A new design method of attitude autopilot for guided aircraft

Abstract

The application discloses a novel guidance aircraft attitude autopilot design method, comprising the following steps: constructing a guidance aircraft attitude control system model; constructing an angle loop and an angular rate loop based on the guidance aircraft attitude control system model, the angle loop taking a reference angle as an input signal to obtain a virtual control rate, and the angular rate loop taking the virtual control rate as an input signal to obtain an attitude control instruction; and the guidance aircraft changing the attitude of the aircraft according to the attitude control instruction. The novel guidance aircraft attitude autopilot design method disclosed by the application estimates and compensates system disturbance through a state observer, reduces the error of the control system, and enhances the robustness of the system.

Description

TECHNICAL FIELD

[0001] The present application relates to a new type of guided aircraft attitude autopilot design method, belonging to the field of aircraft control. BACKGROUND

[0002] The spin gliding guided aircraft not only has high attack precision, but also has the advantage of low cost.

[0003] The roll frame attitude autopilot used by the spin gliding guided aircraft has a coupling effect in the pitch channel and the yaw channel, and the rudder of the spin gliding guided aircraft is small, which amplifies the influence of various interference factors on the aircraft, resulting in high difficulty in controlling the traditional spin gliding guided aircraft, insufficient control precision, and large control error.

[0004] Therefore, it is necessary to conduct more in-depth research on the attitude control of the spin gliding guided aircraft to solve the above problems. SUMMARY

[0005] In order to overcome the above problems, the present application has been studied in depth, and a new type of guided aircraft attitude autopilot design method is proposed, comprising the following steps:

[0006] Constructing a guided aircraft attitude control system model;

[0007] Based on the guided aircraft attitude control system model, an angle loop and an angular rate loop are constructed, the angle loop takes a reference angle as an input signal, and a virtual control rate The angular rate loop takes the virtual control rate as an input signal, and obtains an attitude control instruction U;

[0008] The guided aircraft changes the attitude of the aircraft according to the attitude control instruction U.

[0009] In a preferred embodiment, the guided aircraft attitude control system model is represented as:

[0010]

[0011]

[0012] wherein,

[0013]

[0014]

[0015]

[0016]

[0017]

[0018]

[0019] α represents an angle of attack, β represents a sideslip angle, θ represents a pitch angle, ω x represents a roll rate, ω y represents a yaw rate, ω z represents a pitch rate, δ y represents a rudder deflection, δ z represents an elevator deflection, d1 is a total disturbance in the angle loop, d2 is a total disturbance in the rate loop, ρ represents air density, V represents vehicle speed, S ref represents a vehicle characteristic area, m represents a vehicle mass, represents a partial derivative of a lift coefficient with respect to an angle of attack, represents a partial derivative of a lift coefficient with respect to a sideslip angle, represents a partial derivative of a pitch moment coefficient with respect to an angle of attack, represents a partial derivative of a yaw moment coefficient with respect to a sideslip angle, represents a partial derivative of a pitch damping moment coefficient with respect to a pitch rate, represents a partial derivative of a yaw damping moment coefficient with respect to a yaw rate, represents a partial derivative of a pitch moment coefficient with respect to an elevator deflection δ z , represents a partial derivative of a pitch moment coefficient with respect to a rudder deflection δ y , l represents a vehicle characteristic length, J x represents an x-direction moment of inertia, J y represents a y-direction moment of inertia, J z represents a z-direction moment of inertia.

[0020] In a preferred embodiment, the reference angles are set as where α ref is a reference angle of attack, β ref is a reference sideslip angle, both provided by a guidance command;

[0021] The angle loop includes the following sub-steps:

[0022] S201, based on the reference angles and a guidance vehicle attitude control system, a nominal rate is obtained, the nominal rate is represented as:

[0023]

[0024] S202, an angle loop control subsystem is set, represented as:

[0025]

[0026]

[0027] wherein E1 is the angle loop tracking error;

[0028] S203, setting a state observer to estimate the total disturbance in the angle loop control subsystem;

[0029] S204, obtaining the control rate of the angle loop control subsystem using a sliding mode control

[0030] S205, obtaining a virtual control rate based on the nominal angular rate and the control rate of the angle loop control subsystem is expressed as:

[0031]

[0032] In a preferred embodiment, in S203, the state observer is a reduced-order state observer, preferably set as:

[0033]

[0034] wherein represents the estimated value of d1, p1 represents the auxiliary variable of the state observer, and β1 represents the observer gain of the state observer.

