Novel guidance aircraft attitude autopilot design method

Abstract

The invention discloses a novel guidance aircraft attitude autopilot design method. The method comprises the following steps: constructing a guidance aircraft attitude control system model; based on a 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 to obtain a virtual control rate, and the angular rate loop takes the virtual control rate as an input signal to obtain an attitude control instruction; and guiding the aircraft to change the attitude of the aircraft according to the attitude control instruction. According to the novel guided aircraft attitude autopilot design method disclosed by the invention, system disturbance is estimated and compensated through the state observer, errors of a control system are reduced, and the robustness of the system is enhanced.

Description

Technical Field

[0001] The present invention relates to a design method for an attitude autopilot of a new type of guided vehicle, belonging to the field of vehicle control. Background Art

[0002] The spin-glide guided vehicle not only has high strike accuracy but also has the advantage of low cost.

[0003] The rolling-frame attitude pilot used in the spin-glide guided vehicle has coupling effects in the pitch channel and yaw channel, and the control surfaces of the spin-glide guided vehicle are small, which amplifies the influence of various interference factors on the vehicle, resulting in high difficulty in attitude control, insufficient control accuracy, and large control errors for traditional spin-glide guided vehicles.

[0004] Therefore, it is necessary to conduct a more in-depth study on the attitude control of spin-glide guided vehicles to solve the above problems. Summary of the Invention

[0005] In order to overcome the above problems, the inventors of the present invention have conducted in-depth research and proposed a design method for an attitude autopilot of a new type of guided vehicle, including the following steps:

[0006] Construct a model of the attitude control system of the guided vehicle;

[0007] Based on the model of the attitude control system of the guided vehicle, construct an angle loop and an angular rate loop. The angle loop uses the reference angle as the input signal to obtain the virtual control rate The angular rate loop uses the virtual control rate as the input signal to obtain the attitude control command U;

[0008] The guided vehicle changes its attitude according to the attitude control command U.

[0009] In a preferred embodiment, the model of the attitude control system of the guided vehicle is expressed as:

[0010]

[0011]

[0012] Wherein,

[0013]

[0014]

[0015]

[0016]

[0017]

[0018]

[0019] α represents the angle of attack, β represents the sideslip angle, θ represents the pitch angle, ω x represents the roll rate, ω y represents the yaw rate, ω z represents the pitch rate, δ y represents the yaw rudder deflection angle, δ z represents the pitch rudder deflection angle, d 1 is the total disturbance in the angle loop, d 2 is the total disturbance in the angular rate loop, ρ represents the air density, V represents the aircraft speed, S ref represents the characteristic area of the aircraft, m represents the mass of the aircraft, represents the partial derivative of the lift coefficient with respect to the angle of attack, represents the partial derivative of the lift coefficient with respect to the sideslip angle, represents the partial derivative of the pitch moment coefficient with respect to the angle of attack, represents the partial derivative of the yaw moment coefficient with respect to the sideslip angle, represents the partial derivative of the pitch damping moment coefficient with respect to the pitch rate, represents the partial derivative of the yaw damping moment coefficient with respect to the yaw rate, represents the pitch moment coefficient with respect to the rudder deflection angle δ z of the partial derivative, represents the pitch moment coefficient with respect to the rudder deflection angle δ y of the partial derivative, l represents the characteristic length of the aircraft, J x represents the moment of inertia about the x-axis, J y represents the moment of inertia about the y-axis, J z represents the moment of inertia about the z-axis.

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

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

[0022] S201. Based on the reference angle and the guidance aircraft attitude control system, obtain the nominal angular rate, and the nominal angular rate is expressed as:

[0023]

[0024] S202. Set up the angle loop control subsystem, expressed as:

[0025]

[0026]

[0027] where E 1 is the angle loop tracking error;

[0028] S203. Set up a state observer to estimate the total disturbance in the angle loop control subsystem;

[0029] S204. Obtain the control law of the angle loop control subsystem by using sliding mode control

[0030] S205. Based on the nominal angular rate and the control law of the angle loop control subsystem obtain the virtual control law expressed as:

[0031]

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

[0033]

[0034] where represents the estimated value of d 1 p 1 represents the auxiliary variable of the state observer, and β 1 represents the observer gain of the state observer.

