Integrated guidance and control method for high dynamic aircraft

Abstract

The application discloses an integrated guidance control method applied to a high-dynamic aircraft. In the method, a control signal of the aircraft is obtained in real time, and the control signal is transmitted to a rudder on the aircraft in real time. The control signal comprises equivalent rudder deflection angles for controlling pitching, yawing and rolling of the aircraft. In the process of obtaining the control signal, a tracking differentiator is introduced to obtain a required derivative of a virtual control quantity while ensuring accuracy, and the control signal is obtained by setting an error surface.

Description

TECHNICAL FIELD

[0001] The present application relates to the guidance control of high dynamic aircraft, in particular to an integrated guidance control method for high dynamic aircraft. BACKGROUND

[0002] In the process of tracking a maneuvering target, high dynamic aircraft often flies in a large overload turning mode in the terminal guidance phase. The traditional separated design method separates the design of guidance and control, and it is generally believed that the frequency of the guidance loop signal is slower than that of the control loop signal. However, in the process of large overload turning, if the frequencies of the guidance and control loops tend to be uniform, integrated design will help to improve the guidance effect and also simplify the design.

[0003] The traditional backstepping control method solves the derivative of the virtual control quantity required in the next step design by direct differentiation, but when the virtual control quantity changes rapidly, direct differentiation will make the derivative variable obtained rapidly expand and change, affecting the control process.

[0004] Due to the above reasons, the present inventors have made in-depth research on the integrated guidance control method in the hope of designing an integrated guidance control method for high dynamic aircraft that can solve the above problems. SUMMARY

[0005] In order to overcome the above problems, the present inventors have made intensive research and designed an integrated guidance control method for high dynamic aircraft. In this method, the control signal of the aircraft is obtained in real time, and the control signal is transmitted to the rudder on the aircraft in real time, the control signal including the equivalent rudder deflection angle for controlling the pitch, yaw and roll of the aircraft; in the process of obtaining the control signal, the required derivative of the virtual control quantity is obtained while ensuring the accuracy by introducing a tracking differentiator, and the control signal is fitted by setting an error surface, thereby completing the present application.

[0006] Specifically, the purpose of the present application is to provide an integrated guidance control method for high dynamic aircraft,

[0007] In this guidance method, the control signal of the aircraft is obtained in real time, and the control signal is transmitted to the rudder on the aircraft in real time,

[0008] The control signal includes the equivalent rudder deflection angle for controlling the pitch, yaw and roll of the aircraft;

[0009] The rudder works according to the equivalent rudder deflection angle, adjusts the flight state of the aircraft, so that the aircraft hits the target.

[0010] The control signal is obtained by the following formula (I):

[0011]

[0012] wherein u represents a control signal,

[0013] K3 represents a design parameter,

[0014] s3 represents an error surface three,

[0015] g3 and f3 both represent intermediate variables,

[0016] denotes a derivative of x 3d obtained by a tracking differentiator.

[0017] wherein the error surface three s3 is obtained by the following equation (II):

[0018] s3 = x3 - x 3d (II)

[0019] wherein x3 represents an angular velocity of the aircraft, x3 = [ω z ω y ω x ] T ; wherein ω z represents a pitch angular velocity of the aircraft, ω y represents a yaw angular velocity of the aircraft; and ω x represents a roll angular velocity of the aircraft;

[0020] x 3d represents a real-time tracking virtual control quantity of three variables in x3.

[0021] wherein the x 3d is obtained by the following equation (III):

[0022]

[0023] wherein K2 represents a design parameter,

[0024] s2 represents an error surface two,

[0025] g2 and f2 both represent intermediate variables,

[0026] denotes a derivative of x 2d obtained by a tracking differentiator.

[0027] wherein the error surface two s2 is obtained by the following equation (IV):

[0028] s2 = x2 - x 2d (IV)

[0029] wherein x2 represents an angle of the aircraft; x2 = [α β γ] TWherein, α represents the angle of attack of the aircraft, β represents the sideslip angle of the aircraft, and γ represents the roll angle of the aircraft.

[0030] x 2d The real-time tracking virtual control quantity of the three variables in x2 is represented.

[0031] Wherein, the x 2d Obtained by the following formula (five):

[0032]

[0033] Wherein, K1 represents a design parameter,

[0034] s1 represents the error surface one,

[0035] f1 and g1 are both intermediate variables.