[0035] In a preferred embodiment, in S204, the sliding surface S1 is set as:

[0036]

[0037] wherein τ represents the time constant, E1(0) represents the initial state of E1, is the nominal value of .

[0038] In a preferred embodiment, the angular rate loop includes the following sub-steps:

[0039] S301, setting a nominal control rate

[0040]

[0041] S302, setting the angle loop control subsystem, expressed as:

[0042]

[0043] ​

[0044] E2 is the angle rate loop tracking error;

[0045] S303, setting a state observer to estimate the total disturbance in the angle loop control subsystem;

[0046] S304, using a sliding mode control to obtain the control rate of the angle rate loop control subsystem

[0047] S305, based on the nominal control rate and the control rate of the angle rate loop control subsystem obtaining the attitude control instruction U, expressed as:

[0048]

[0049] In a preferred embodiment, in S302, the angle loop tracking error E1 is set as:

[0050]

[0051] wherein, obtained by measuring the aircraft sensors.

[0052] In a preferred embodiment, in S303, the state observer is a reduced-order state observer, preferably set as:

[0053]

[0054] wherein, represents the estimated value of d2, p2 represents the auxiliary variable of the state observer, and β2 represents the observer gain of the state observer.

[0055] In a preferred embodiment, in S304, the sliding surface S2 is set as:

[0056]

[0057] wherein, E2(0) represents the initial state of E2, is the nominal value of .

[0058] In a preferred embodiment, the obtained control rate of the angle rate loop control subsystem is expressed as:

[0059]

[0060]

[0061] wherein, η21 , η 22 , η 23 is a coefficient matrix, and v1, v2 are constants.

[0062] The present application has the beneficial effects including:

[0063] (1) The system state variable converges in fixed time, improving the convergence speed;

[0064] (2) The system disturbance is estimated and compensated by the state observer, reducing the error of the control system and enhancing the robustness of the system. BRIEF DESCRIPTION OF DRAWINGS

[0065] Figure 1 A flow chart of a design method of a new type of guided vehicle attitude autopilot according to a preferred embodiment of the present application is shown;

[0066] Figure 2 A simulation result of an angle loop tracking curve in Example 1 is shown;

[0067] Figure 3 A simulation result of an angular rate loop tracking curve in Example 1 is shown;

[0068] Figure 4 A simulation result of a change curve of equivalent pitch rudder deflection angle and equivalent yaw rudder deflection angle in Example 1 is shown;

[0069] Figure 5 A simulation result of a tracking error curve in Example 1 is shown;

[0070] Figure 6 A simulation result of estimation of disturbance d1 by the state observer in Example 1 is shown;

[0071] Figure 7 A simulation result of estimation of disturbance d2 by the state observer in Example 1 is shown. DETAILED DESCRIPTION

[0072] The present application will be further described in detail by the accompanying drawings and examples. Through these descriptions, the features and advantages of the present application will become more apparent.

[0073] The word "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 the disclosure is not limited to the specific embodiments illustrated in the drawings.

[0074] The control of the guided aircraft includes control of the aircraft mass center movement and control of the aircraft attitude, wherein the control of the aircraft mass center movement utilizes the seeker to measure the relative position of the missile to the target or the deviation from the predetermined orbit to form a guidance command; the control of the aircraft attitude utilizes the guidance command to control the actuator to generate the force required for controlling the missile.

[0075] Further, in the present application, the guided aircraft is controlled by a canard so as to be launched from the barrel to effectively improve the lift-drag ratio.

[0076] Further, the tail wing of the guided aircraft is designed to be pop-off and has a certain oblique angle, so as to be launched from the barrel, and meanwhile the missile body can self-rotate at a certain angular velocity, which not only saves the additional stability control on the roll channel and reduces the cost of the guidance control system, but also effectively reduces the adverse effects of the aerodynamic shape asymmetry and the eccentricity of the rocket engine thrust.

[0077] According to the present application, a new design method of the attitude autopilot of the guided aircraft is provided, which comprises the following steps as shown in the figure: Figure 1

[0078] constructing a model of the attitude control system of the guided aircraft;

[0079] constructing an angle loop and an angular velocity loop based on the model of the attitude control system of the guided aircraft, wherein the angle loop takes a reference angle as an input signal to obtain a virtual control rate the angular velocity loop takes the virtual control rate as an input signal to obtain an attitude control command U;

[0080] the guided aircraft changes the attitude of the aircraft according to the attitude control command U.

