A gust load mitigation method for large flexible aircraft

Abstract

The present invention relates to the field of aerospace technology, and in particular to a method for alleviating gust loads applicable to a large flexible aircraft. The method comprises the following steps: arranging a flight parameter sensor on the large flexible aircraft; arranging an optical fiber strain sensor on the wing of the large flexible aircraft; establishing an inner-loop control loop, wherein the inner-loop control loop performs feedback adjustment based on the local curvature of the wing and determines an inner-loop rudder control adjustment amount; establishing an outer-loop control loop, wherein the outer-loop control loop performs tracking compensation based on a measurement value of the flight parameter sensor and determines an outer-loop rudder control adjustment amount; combining the inner-loop control loop and the outer-loop control loop to establish a gust mitigation control model for the large flexible aircraft, determine a final gust mitigation rudder control adjustment amount, and control the rudder of the large flexible aircraft. The present invention can achieve aircraft gust load mitigation based on large wing deformation.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and in particular to a gust load mitigation method applicable to large flexible aircraft. Background Art

[0002] With the development of composite materials and aircraft design technologies, high-aspect-ratio, highly flexible aircraft are gaining widespread application. These aircraft exhibit low stiffness and significant elastic effects. Typically, these aircraft possess low structural weight and a high lift-to-drag ratio. Disturbances during flight can affect the aircraft's rigid-body motion and induce elastic vibrations in the structure, impacting flight performance and safety. Strong gusts and turbulence can impose significant additional aerodynamic forces and moments on highly flexible aircraft, shortening the fatigue life of the aircraft structure. Therefore, gust mitigation analysis for highly flexible aircraft is crucial for ensuring flight performance and safety.

[0003] Previous aircraft gust mitigation solutions were mostly linear methods based on the assumption of small deformations, making them unsuitable for mitigating gust loads on large, flexible aircraft with geometrically nonlinear, large deformation characteristics. Chinese invention patent application publication number CN107765698A discloses a control method for mitigating vertical gust loads on large aircraft, but it fails to consider the nonlinear effects caused by large deformations and is solely a linear gust mitigation approach. Summary of the Invention

[0004] In view of the above problems, the present invention provides a gust load mitigation method suitable for large flexible aircraft, which solves the technical problem in the prior art that it is difficult to consider the changes in aerodynamic shape and structural stiffness caused by the large deformation of the wings of flexible aircraft with a large aspect ratio.

[0005] The present invention provides a method for alleviating gust loads applicable to a large flexible aircraft, comprising the following steps:

[0006] Step S1, setting a flight parameter sensor on the large flexible aircraft; setting an optical fiber strain sensor on the wing of the large flexible aircraft; measuring the local structural strain using the optical fiber strain sensor to obtain the local curvature of the wing;

[0007] Step S2: establishing an inner control loop, wherein the inner control loop performs feedback adjustment based on the local curvature of the wing to determine an inner control surface control adjustment amount; and determining parameters of the feedback adjustment by optimizing a minimized performance index;

[0008] Step S3: establishing an outer control loop, wherein the outer control loop performs tracking compensation based on the measurement value of the flight parameter sensor to determine the outer control surface control adjustment amount;

[0009] Step S4: combining the inner control loop and the outer control loop to establish a gust mitigation control model for the large flexible aircraft, and determining a final gust mitigation control surface adjustment amount;

[0010] Step S5: Control the control surfaces of the highly flexible aircraft based on the gust mitigation control surface adjustment amount to achieve gust load mitigation.

[0011] Preferably, in step S1: the flight parameter sensors include angle, height and speed sensors; the optical fiber strain sensors are set on the upper surface or lower surface of the wing and arranged in a grid or linear manner, and the number of the optical fiber strain sensors is determined by the size of the detection area.