[0035] In a preferred embodiment, in S204, set the sliding mode surface S 1 as:

[0036]

[0037] where τ represents the time constant, and E 1 (0) represents the initial state of E 1 , is the nominal value of.

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

[0039] S301. Set the nominal control law

[0040]

[0041] S302. Set up the angle loop control subsystem, expressed as:

[0042]

[0043]

[0044] where E 2 is the tracking error of the angular rate loop;

[0045] S303. Set up a state observer to estimate the total disturbance in the angle loop control subsystem;

[0046] S304. Obtain the control law of the angular rate loop control subsystem by using sliding mode control

[0047] S305. Based on the nominal control law and the control law of the angular rate loop control subsystem obtain the attitude control command U, expressed as:

[0048]

[0049] In a preferred embodiment, in S302, the angle loop tracking error E 1 is set to:

[0050]

[0051] where is obtained by measuring through the aircraft sensor.

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

[0053]

[0054] where represents the estimated value of d 2 and p 2 represents the auxiliary variable of the state observer, and β 2 represents the observer gain of the state observer.

[0055] In a preferred embodiment, in S304, set the sliding mode surface S 2 to:

[0056]

[0057] where E 2 (0) represents the initial state of E 2 and is the nominal value.

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

[0059]

[0060]

[0061] where η 21 , η 22 , η 23 are coefficient matrices, and v 1 , v 2 are constants.

[0062] The beneficial effects of the present invention include:

[0063] (1) The system state variables converge within a fixed time, improving the convergence speed;

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

[0065] Figure 1 Shows a schematic flow diagram of a new guidance vehicle attitude autopilot design method according to a preferred embodiment of the present invention;

[0066] Figure 2 Shows the simulation results of the angle loop tracking curve in Example 1;

[0067] Figure 3 Shows the simulation results of the angular rate loop tracking curve in Example 1;

[0068] Figure 4 Shows the simulation results of the change curves of the equivalent pitch rudder deflection angle and the equivalent yaw rudder deflection angle in Example 1;

[0069] Figure 5 Shows the simulation results of the tracking error curve in Example 1;

[0070] Figure 6 Shows the estimation simulation results of the disturbance d 1 by the state observer in Example 1;

[0071] Figure 7 Shows the estimation simulation results of the disturbance d 2 by the state observer in Example 1. DETAILED DESCRIPTION OF THE INVENTION

[0072] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Through these descriptions, the features and advantages of the present invention will become more clearly defined.

[0073] As used herein, the term "exemplary" means "serving as an example, embodiment, or illustration". Any embodiment described herein as "exemplary" is not necessarily to be construed as superior or better than other embodiments. Although various aspects of the embodiments are shown in the drawings, the drawings do not have to be drawn to scale unless otherwise specified.

[0074] The control of a guided vehicle includes the control of the movement of the vehicle's center of mass and the control of the vehicle's attitude. Among them, the control of the movement of the vehicle's center of mass uses a seeker to measure the relative position between the missile and the target or the deviation from a predetermined trajectory to form a guidance command; the control of the vehicle's attitude uses the guidance command to control an actuator to generate the force required to control the missile.

[0075] Furthermore, in the present invention, the guided vehicle is controlled by canards to facilitate launch from a gun barrel, thereby effectively increasing the lift-to-drag ratio.

[0076] Even further, the tail fins of the guided vehicle are designed to be spring-loaded and have a certain skew angle, which facilitates launch from a gun barrel. At the same time, the missile body can rotate at a certain angular velocity, which not only eliminates the need for additional stability control of the roll channel and reduces the cost of the guidance control system, but also effectively reduces the adverse effects of aerodynamic shape asymmetry and rocket engine thrust eccentricity.

[0077] According to a novel design method for the attitude autopilot of a guided vehicle provided by the present invention, as Figure 1 shown, it includes the following steps:

[0078] Construct a model of the attitude control system of the guided vehicle;

[0079] Based on the model of the attitude control system of the guided vehicle, construct an angle loop and an angular rate loop. The angle loop uses the reference angle as the input signal to obtain the virtual control rate The angular rate loop uses the virtual control rate as the input signal to obtain the attitude control command U;

[0080] The guided vehicle changes its attitude according to the attitude control command U.