[0081] In a preferred embodiment, the model of the attitude control system of the guided aircraft is represented as:

[0082]

[0083]

[0084] wherein,

[0085]

[0086]

[0087]

[0088]

[0089] ​

[0090]

[0091] α represents an angle of attack, β represents a sideslip angle, θ represents a pitch angle, ω x represents a roll rate, ω y represents a yaw rate, ω z represents a pitch rate, δ y represents a rudder deflection angle, δ z represents an elevator deflection angle, d1 is a total disturbance in the angle loop, including system modeling error, aerodynamic parameter disturbance and external disturbance, d2 is a total disturbance in the angular rate loop, including system modeling error, aerodynamic parameter disturbance and external disturbance, ρ represents air density, V represents aircraft speed, S ref represents aircraft characteristic area, m represents aircraft mass, represents a partial derivative of a lift coefficient with respect to an angle of attack, represents a partial derivative of a lift coefficient with respect to a sideslip angle, represents a partial derivative of a pitch moment coefficient with respect to an angle of attack, represents a partial derivative of a yaw moment coefficient with respect to a sideslip angle, represents a partial derivative of a pitch damping moment coefficient with respect to a pitch rate, represents a partial derivative of a yaw damping moment coefficient with respect to a yaw rate, represents a partial derivative of a pitch moment coefficient with respect to a rudder deflection angle δ z , represents a partial derivative of a pitch moment coefficient with respect to an elevator deflection angle δ y , l represents an aircraft characteristic length, J x represents an x-direction rotational inertia, J y represents a y-direction rotational inertia, J z represents a z-direction rotational inertia.

[0092] According to the present application, the reference angle is set as wherein α ref is a reference angle of attack, β ref is a reference sideslip angle, both of which are provided by a guidance command;

[0093] According to the present application, the angle loop comprises the following sub-steps:

[0094] S201, based on the reference angle and the guidance aircraft attitude control system, a nominal angular rate is obtained, the nominal angular rate is represented as:

[0095]

[0096] S202, setting an angle loop control subsystem, denoted as:

[0097]

[0098]

[0099] wherein E1 is an angle loop tracking error;

[0100] S203, setting a state observer to estimate total disturbance in the angle loop control subsystem;

[0101] S204, obtaining a control rate of the angle loop control subsystem using a sliding mode control

[0102] S205, obtaining a virtual control rate based on a nominal angular rate and the control rate of the angle loop control subsystem denoted as:

[0103]

[0104] In one preferred embodiment, in S201, the derivative of the reference angle is obtained by a low-pass filter, which is set as:

[0105]

[0106] wherein τ1 is a time constant to be designed, and τ1>0.

[0107] In one preferred embodiment, in S202, the angle loop tracking error E1 is set as:

[0108]

[0109] wherein, obtained by measuring the aircraft sensors.

[0110] In one preferred embodiment, in S203, the state observer is a reduced-order state observer, which is preferably set as:

[0111]

[0112] wherein, denotes the estimated value of d1, p1 denotes an auxiliary variable of the state observer, and β1 denotes an observer gain of the state observer.

[0113] In one preferred embodiment, in S204, the sliding surface S1 is set as:

[0114] ​

[0115] where τ denotes the time constant, E1(0) represents the initial state of E1, is the nominal value of .

[0116] Preferably, is set to

[0117]

[0118] where χ 11 , χ 12 , χ 13 is a coefficient matrix, m1, m2 are constants, and have m1 > 1, 0 < m2 < 1.

[0119] More preferably, the coefficient matrix χ 11 , χ 12 , χ 13 is set to

[0120] χ 1i = diag(χ 1i,1 , χ 1i,2 )

[0121] where i denotes different elements, χ 1i,1 > 0, χ 1i,2 > 0, i = 1, 2, 3.

[0122] According to the present application, the control rate of the angle loop control subsystem obtained is expressed as:

[0123]

[0124]

[0125] where η 11 , η 12 , η 13 is a coefficient matrix, n1, n2 are constants, and have n1 > 1, 0 < n2 < 1.

[0126] More preferably, the coefficient matrix η 11 , η 12 , η 13 is set to

[0127] η 1i = diag(η 1i，1 , η 1i，2 )

[0128] where η 1i,1 > 0, η 1i,2 > 0, i = 1, 2, 3.