[0012] Preferably, in step S1, the calculation expression for obtaining the local curvature of the wing by measuring the local structural strain using the optical fiber strain sensor is:

[0013] E 11 =-x2(K3-k3)+x3(K2-k2)

[0014]

[0015] E 22 =E 33 =E 23 =E 32 =0

[0016] Among them, E ij is the Green strain tensor, i and j are the indices of the coordinate axes, and the value range of i and j is {1, 2, 3}. x2 and x3 are the coordinates of the y-axis and z-axis of each strain sensor in the local coordinate system of the beam section, respectively. K1, K2, K3 are the curvatures at the sensor measuring points, and k1, k2, k3 are the initial curvatures at the sensor measuring points. If the initial state of the beam remains straight, then k1 = k2 = k3 = 0.

[0017] Preferably, step S2 specifically includes:

[0018] Step S2-1, establishing a linear model that describes the relationship between the state vector, the inner loop control signal, and the measurement output, wherein the measurement output is the local curvature of the wing;

[0019] Step S2-2: Optimizing the linear model using a linear quadratic regulator method to obtain feedback gain parameters;

[0020] Step S2-3: Based on the feedback gain parameter, determine the inner loop control signal as the inner loop control adjustment amount.

[0021] Preferably, step S2-1 specifically includes:

[0022] Establish the following linear model:

[0023]

[0024] y=Cx

[0025] in, is the derivative of the state vector, x is the state vector, u i is the inner loop control signal, y is the measured output, A is the state matrix, B is the input matrix, and C is the output matrix;

[0026] The inner loop control signal expression is:

[0027] u i =-Ky

[0028] Where K is the feedback gain matrix, and the final inner loop closed-loop system is expressed as:

[0029]

[0030] Among them, A c =A-BKC represents the closed-loop state matrix.

[0031] Preferably, step S2-2 specifically includes:

[0032] The value of P in the LQR problem is obtained by iteratively solving the following Lyapunov equation:

[0033]

[0034] Where R is the positive definite control weight matrix, Q is the positive semi-steady weight matrix, X = x(0) T x(0), x(0) is the initial value of the state vector x, P is the positive definite matrix representing the state feedback weight, S is the Lagrange multiplier; the feedback gain matrix is ​​K = R - 1 B T PSC T (CSC T ) -1 ;

[0035] The performance index J is calculated from the value of P:

[0036]

[0037] Where tr(·) represents the trace of the matrix;

[0038] Finally, the feedback gain matrix K is calculated by minimizing the performance index J.

[0039] Preferably, step S3 specifically includes:

[0040] Step S3-1, establishing a compensator model that describes the relationship between the tracking error and the outer loop control signal;

[0041] Step S3-2: using the compensator model, determining an outer loop control signal from the tracking error; the tracking error is determined by a measurement value of the flight parameter sensor.

[0042] Preferably, step S3-1 specifically includes:

[0043] The reference input expression of the outer control loop is:

[0044] z T =H a x

[0045] where z T is a vector containing the variables to be tracked, H a Represents the selector of the feedback variable; the compensator expression of the outer control loop is:

[0046]

[0047] u v =Dw c +J c e

[0048] Among them, w c is the state of the compensator, F, G, D and J c is the compensator structure matrix, e is the tracking error, u v is the control signal output by the outer control loop, and the tracking error e is calculated as:

[0049] e=rz T

[0050] Where r is a vector containing the references to be tracked.

[0051] Preferably, the control signals of the outer control loop and the inner control loop are superimposed to obtain the final gust mitigation control signal u, and the final gust mitigation control model is expressed as an augmented form expression:

[0052]

[0053] The expression of gust mitigation control signal u is:

[0054]

[0055] Where I is the unit matrix.

[0056] Compared with the prior art, the present invention has at least the following beneficial effects:

[0057] (1) The present invention fully considers the large deformation of the wing of a highly flexible aircraft, controls the wing within a small deformation range through deformation identification and control, and then performs linear gust mitigation. Compared with conventional gust mitigation methods, the present invention has a wider range of applications and is of great significance for mitigating gust loads on highly flexible aircraft with large aspect ratios.

[0058] (2) The method provided by the present invention can not only be used for gust load mitigation analysis of large flexible aircraft in the design stage, but can also play an important role in experimental testing and control law design. BRIEF DESCRIPTION OF THE DRAWINGS

[0059] The drawings are only for purposes of illustrating particular embodiments and are not to be considered limiting of the invention.