[0036] Wherein, the error surface one s1 is obtained by the following formula (six):

[0037] s1=x1 (six)

[0038] Wherein, x1 represents the line-of-sight angular velocity of the aircraft relative to the target, ε represents the line-of-sight inclination angle of the target relative to the aircraft, and η represents the line-of-sight deflection angle of the target relative to the aircraft; is the derivative of ε, representing the line-of-sight angular velocity in the pitch direction of the target relative to the aircraft; is the derivative of η, representing the line-of-sight angular velocity in the yaw direction of the target relative to the aircraft.

[0039] The beneficial effects of the present application include:

[0040] According to the integrated guidance control method for a high-dynamic aircraft provided by the present application, the rudder signal is designed by backstepping based on a three-dimensional integrated guidance control model according to the guidance control requirements of the aircraft, so as to obtain the control signal in real time, which can optimize the differential expansion problem generated in the traditional backstepping control process. Meanwhile, the virtual control quantity generated in the backstepping design process is differentiated by using a tracking differentiator, so that the obtained result has higher precision, thereby generating higher control precision and realizing better guidance control effect. BRIEF DESCRIPTION OF DRAWINGS

[0041] Figure 1 The motion trajectory of the aircraft and the target in the embodiment of the present application is shown;

[0042] Figure 2 The equivalent rudder deflection angle change curve obtained in real time in the embodiment of the present application is shown;

[0043] Figure 3Figures showing the virtual control variable x in the experimental examples and comparative examples of the present application 2dε Figures showing the virtual control variable x in the experimental examples and comparative examples of the present application

[0044] Figure 4 Figures showing the virtual control variable x in the experimental examples and comparative examples of the present application 2dη Figures showing the virtual control variable x in the experimental examples and comparative examples of the present application

[0045] Figure 5 Figures showing the virtual control variable x in the experimental examples and comparative examples of the present application 2d Figures showing the virtual control variable x in the experimental examples and comparative examples of the present application DETAILED DESCRIPTION

[0046] The application will be further described in the following drawings and examples. The features and advantages of the present application will become more apparent from these descriptions.

[0047] The term "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Unless specifically stated otherwise, the drawings are not drawn to scale and are merely intended to conceptually illustrate the features described herein.

[0048] According to the present application, an integrated guidance control method for a high-dynamic aircraft is provided. In the guidance method, a control signal of the aircraft is obtained in real time, and the control signal is transmitted to a rudder on the aircraft in real time. The control signal includes an equivalent rudder deflection angle for controlling the pitch, yaw and roll of the aircraft. The rudder operates according to the equivalent rudder deflection angle to adjust the flight state of the aircraft, so that the aircraft hits the target.

[0049] Preferably, u represents the control signal, u = [δ z δ y δ x ] T ; δ z represents the equivalent rudder deflection angle for controlling the pitch of the aircraft, δ y represents the equivalent rudder deflection angle for controlling the yaw of the aircraft, and δ x represents the equivalent rudder deflection angle for controlling the roll of the aircraft. In the present application, the guidance system and the control system are integrated and processed directly to obtain the equivalent rudder deflection angle converted by the control system in the traditional scheme, and the rudder is controlled directly according to the equivalent rudder deflection angle. This can eliminate the defect of the frequency inconsistency between the guidance loop and the control loop, and is helpful to improve the guidance effect.

[0050] In a preferred embodiment, the control signal is obtained by the following formula (I):

[0051]

[0052] wherein u represents the control signal,

[0053] K3 represents the design parameter, and its preferred value is 50;

[0054] s3 indicates error surface three.

[0055] Both g3 and f3 represent intermediate variables;

[0056] x represents the value obtained by the tracking differentiator. 3d The derivative of .

[0057] The tracking differentiator can be represented as:

[0058]

[0059] Where R3 takes the value 30; the function F(·) can be expressed as:

[0060] F(h1,h2)=υ(h1)+υ(h2)

[0061]

[0062] constant a i >0, (i=2,3); In this application, the values ​​of a2 and a3 are preferably 400.

[0063] By tracking the differentiator, we can obtain... Approximate value z 32 In this application, To represent z 32 .

[0064] In a preferred embodiment, the error surface s3 is obtained by the following equation (ii):

[0065] s3 = x3 - x 3d (two)

[0066] Where x3 represents the angular velocity of the aircraft; preferably, x3 = [ω z ω y ω x ] T ;ω z ω represents the pitch angular velocity of the aircraft. y ω represents the yaw rate of the aircraft. x This indicates the roll angular velocity of the aircraft;

[0067] The error surface set in the present application can be derived in the process of "backstepping control", and then the control signal u is finally obtained, which is a necessary step to obtain the control signal. In the present application, by making the error surface as small as possible, the tracking amount of the virtual control quantity can track the tracked quantity; such setting can ensure that the finally obtained control signal u can make each error surface close to 0, that is, the designed virtual control quantity can well track the target variable, and the expected guidance and control effect is achieved.

