt angle of the aircraft thruster, a control efficiency model of the aircraft in the rotated body coordinate system is constructed. The aircraft control efficiency model in the rotated body coordinate system is simplified by linear constraint processing to obtain the simplified aircraft control efficiency model. The simplified aircraft control efficiency model is solved by multi-level generalized inverse solution until the number of solutions meets the preset number of solutions, and the solution result of multi-level generalized inverse is obtained. The results of solving the multi-level generalized inverse are accumulated to obtain the restricted solution.

4. The control allocation method for a distributed propulsion tiltrotor aircraft according to claim 3, characterized in that, The expression for the aircraft control efficiency model in the rotated body coordinate system is as follows: In the above formula, express and The desired torque and triaxial desired tension in the rotated body coordinate system. This represents the control efficiency matrix in the rotated body coordinate system. The weight matrix in the rotated body coordinate system No. In the next iteration, the tension vector of each thruster in the rotated body coordinate system. No. The tension on each wing along the lower edge of the rotated body coordinate system The tension of the shaft, Indicates the first The tension on each wing along the lower edge of the rotated body coordinate system The tension of the shaft, Indicates the first The maximum thrust that the thrusters on each wing can provide. In the rotated body coordinate system, the first... The minimum tilt angle of each wing In the rotated body coordinate system, the first... The maximum tilt angle of each wing.

5. The control allocation method for a distributed propulsion tiltrotor aircraft according to claim 3, characterized in that, The expression for the solution result after the restrictions is as follows: In the above formula, This represents the final calculated thrust vector of each thruster in the rotated body coordinate system. Indicates the first The tension vectors of each thruster in the rotated body coordinate system obtained from the next iteration. This represents the final solution obtained from the first... The thrust vector of each thruster in the rotated body coordinate system Components of the axis, This represents the final solution obtained from the first... The thrust vector of each thruster in the rotated body coordinate system Components of the axis, Indicates the desired tilt angle.

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