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

[0082]

[0083]

[0084] Among them,

[0085]

[0086]

[0087]

[0088]

[0089]

[0090]

[0091] α represents the angle of attack, β represents the sideslip angle, θ represents the pitch angle, ω x represents the roll rate, ω y represents the yaw rate, ω z represents the pitch rate, δ y represents the yaw rudder deflection angle, δ z represents the pitch rudder deflection angle, d 1 is the total disturbance in the angle loop, including system modeling errors, aerodynamic parameter disturbances, and external disturbances, d 2 is the total disturbance in the angular rate loop, including system modeling errors, aerodynamic parameter disturbances, and external disturbances, ρ represents the air density, V represents the aircraft speed, S ref represents the characteristic area of the aircraft, m represents the mass of the aircraft, represents the partial derivative of the lift coefficient with respect to the angle of attack, represents the partial derivative of the lift coefficient with respect to the sideslip angle, represents the partial derivative of the pitch moment coefficient with respect to the angle of attack, represents the partial derivative of the yaw moment coefficient with respect to the sideslip angle, represents the partial derivative of the pitch damping moment coefficient with respect to the pitch rate, represents the partial derivative of the yaw damping moment coefficient with respect to the yaw rate, represents the pitch moment coefficient with respect to the rudder deflection angle δ z partial derivative of, represents the pitch moment coefficient with respect to the rudder deflection angle δ y partial derivative of, l represents the characteristic length of the aircraft, J x represents the moment of inertia about the x-axis, J y represents the moment of inertia about the y-axis, J z represents the moment of inertia about the z-axis.

[0092] According to the present invention, the reference angle is set to where α ref is the reference angle of attack, β refis the reference sideslip angle, both provided by the guidance command;

[0093] According to the present invention, the angle loop includes the following sub-steps:

[0094] S201. Based on the reference angle and the attitude control system of the guided vehicle, obtain the nominal angular rate, and the nominal angular rate is expressed as:

[0095]

[0096] S202. Set the angle loop control subsystem, expressed as:

[0097]

[0098]

[0099] where E 1 is the angle loop tracking error;

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

[0101] S204. Use sliding mode control to obtain the control law of the angle loop control subsystem

[0102] S205. Based on the nominal angular rate and the control law of the angle loop control subsystem obtain the virtual control law expressed as:

[0103]

[0104] In a preferred embodiment, in S201, the differential of the reference angle is obtained through a low-pass filter, and the low-pass filter is set as:

[0105]

[0106] where τ 1 is the time constant to be designed, and τ 1 > 0.

[0107] In a preferred embodiment, in S202, the angle loop tracking error E 1 is set as:

[0108]

[0109] where is obtained by measuring through the vehicle sensor.

[0110] In a preferred embodiment, in S203, the state observer is a reduced-order state observer. Preferably, it is set as:

[0111]

[0112] Wherein, represents the estimated value of d 1 , p 1 represents the auxiliary variable of the state observer, and β 1 represents the observer gain of the state observer.

[0113] In a preferred embodiment, in S204, the sliding mode surface S 1 is set as:

[0114]

[0115] Wherein, τ represents the time constant, and E 1 (0) represents the initial state of E 1 , is the nominal value of.

[0116] Preferably, it is set as:

[0117]

[0118] Wherein, χ 11 , χ 12 , χ 13 are coefficient matrices, m 1 , m 2 are constants, and there is m 1 > 1, 0 < m 2 < 1.

[0119] More preferably, the coefficient matrices χ 11 , χ 12 , χ 13 are set as:

[0120] χ 1i = diag(χ 1i,1 , χ 1i,2 )

[0121] Wherein, i represents different elements, χ 1i,1 > 0, χ 1i,2 > 0, i = 1, 2, 3.

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

[0123]

[0124]

[0125] Among them, η 11 、η 12 、η 13 are coefficient matrices, n 1 、n 2 are constants, and there is n 1 > 1, 0 < n 2 < 1.