[0068] x 3d The real-time tracking virtual control quantity of the three variables in x3.

[0069] In the present application, the angular velocity of the aircraft is obtained in real time by sensors carried in the aircraft, such as gyroscopes and inertial navigation systems.

[0070] In a preferred embodiment, the x 3d is obtained by the following formula (three):

[0071]

[0072] wherein K2 represents a design parameter, and preferably its value is 10; by designing each design parameter, the stability of the overall system can be ensured, and the control signal U obtained based on each design parameter is input into the aircraft, and the aircraft flies accordingly, which will not cause unstable response of the aircraft.

[0073] s2 represents error surface two,

[0074] g2 and f2 are intermediate variables,

[0075] represents the derivative of x 2d obtained by the tracking differentiator.

[0076] The tracking differentiator can be represented as:

[0077]

[0078] wherein R2 takes a value of 30; the function F(·) can be represented as:

[0079] F(h1, h2) = υ(h1) + υ(h2)

[0080]

[0081] The constant a i > 0, (i = 2, 3); the values of a2 and a3 are preferably 400.

[0082] The approximate value z of can be obtained by the tracking differentiator22 , in the present application to represent z 22 .

[0083] In a preferred embodiment, the error surface two s2 is obtained by the following formula (four):

[0084] s2 = x2 - x 2d (four)

[0085] wherein x2 represents the angles of the aircraft; x2 = [α β γ] T ; α represents the angle of attack of the aircraft, β represents the sideslip angle of the aircraft, and γ represents the roll angle of the aircraft; preferably, the angles of the aircraft are obtained in real time by sensors carried on the aircraft, such as gyroscopes and inertial navigation systems, etc.

[0086] x 2d represents the virtual control quantity that the variables in x2 need to track.

[0087] In a preferred embodiment, the x 2d is obtained by the following formula (five):

[0088]

[0089] By adding the term 0 in formula (five), the order of the system is consistent.

[0090] wherein K1 represents a design parameter, and its value is preferably 10;

[0091] s1 represents the error surface one,

[0092] f1 and g1 both represent intermediate variables.

[0093] In a preferred embodiment, the error surface one s1 is obtained by the following formula (six):

[0094] s1 = x1 - 0 (six)

[0095] The purpose of formula six in the present application is to control the relative velocity perpendicular to the line connecting the aircraft and the target to converge to 0, so that the aircraft completely flies towards the target.

[0096] wherein x1 represents the line-of-sight angular velocity of the aircraft relative to the target, ε represents the line-of-sight inclination angle of the target relative to the aircraft as the origin; η represents the line-of-sight deflection angle of the target relative to the aircraft as the origin; is the derivative of ε, representing the line-of-sight angular velocity in the pitch direction of the target relative to the aircraft as the origin; is the derivative of η, representing the yaw direction line-of-sight angular velocity of the target relative to the aircraft. In the present application, the line-of-sight angular velocity of the target relative to the aircraft is obtained in real time by a seeker on the aircraft.

[0097] Preferably, the line-of-sight coordinate system formed by the aircraft and the target is defined as follows: taking the center of gravity of the aircraft as the coordinate origin, taking the line connecting the center of gravity of the aircraft and the center of gravity of the target as the X axis, taking the direction from the aircraft to the target as positive, and taking the yaw direction of the aircraft as the Y axis and the pitch direction of the aircraft as the Z axis during the flight of the aircraft along the X axis.

[0098] In a preferred embodiment, the intermediate variable is obtained by the following formula:

[0099]

[0100]

[0101]

[0102]

[0103]

[0104]

[0105] The above intermediate variable is a dynamic model based on the flight mechanics of the aircraft and set for the missile form; wherein: R represents the relative distance between the aircraft and the target, which is obtained in real time by a seeker on the aircraft, represents the derivative of R; (a Tr ,a Tε ,a Tη represents the acceleration of the target in the line-of-sight coordinate system formed by the aircraft and the target; q represents the dynamic pressure of the aircraft, q = 0.5pV 2 ; p represents the air density, and V represents the speed of the aircraft in the earth coordinate system;

[0106] The following aircraft-related parameters are known, specifically, P represents the thrust of the aircraft, m represents the mass of the aircraft, S represents the reference area of the aircraft, and L represents the characteristic length of the aircraft; represents the lift coefficient; represents the side force coefficient; represents the lateral static stability derivative induced by the angle of attack; represents the lateral static stability derivative induced by the side slip angle; represents the yaw static stability derivative, represents the pitch static stability derivative, represents the aileron control efficiency, represents the rudder control efficiency, denotes the elevator control efficiency, J x denotes the aircraft moment of inertia about the longitudinal axis, J y denotes the aircraft moment of inertia about the vertical axis, J z denotes the aircraft moment of inertia about the lateral axis.