[0060] Figure 1 The present invention provides a flow chart of the method for alleviating gust loads on a large flexible aircraft.

[0061] Figure 2 This is a schematic diagram of the highly flexible aircraft structure model and coordinate system provided by the present invention.

[0062] Figure 3 Schematic diagram of the geometric configuration of the nonlinear beam model provided by the present invention.

[0063] Figure 4 Schematic diagram of the control flow for alleviating gust loads on a large flexible aircraft provided by the present invention. DETAILED DESCRIPTION

[0064] In order to more clearly understand the above-mentioned objects, features and advantages of the present invention, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments can be combined with each other. In addition, the present invention can also be implemented in other ways different from those described herein. Therefore, the scope of protection of the present invention is not limited by the specific embodiments disclosed below.

[0065] The gust load mitigation method for a large flexible aircraft provided by the present invention includes the processes of wing deformation identification, wing deformation control and linear gust mitigation of the large flexible aircraft; deformation identification of the large flexible wing is performed based on an optical fiber strain sensor; inner loop control is used to control the wing within a small deformation range; and a linear method is used to mitigate gusts on the aircraft controlled within the small deformation range.

[0066] like Figure 2 、 Figure 3As shown in the figure, the wing of a high-aspect-ratio aircraft is considered as a nonlinear beam, and a geometric model is established. Three coordinate systems are defined in the geometric model: the earth coordinate system E, the body axis system B, and the local coordinate system G. The earth coordinate system is an inertial system, ignoring the effects of the earth's curvature and rotation; the origin of the body axis system B is located on the fuselage and moves with the aircraft; the origin of the local coordinate system G is located on the wing reference line and moves with the elastic deformation of the wing.

[0067] In order to illustrate the effectiveness of the method proposed by the present invention, the above technical solution of the present invention is described in detail below through a specific embodiment. Figure 1 As shown, a method for alleviating gust loads applicable to large flexible aircraft is disclosed, and the specific implementation steps are as follows:

[0068] Step S1: a flight parameter sensor is set on the large flexible aircraft; an optical fiber strain sensor is set on the wing of the large flexible aircraft; and the local curvature of the wing is obtained by measuring the local structural strain using the optical fiber strain sensor.

[0069] An optical fiber strain sensor is installed on the wing of a large flexible aircraft, and the strain information obtained by the optical fiber strain sensor is used to obtain wing deformation information; the strain information includes local structural strain, and the wing deformation information includes the local curvature vector of the wing.

[0070] Firstly, the relationship between the deformation of the wing of a large flexible aircraft and the response value of the optical fiber strain sensor is determined by analysis.

[0071] Define the coordinate transformation matrix from the local coordinate system to the body axis system at time t and arc length coordinate s as C BG , its inverse matrix C GB =[C BG ] -1 , and then the curvature vector can be determined

[0072]

[0073] Among them C BG′ C BG Derivatives with respect to arc length, curvature vector Describes the rate of change of the local coordinate system relative to the body axis system as the arc length s changes.

[0074] Assume that the coordinates of the origin of the local coordinate system G in the body axis system are R B , then the strain expression is as follows:

[0075] γ=C GB R′ B -E1

[0076]

[0077] Where γ=[γ 11 2γ 12 2γ 13 ] T is the force strain, κ=[κ1 κ2 κ3] T is the moment strain, R′ B is the local coordinate system displacement vector R B The derivative of E1 = [1 0 0] T , is the coordinate transformation matrix at time 0, is the derivative of the coordinate transformation matrix with respect to the arc length at time 0. In the above way, the force strain describes the difference between the displacement in the current state and the displacement in the initial state, and the moment strain describes the difference between the rotation in the current state and the initial state.

[0078] The complete strain vector is denoted as ε, and its expression is:

[0079] ε=[γ T κ T ] T

[0080] In the above way, the force strain and moment strain are combined into a strain vector, which is used to describe the overall deformation state of the object under the action of force.