[0129] In one preferred embodiment, the angular rate loop comprises the following sub-steps:

[0130] S301, setting a nominal control rate

[0131]

[0132] S302, setting an angular loop control subsystem, denoted as:

[0133]

[0134]

[0135] wherein E2 is the angular rate loop tracking error;

[0136] S303, setting a state observer to estimate the total disturbance in the angular loop control subsystem;

[0137] S304, using a sliding mode control to obtain the control rate of the angular rate loop control subsystem

[0138] S305, obtaining a pose control command U based on the nominal control rate and the control rate of the angular rate loop control subsystem denoted as:

[0139]

[0140] In one preferred embodiment, in S301, the derivative of the virtual control rate is obtained through a low-pass filter, which is set as:

[0141]

[0142] wherein τ2 is a time constant to be designed, and τ2 > 0.

[0143] In one preferred embodiment, in S302, the angular loop tracking error E1 is set as:

[0144]

[0145] wherein is obtained through measurements of the aircraft sensors.

[0146] In one preferred embodiment, in S303, the state observer is a reduced-order state observer, which is preferably set as:

[0147]

[0148] wherein, represents an estimated value of d2, p2 represents an auxiliary variable of the state observer, and β2 represents an observer gain of the state observer.

[0149] In a preferred embodiment, in S304, the sliding surface S2 is set as:

[0150]

[0151] wherein E2(0) represents an initial state of E2, is a nominal value of .

[0152] Preferably, is set as:

[0153]

[0154] wherein χ 21 , χ 22 , χ 23 is a coefficient matrix, and u1, u2 are constants, and have u1>1, 0

[0155] More preferably, the coefficient matrix χ 21 , χ 22 , χ 23 is set as:

[0156] χ 2i = diag(χ 2i,1 , χ 2i,2 )

[0157] wherein i represents different elements, χ 2i,1 >0, χ 2i,2 >0, i=1, 2, 3.

[0158] According to the present application, the control rate of the angular rate loop control subsystem obtained is represented as:

[0159]

[0160]

[0161] wherein η 21 , η 22 , η 23 is a coefficient matrix, and v1, v2 are constants.

[0162] More preferably, the coefficient matrix η 21 , η 22 , η 23 is set as:

[0163] η 2i = diag(η 2i,1 ,η 2i,2 )

[0164] wherein η 2i,1 > 0, η 2i,2 > 0, i = 1, 2, 3.

[0165] Embodiment

[0166] Embodiment 1

[0167] A simulation experiment is performed to obtain the attitude control instruction of the guided aircraft, including the following steps:

[0168] including the following steps:

[0169] A model of the attitude control system of the guided aircraft is constructed;

[0170] Based on the model of the attitude control system of the guided aircraft, an angle loop and an angular rate loop are constructed, the angle loop taking the reference angle as an input signal to obtain a virtual control rate The angular rate loop takes the virtual control rate as an input signal to obtain the attitude control instruction U;

[0171] The model of the attitude control system of the guided aircraft is represented as:

[0172]

[0173]

[0174] wherein,

[0175]

[0176]

[0177]

[0178]

[0179]

[0180]

[0181] The angle loop includes the following sub-steps:

[0182] S201, based on the reference angle and the attitude control system of the guided aircraft, a nominal angular rate is obtained, the nominal angular rate is represented as:

[0183]

[0184] S202, setting an angle loop control subsystem, denoted as:

[0185]

[0186]

[0187] S203, setting a state observer to estimate total disturbance in the angle loop control subsystem;

[0188] S204, using a sliding mode control to obtain a control rate of the angle loop control subsystem

[0189] S205, based on a nominal angular rate and a control rate of the angle loop control subsystem obtaining a virtual control rate denoted as:

[0190]

[0191] In S201, the differential of the reference angle is obtained through a low-pass filter, which is set as:

[0192]

[0193] In S202, the angle loop tracking error E1 is set as:

[0194]

[0195] In S203, the state observer is a reduced-order state observer, which is preferably set as:

[0196]

[0197] In S204, the sliding surface S1 is set as:

[0198]

[0199]

[0200] obtaining a control rate of the angle loop control subsystem denoted as:

[0201]

[0202]

[0203] The angular rate loop includes the following sub-steps:

[0204] S301, setting a nominal control rate

[0205]

[0206] S302, set angle loop control subsystem, denoted as:

[0207]

[0208]

[0209] S303, set state observer to estimate the total disturbance in angle loop control subsystem;

[0210] S304, using sliding mode control to obtain the control rate of angular rate loop control subsystem

[0211] S305, based on the nominal control rate And the control rate of angular rate loop control subsystem Obtain attitude control instruction U, denoted as:

[0212]

[0213] In S301, the differential of virtual control rate Obtained by low-pass filter, the low-pass filter is set to:

[0214]

[0215] In S302, the angle loop tracking error E1 is set to:

[0216]

[0217] In S303, the state observer is set to:

[0218]

[0219] In S304, set the sliding surface S2 to:

[0220]

[0221]

[0222] The control rate of angular rate loop control subsystem obtained Denoted as:

[0223]

[0224]

[0225] In the simulation process, the disturbance is set as:

[0226]

[0227] Reference input instruction is designed as:

[0228] Considering the limitations of the actual guided shell actuator, the rudder angle is limited:

[0229]

[0230] In the simulation, the relevant parameters are set as:

[0231] χ 11 = χ 12 = χ 13 = χ 21 = χ 22 = χ 23 = η 11 = η 12 = η 13 = η 21 = η 22 = η 23 = diag(0.01,0.01),

[0232] m1 = n1 = μ1 = ν1 = 2-7 / 9, m2 = n2 = μ2 = ν2 = 7 / 9, β1 = 100, β2 = 20.

[0233] The simulation results are shown in Figures 2-7 .

[0234] wherein, Figure 2 shows the angle loop tracking curve simulation results, where α, β are the actual angles obtained by simulation, α ref , β ref are the input reference angle, from the figure, it can be seen that in the presence of external disturbance, the angle of attack and sideslip angle can be well tracked by the command signal, the adjustment time is less than 2s.

[0235] Figure 3 shows the angular rate loop tracking curve simulation results, where ω z , ω y are the actual angular velocity obtained by simulation, ω zref , ω yref are the reference angular velocity obtained by angle loop, from the figure, it can be seen that the pitch rate and yaw rate can quickly track the corresponding virtual control command, the adjustment time is less than 0.5s.

[0236] Figure 4The variation curves of equivalent pitch rudder deflection angle and equivalent yaw rudder deflection angle are shown, and it can be seen from the figure that the amplitude is always within the limit of the maximum rudder deflection angle, the equivalent input control instruction curve is smooth, and the engineering requirements are met.

[0237] Figure 5 The tracking error curve simulation results are shown, wherein E1(1) is the error of the actual value and the reference value of the attack angle, E2(1) is the error of the actual value and the reference value of the pitch angle rate, E1(2) is the error of the actual value and the reference value of the sideslip angle, and E2(2) is the error of the actual value and the reference value of the yaw angle rate, and it can be seen from the figure that the error can quickly converge in the presence of external interference.

[0238] Figure 6 、 7 The estimation effect of the state observer on the disturbances d1 and d2 is shown, wherein d1(1) is the total disturbance component 1 of the angle loop, d1(1) esm is the estimated value of the total disturbance component 1 of the angle loop, d1(2) is the total disturbance component 2 of the angle loop, d1(2) esm is the estimated value of the total disturbance component 2 of the angle loop, d2(1) is the total disturbance component 1 of the angular rate loop, d2(1) esm is the estimated value of the total disturbance component 1 of the angular rate loop, d2(2) is the total disturbance component 2 of the angular rate loop, and d2(2) esm is the estimated value of the total disturbance component 2 of the angular rate loop, and it can be obviously seen from the figure that the reduced-order state observer can quickly and accurately estimate the disturbance value, and the control accuracy of the closed-loop system is improved.

[0239] In the description of the present application, it should be pointed out that the terms "upper", "lower", "inner", "outer", "front", "rear" and the like indicate the orientation or positional relationship in the working state of the present application, and are only for the convenience of describing the present application and simplifying the description, and therefore cannot be understood as indicating or implying that the devices or elements referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as limiting the present application. In addition, the terms "first", "second", "third", "fourth" are only for descriptive purposes and cannot be understood as indicating or implying relative importance.

[0240] In the description of the present application, it should be pointed out that unless otherwise explicitly specified and limited, the terms "mounting", "connection", "connection" should be understood in a broad sense, for example, it can be fixed connection, or detachable connection, or integral connection; it can be mechanical connection, or electrical connection; it can be direct connection, or indirect connection through intermediate medium; it can be the communication inside two elements. For ordinary skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0241] The application has been described above with reference to preferred embodiments. However, these embodiments are merely exemplary and are intended to be illustrative only. Various substitutions and alterations are possible in view of the disclosure of this application without departing from the spirit and scope of the application.