[0107] Experimental example

[0108] The initial position of the aircraft is (0 0 0), the target initial position is (5km 5km 5km), the initial speed of the aircraft is 600m / s, and the target has an irregular acceleration form,

[0109] An integrated guidance control method for high-dynamic aircraft is used to guide and control the aircraft, and a control signal is directly obtained, the control signal including an equivalent rudder deflection angle for controlling the pitch, yaw and roll of the aircraft; wherein the control signal is obtained in real time by the following formula (I):

[0110]

[0111] denotes the derivative of x 3d obtained by a tracking differentiator; the tracking differentiator is denoted as:

[0112]

[0113] wherein R3 is 30; the function F(·) is denoted as:

[0114] F(h1, h2) = υ(h1) + υ(h2)

[0115]

[0116] a constant a i > 0, (i = 2, 3), the values of a2 and a3 are both 400,

[0117] The approximate value z 32 of x obtained by a tracking differentiator is:

[0118] wherein the error surface s3 is obtained by the following formula (II):

[0119] s3 = x3 - x 3d (II)

[0120] x3 = [ω z ω y ω x ] T

[0121] x3d is obtained by the following equation (three):

[0122]

[0123] Error surface two s2 is obtained by the following equation (four):

[0124] s2 = x2 - x 2d (four)

[0125] x2 = [a b g] T

[0126] x 2d is obtained by the following equation (five):

[0127]

[0128] Error surface one s1 is obtained by the following equation (six):

[0129] s1 = x1 - 0 (six)

[0130]

[0131] The design parameter K1 takes a value of 10, K2 takes a value of 10, and K3 takes a value of 50;

[0132] The intermediate variable is obtained by the following equation:

[0133]

[0134]

[0135]

[0136]

[0137]

[0138]

[0139] ε represents the line-of-sight inclination angle of the target relative to the aircraft as the origin; η represents the line-of-sight declination angle of the target relative to the aircraft as the origin; is the derivative of ε, representing the line-of-sight angular velocity in the pitch direction of the target relative to the aircraft as the origin; is the derivative of η, representing the line-of-sight angular velocity in the yaw direction of the target relative to the aircraft as the origin;

[0140] α represents the angle of attack of the aircraft, β represents the sideslip angle of the aircraft, and γ represents the roll angle of the aircraft;

[0141] ω z represents the pitch angular velocity of the aircraft, and ωy ω represents the yaw rate of the aircraft. x This indicates the roll angular velocity of the aircraft;

[0142] R represents the relative distance between the aircraft and the target. Represents the derivative of R; (a Tr ,a Tε ,a Tη ) represents the target acceleration in the line-of-sight coordinate system formed by the aircraft and the target; q represents the dynamic pressure of the aircraft, q = 0.5ρV 2 ρ represents air density, and V represents the speed of the aircraft.

[0143] P represents the thrust of the aircraft, m represents the mass of the aircraft, S represents the reference area of ​​the aircraft, and L represents the characteristic length of the aircraft. Indicates the lift coefficient; Indicates the lateral force coefficient; This represents the lateral static stability derivative caused by the angle of attack; This represents the lateral static stability derivative induced by the sideslip angle; This represents the static stability derivative of yaw. This represents the statically stable derivative of pitch. Indicates aileron control efficiency. Indicates rudder control efficiency. J represents the elevator control efficiency. x J represents the moment of inertia of an aircraft about its longitudinal axis. y J represents the moment of inertia of an aircraft about its vertical axis. z This represents the moment of inertia of an aircraft rotating about its horizontal axis.

[0144] Equation (1) provides the real-time control signal u, u = [δ z δ y δ x ] T ;δ z δ represents the equivalent control deflection angle for controlling the pitch of an aircraft. y δ represents the equivalent control deflection angle for controlling the yaw of an aircraft. x This represents the equivalent rudder deflection angle used to control the aircraft's roll; based on this equivalent rudder deflection angle, the aircraft is controlled to track the target, and the final guidance terminal accuracy is 0.159m.

[0145] Figure 1 The image shows the trajectories of the aircraft and the target within 12 seconds of launch. Figure 2 This shows the equivalent rudder deflection angle obtained in real time using Equation 1 within 12 seconds of launch. From Figure 1 and Figure 2 It can be seen that the integrated guidance and control method for high-dynamic aircraft uses the equivalent rudder deflection angle δ y and δz That is, the aircraft can be controlled to fly towards the target and finally hit the target.