[0081] The wing of a high aspect ratio aircraft is regarded as a nonlinear beam, and the nonlinear beam is divided into N equal strain units with node positions s0, s1, ..., s n ,…,s N , assuming that the internal strain value of each unit remains unchanged along the length direction, then the nth unit s n-1 ≤s n The internal strain is Assume that both shear strain and tensile strain can be neglected, that is, γ n =0, then we have the following relational expression:

[0082]

[0083] in, Represents the curvature vector The modulus of ||·|| represents the modulus of the calculated vector, is the coordinate transformation matrix of the n-1th unit, I is the unit matrix, R B,n-1 is the displacement vector of the n-1th element. The above formula defines the analytical relationship between the beam configuration and curvature under the constant strain element assumption. Once the curvature vector within the element is determined, the spatial position of any point on the beam reference line can be calculated based on the obtained coordinate transformation matrix and the displacement vector. Using the above formula, the spatial position of any point on the beam reference line can be calculated using the curvature vector.​

[0084] Through the above analysis, combined with Green's strain theory, the relationship between the structural strain measured by the optical fiber strain sensor installed on the wing and the curvature vector can be obtained as follows:

[0085] E 11 =-x2(K3-k3)+x3(K2-k2)

[0086]

[0087] E 22 =E 33 =E 23 =E 32 =0

[0088] Among them, E ij is the Green strain tensor, i and j are coordinate axis indices, and both range from {1, 2, 3}. x2 and x3 are the y- and z-axis coordinates of each strain sensor in the local coordinate system of the beam cross section, respectively. K1, K2, and K3 are the curvatures at the sensor measurement points, and k1, k2, and k3 are the initial curvatures at the sensor measurement points. If the beam is initially straight, then k1 = k2 = k3 = 0. Using the above formula, the curvature of the wing in different directions can be calculated from the strain data measured by the fiber optic strain sensor.

[0089] In some embodiments, the optical fiber strain sensor can be set on the upper surface or lower surface of the wing, and can be installed by pasting, embedding or other fixing methods, arranged in a grid or linear manner, and the number of sensors can be configured according to the size of the detection area.

[0090] Step S2: establishing an inner control loop, wherein the inner control loop performs feedback adjustment based on the local curvature of the wing to determine the inner control surface control adjustment amount; and determining the parameters of the feedback adjustment by optimizing the minimized performance index.

[0091] In this step, the deformation information of the wing curvature change, obtained by strain detection using the optical fiber strain sensor, is used to adjust the control input to control the deformation within a small deformation range. Specifically, this is achieved by establishing a linearized state-space model and using feedback control to reduce deformation.

[0092] In this step, the wing curvature change measured in the previous step is used as the measurement output y for linear modeling. The following linear model is established:

[0093]

[0094] y=Cx

[0095] in, is the derivative of the state vector, x is the state vector, u i is the inner loop control signal, y is the measured output, A is the state matrix, B is the input matrix, and C is the output matrix.

[0096] The inputs to the inner control loop include a state vector, historical control signals, and measurement outputs. The state vector describes the current state of the system, which can be displacement, velocity, or other examples. The historical control signal is the control signal from the control system prior to the current moment, used to control the deflection angle of the control surface. The measurement output is the system output measured by the sensor, namely, the change in wing curvature. The output of the inner control loop is the current control signal, a control input calculated through feedback control, used to adjust actuators such as the control surface.

[0097] Through output feedback control, the control signal is obtained by the following expression:

[0098] u i =-Ky

[0099] Where K is the feedback gain matrix, and the final inner loop closed-loop system is expressed as:

[0100]

[0101] Among them, A c =A-BKC represents the closed-loop state matrix.

[0102] The present invention adopts the linear quadratic regulator (LQR) method with output feedback to calculate the feedback gain, and calculates the feedback gain matrix K by minimizing the performance index J. The performance index is defined as:

[0103]

[0104] Where R is the positive definite control weight matrix, Q is the positive semi-steady weight matrix, and t is time. This expression is used to regulate the performance of the measured output.