Claims

1. A novel design method for an autopilot of a guided aircraft, characterized in that, The method comprises the following steps: constructing a model of a guidance vehicle attitude control system; Based on the model of the attitude control system of the guided aircraft, an angle loop and an angular rate loop are constructed, the angle loop taking a reference angle As an input signal, a virtual control rate is acquired The angular rate loop takes the virtual control rate As an input signal, an attitude control instruction U is acquired the guidance vehicle changing the attitude of the vehicle according to the attitude control instruction U, the model of the guidance vehicle attitude control system is expressed as: wherein, α denotes the angle of attack, β denotes the angle of sideslip, θ denotes the angle of pitch, ω x denotes the roll rate, ω y denotes the yaw rate, ω z denotes the pitch rate, δ y denotes the rudder deflection, δ z denotes the elevator deflection, di denotes the total disturbance in the angle loop, d2 denotes the total disturbance in the rate loop, p denotes the air density, V denotes the vehicle velocity, S ref denotes the vehicle characteristic area, m denotes the vehicle mass, denotes the partial derivative of the lift coefficient with respect to the angle of attack, denotes the partial derivative of the lift coefficient with respect to the angle of sideslip, denotes the partial derivative of the pitch moment coefficient with respect to the angle of attack, denotes the partial derivative of the yaw moment coefficient with respect to the angle of sideslip, denotes the partial derivative of the pitch damping moment coefficient with respect to the pitch rate, denotes the partial derivative of the yaw damping moment coefficient with respect to the yaw rate, denotes the partial derivative of the pitch moment coefficient with respect to the elevator deflection δ z , denotes the partial derivative of the pitch moment coefficient with respect to the elevator deflection δ y , I denotes the vehicle characteristic length, J x denotes the x-direction moment of inertia, J y denotes the y-direction moment of inertia, J z denotes the z-direction moment of inertia, The reference angle is set as where α ref is a reference angle of attack, β ref is a reference sideslip angle, both provided by the guidance command; the angle loop comprises the following sub-steps: S201、based on the reference angle and the guided aircraft attitude control system, a nominal angular rate is obtained, and the nominal angular rate is represented as: S202, setting an angle loop control subsystem, expressed as: wherein, E1 is an angle loop tracking error; S203, setting a state observer to estimate total disturbance in the angle loop control subsystem; S204, obtaining the control rate of the angle loop control subsystem by using the sliding mode control S205、based on the nominal angular rate and the control rate of the angle loop control subsystem obtaining a virtual control rate is expressed as:

2. The guidance vehicle attitude autopilot design method according to claim 1, wherein, in S203, the state observer is a reduced-order state observer.

3. The guided vehicle attitude autopilot design method of claim 2, wherein, the state observer is set as: wherein represents an estimate of d1, pi represents an auxiliary variable of the state observer, and β1represents an observer gain of the state observer.

4. The guidance vehicle attitude autopilot design method according to claim 1, wherein, in S204, the sliding surface S1 is set as: where τ represents a time constant, Ei(0) represents an initial state of Ei, is a nominal value of is a nominal value of 5. The guidance vehicle attitude autopilot design method according to claim 1, wherein, the angular rate loop comprises the following sub-steps: S301, set a nominal control rate S302, setting an angle loop control subsystem, expressed as: wherein, E2 is an angular rate loop tracking error; S303, setting a state observer to estimate total disturbance in the angle loop control subsystem; S304, obtaining the control rate of the angular rate loop control subsystem by using the sliding mode control S305、based on the nominal control rate and the control rate of the angular rate loop control subsystem Obtain the attitude control instruction U, denoted as:

6. The guidance vehicle attitude autopilot design method according to claim 5, wherein, in S302, the angle loop tracking error E1 is set as: wherein, Obtained by aircraft sensor measurements.

7. The guidance vehicle attitude autopilot design method according to claim 5, wherein, in S303, the state observer is a reduced-order state observer.

8. The guided vehicle attitude autopilot design method of claim 7, wherein, the state observer is set as: wherein represents an estimate of d2, p2 represents an auxiliary variable of the state observer, and β2 represents an observer gain of the state observer.

9. The guidance vehicle attitude autopilot design method according to claim 5, wherein, in S304, the sliding surface S2 is set as: where E2(0) represents the initial state of E2, is the nominal value of E2.

10. The guidance vehicle attitude autopilot design method according to claim 5, wherein, The control rate of the angular rate loop control subsystem obtained is represented as: wherein η 21 , η 22 , η 23 are coefficient matrices, and v1, v2 are constants.

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