[0146] Comparative Example

[0147] Based on the same initial conditions as the experimental example, that is, the initial position of the aircraft is (0 0 0), the initial position of the target is (5km 5km 5km), the initial speed of the aircraft is 600m / s, and the target has an irregular acceleration form,

[0148] An integrated guidance control method similar to that in the experimental example is used to guide and control the aircraft, the only difference being that in the experimental example, the derivative of the virtual control quantity is obtained by a tracking differentiator, while in the comparative example, the derivative of the virtual control quantity is obtained by a first-order low-pass filter; the final guidance terminal accuracy is 0.283m.

[0149] The intermediate variables required to obtain the control command in the experimental example and the comparative example, that is, the virtual control quantity, are called up and the virtual control quantity is continuously tracked, the virtual control quantity x 2d = [x 2dε x 2dη x 2dγ ] Τ The virtual control quantity obtained by tracking in the experimental example is denoted by subscript td, which is Figure 3 、 Figure 4 and Figure 5 in red dotted lines, and the virtual control quantity obtained by tracking in the comparative example is Figure 3 、 Figure 4 and Figure 5 in blue dot-dashed lines, denoted by subscript dsc;

[0150] The call results of the virtual control quantity are shown in Figure 3 、 Figure 4 and Figure 5 According to the call results, the virtual control quantity obtained in the experimental example is closer to the original variable than the virtual control quantity obtained in the comparative example, that is, closer to the green line in Figure 3 、 Figure 4 and Figure 5 , so that the entire control process has higher tracking accuracy, and thus better guidance accuracy can also be obtained.

[0151] The above describes the present application in combination with preferred embodiments, but these embodiments are only exemplary and serve only to illustrate. On this basis, various substitutions and improvements can be made to the present application, which all fall within the scope of protection of the present application.

Claims

1. An integrated guidance and control method for high-dynamic aircraft, characterized in that, In this guidance and control method, the control signals of the aircraft are obtained in real time, and the control signals are transmitted to the servo motors on the aircraft in real time. The control signals include the equivalent rudder deflection angles that control the aircraft's pitch, yaw, and roll. The servo motor operates according to the equivalent rudder deflection angle to adjust the flight state of the aircraft, thereby enabling the aircraft to hit the target; The control signal is obtained through the following formula (I): Where u represents the control signal. K3 represents the design parameters. s3 indicates error surface three. Both g3 and f3 represent intermediate variables. x represents the value obtained by the tracking differentiator. 3d The derivative; The error surface s3 is obtained by the following formula (ii): s3 = x3 - x 3d (two) Where x3 represents the angular velocity of the aircraft, x3=[ω z ω y ω x ] T , where ω z ω represents the pitch angular velocity of the aircraft. y ω represents the yaw rate of the aircraft. x This indicates the roll angular velocity of the aircraft; x 3d This represents the real-time tracking virtual control quantity of the three variables in x3; The x 3d We obtain it through the following formula (iii): Where K2 represents the design parameters, s2 represents error surface two. Both g2 and f2 represent intermediate variables. x represents the value obtained by the tracking differentiator. 2d The derivative of .

2. The integrated guidance and control method for high-dynamic aircraft according to claim 1, characterized in that, The error surface s2 is obtained by the following equation (iv): s2=x2-x 2d (Four) Where x2 represents the angle of the aircraft, x2 = [αβγ] T , where α represents the angle of attack of the aircraft, β represents the sideslip angle of the aircraft, and γ represents the roll angle of the aircraft; x 2d This represents the real-time tracking virtual control quantity of the three variables in x2.

3. The integrated guidance and control method for high-dynamic aircraft according to claim 2, characterized in that, The x 2d Obtained through the following formula (5): Where K1 represents the design parameters, s1 represents error surface one. Both f1 and g1 represent intermediate variables.

4. The integrated guidance and control method for high-dynamic aircraft according to claim 3, characterized in that, The error surface s1 is obtained by the following equation (vi): s1 = x1 (VI) Where x1 represents the line-of-sight angular velocity of the aircraft relative to the target. Where ε represents the line-of-sight angle relative to the target with the aircraft as the origin. η is the derivative of ε, representing the line-of-sight angular velocity in the pitch direction relative to the target with the aircraft as the origin, and η represents the line-of-sight deflection angle relative to the target with the aircraft as the origin. It is the derivative of η, representing the line-of-sight angular velocity relative to the target with the aircraft as the origin.

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