[0105] Then, the solution to the LQR problem is obtained by solving the following Lyapunov equation:

[0106]

[0107] Where X = x(0) T x(0), x(0) is the initial value of the state vector x, P is the positive definite matrix representing the state feedback weight, and S is the Lagrange multiplier. The feedback gain matrix is ​​K = R -1 B T PSC T (CSC T ) -1

[0108] These equations are solved iteratively to obtain a suboptimal solution that moves towards minimizing the quadratic performance indicator, and the solution depends on the initial return. The final J can be written as:

[0109]

[0110] where tr(·) represents the trace of the matrix.

[0111] The present invention calculates the feedback gain matrix through the above-mentioned LQR method to obtain a control signal, and forms an inner control loop for adjusting the control signal in real time to control the deformation within a small deformation range.

[0112] Step S3: establishing an outer control loop, wherein the outer control loop performs tracking compensation based on the measurement value of the flight parameter sensor to determine the outer control surface control adjustment amount.

[0113] The outer control loop handles slower dynamic changes and compensates for tracking variables such as angle, altitude, or speed based on measurements from the aircraft's flight parameter sensors, which can measure the aircraft's angle, altitude, or speed.

[0114] The reference input of the outer control loop is given by:

[0115] z T =H a x

[0116] where z T is a vector containing the variable to be tracked, such as angle, height, or speed, etc. a Represents the selector of the feedback variable. The compensator expression of the outer control loop is:

[0117]

[0118] u v =Dw c +J c e

[0119] Among them, w c is the state of the compensator, F, G, D and J c is the compensator structure matrix, e is the tracking error, u v is the control signal output by the outer control loop. The tracking error e is calculated as:

[0120] e=rz T

[0121] Where r is a vector containing the references to be tracked.

[0122] For the outer control loop, the tracking error e is obtained by the calculated error between the measured value of the flight parameter sensor on the aircraft and the reference value. The control signal u output by the outer control loop can be obtained from the tracking error e and the state of the compensator. v , used to adjust actuators such as rudders.

[0123] Step S4, combining the inner control loop and the outer control loop, establishing a gust mitigation control model for a large flexible aircraft, and determining the final gust mitigation control adjustment amount of the control surface. The final control flow is as follows Figure 4 shown.

[0124] The control signals of the outer control loop and the inner control loop are superimposed to obtain the final control signal u:

[0125] u=-Ky-u v

[0126] The final gust mitigation control model is written in augmented form as:

[0127]

[0128] Gust mitigation control signal:

[0129]

[0130] Step S5: Control the control surfaces of the highly flexible aircraft based on the gust mitigation control surface adjustment amount to achieve gust load mitigation.

[0131] Through the above scheme, the present invention fully considers the large deformation of the wings of highly flexible aircraft, controls the wings within a small deformation range through deformation identification and control, and then performs linear gust mitigation. Compared with conventional gust mitigation methods, the present invention has a wider range of applications and is of great significance for mitigating gust loads on highly flexible aircraft with large aspect ratios.

[0132] Although the specific embodiments of the present invention have been described in a particular order, it should be understood that such actions or steps are required to be performed in the particular order shown or in a sequential order, or that all illustrated actions or steps are required to be performed to obtain the desired result. Under certain circumstances, multitasking and parallel processing may be advantageous. Similarly, although some specific implementation details have been included in the above discussion, these should not be construed as limiting the scope of the present disclosure. Some features described in the context of a separate embodiment can also be implemented in a single implementation in combination. On the contrary, the various features described in the context of a single implementation can also be implemented in multiple implementations individually or in any suitable sub-combination.

[0133] The above description is only a preferred specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by any technician familiar with this technical field within the technical scope disclosed by the present invention should be covered by the scope of protection of the present invention.

Claims

1. A method for alleviating gust loads for a large flexible aircraft, characterized in that: The following steps are involved: Step S1, setting a flight parameter sensor on the large flexible aircraft; setting an optical fiber strain sensor on the wing of the large flexible aircraft; measuring the local structural strain using the optical fiber strain sensor to obtain the local curvature of the wing; Step S2: establishing an inner control loop, wherein the inner control loop performs feedback adjustment based on the local curvature of the wing to determine an inner control surface control adjustment amount; and determining parameters of the feedback adjustment by optimizing a minimized performance index; Step S3: establishing an outer control loop, wherein the outer control loop performs tracking compensation based on the measurement value of the flight parameter sensor to determine the outer control surface control adjustment amount; Step S4: combining the inner control loop and the outer control loop to establish a gust mitigation control model for the large flexible aircraft, and determining a final gust mitigation control surface adjustment amount; Step S5: controlling the control surface of the highly flexible aircraft based on the gust mitigation control surface adjustment amount to achieve gust load mitigation; Step S4 specifically includes: The control signals of the outer control loop and the inner control loop are superimposed to obtain the final gust mitigation control signal , and the final gust mitigation control model is expressed as an augmented form expression: Gust mitigation control signal The expression is: in, is the unit matrix, is the state vector, is the state of the compensator, is the state matrix, 、 、 and is the compensator structure matrix, is the input matrix, is the inner loop control signal, is a vector containing the references to be tracked, To measure the output, is the control signal output by the outer control loop, is the output matrix, A selector representing a feedback variable, is a vector containing the variables to be tracked, is the feedback gain matrix.

2. The gust load mitigation method for a large flexible aircraft according to claim 1, characterized in that: In the step S1: The flight parameter sensors include angle, height and speed sensors; the optical fiber strain sensors are set on the upper surface or lower surface of the wing and arranged in a grid or linear manner. The number of the optical fiber strain sensors is determined by the size of the detection area.

3. The gust load mitigation method for a large flexible aircraft according to claim 2, characterized in that: In step S1, the calculation expression for obtaining the local curvature of the wing by measuring the local structural strain using the optical fiber strain sensor is: in, is the Green strain tensor, i and j are the indices of the coordinate axes, and the value range of i and j is {1, 2, 3}. 、 are the y-axis and z-axis coordinates of each strain sensor in the local coordinate system of the beam section, is the curvature at the sensor point, is the initial curvature at the sensor measuring point. If the initial state of the beam remains straight, then .

4. The gust load mitigation method for a large flexible aircraft according to claim 3, characterized in that: Step S2 specifically includes: Step S2-1, establishing a linear model that describes the relationship between the state vector, the inner loop control signal, and the measurement output, wherein the measurement output is the local curvature of the wing; Step S2-2: Optimizing the linear model using a linear quadratic regulator method to obtain feedback gain parameters; Step S2-3: Based on the feedback gain parameter, determine the inner loop control signal as the inner loop control adjustment amount.

5. The gust load mitigation method for a large flexible aircraft according to claim 4, characterized in that: Step S2-1 specifically includes: Establish the following linear model: in, is the derivative of the state vector, is the state vector, is the inner loop control signal, To measure the output, is the state matrix, is the input matrix, is the output matrix; The inner loop control signal expression is: in is the feedback gain matrix, and the final inner loop closed-loop system is expressed as: in, represents the closed-loop state matrix.

6. The gust load mitigation method for a large flexible aircraft according to claim 5, characterized in that: Step S2-2 specifically includes: By iteratively solving the following Lyapunov equation, we can obtain Value: in, is the positive definite control weight matrix, is the positive semi-stationary weight matrix, , is the state vector The initial value of is the positive definite matrix representing the state feedback weight, is the Lagrange multiplier; the feedback gain matrix is ; Depend on The performance index is calculated by : in, represents the trace of the matrix; Finally, by minimizing the performance index To get the feedback gain matrix .

7. The gust load mitigation method for a large flexible aircraft according to any one of claims 1 to 6, characterized in that: Step S3 specifically includes: Step S3-1, establishing a compensator model that describes the relationship between the tracking error and the outer loop control signal; Step S3-2: using the compensator model, determining an outer loop control signal from the tracking error; the tracking error is determined by a measurement value of the flight parameter sensor.

8. The gust load mitigation method for a large flexible aircraft according to claim 7, characterized in that: Step S3-1 specifically includes: The reference input expression of the outer control loop is: in is a vector containing the variables to be tracked, Represents the selector of the feedback variable; the compensator expression of the outer control loop is: in, is the state of the compensator, 、 、 and is the compensator structure matrix, is the tracking error, is the control signal output by the outer control loop, tracking error The calculation method is: in is a vector containing the references to be tracked.

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