[0126] More preferably, the coefficient matrices η 11 、η 12 、η 13 are set as:

[0127] η 1i = diag(η 1i，1 , η 1i，2 )

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

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

[0130] S301. Set the nominal control rate

[0131]

[0132] S302. Set the angle loop control subsystem, expressed as:

[0133]

[0134]

[0135] Among them, E 2 is the angular rate loop tracking error;

[0136] S303. Set a state observer to estimate the total disturbance in the angle loop control subsystem;

[0137] S304. Obtain the control rate

[0138] of the angular rate loop control subsystem by using 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 command U, expressed as:

[0139]

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

[0141]

[0142] where τ 2 is the time constant to be designed, and τ 2 > 0.

[0143] In a preferred embodiment, in S302, the angle loop tracking error E 1 is set as:

[0144]

[0145] where is obtained by measuring with an aircraft sensor.

[0146] In a preferred embodiment, in S303, the state observer is a reduced-order state observer. Preferably, it is set as:

[0147]

[0148] where represents the estimated value of d 2 , p 2 represents the auxiliary variable of the state observer, and β 2 represents the observer gain of the state observer.

[0149] In a preferred embodiment, in S304, the sliding mode surface S 2 is set as:

[0150]

[0151] where E 2 (0) represents the initial state of E 2 , is the nominal value of.

[0152] Preferably, is set as:

[0153]

[0154] where χ 21 , χ 22 , χ 23 are coefficient matrices, u 1 , u 2 are constants, and there is u 1 > 1, 0 < u2 < 1.

[0155] More preferably, the coefficient matrices χ 21 , χ 22 , χ 23 are set as:

[0156] χ 2i = diag(χ 2i,1 , χ 2i,2 )

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

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

[0159]

[0160]

[0161] where η 21 , η 22 , η 23 are coefficient matrices, and v 1 , v 2 are constants.

[0162] More preferably, the coefficient matrices η 21 , η 22 , η 23 are set as:

[0163] η 2i = diag(η 2i,1 , η 2i,2 )

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

[0165] Embodiment

[0166] Embodiment 1

[0167] A simulation experiment is carried out to obtain the attitude control command of the guided vehicle, including the following steps:

[0168] Including the following steps:

[0169] Construct a model of the attitude control system of the guided vehicle;

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

[0171] The attitude control system model of the guided vehicle is expressed as:

[0172]

[0173]

[0174] wherein,

[0175]

[0176]

[0177]

[0178]

[0179]

[0180]

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

[0182] S201. Obtain the nominal angular rate based on the reference angle and the attitude control system of the guided vehicle, and the nominal angular rate is expressed as:

[0183]

[0184] S202. Set up the angle loop control subsystem, which is expressed as:

[0185]

[0186]

[0187] S203. Set up a state observer to estimate the total disturbance in the angle loop control subsystem;

[0188] S204. Obtain the control rate of the angle loop control subsystem by using sliding mode control

[0189] S205. Based on the nominal angular rate and the control rate of the angle loop control subsystem obtain the virtual control rate which is expressed as:

[0190]

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

[0192]

[0193] In S202, the angle loop tracking error E 1 is set to:

[0194]

[0195] In S203, the state observer is a reduced-order state observer, and preferably, it is set to:

[0196]

[0197] In S204, the sliding mode surface S 1 is:

[0198]

[0199]

[0200] The control rate of the obtained angle loop control subsystem is expressed as:

[0201]

[0202]

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

[0204] S301. Set the nominal control rate

[0205]

[0206] S302. Set the angle loop control subsystem, which is expressed as:

[0207]

[0208]

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

[0210] S304. Obtain the control rate of the angular rate loop control subsystem by using sliding mode control

[0211] S305. Based on the nominal control rate Control law of angular rate loop control subsystem Obtain the attitude control command U, expressed as:

[0212]

[0213] In S301, the differential of the virtual control law Is obtained through a low-pass filter, and the low-pass filter is set as:

[0214]

[0215] In S302, the angle loop tracking error E 1 Is set as:

[0216]

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

[0218]

[0219] In S304, set the sliding mode surface S 2 As:

[0220]

[0221]

[0222] The obtained control law of the angular rate loop control subsystem Is expressed as:

[0223]

[0224]

[0225] During the simulation, set the disturbance as:

[0226]

[0227] Reference input command Is designed as:

[0228] Considering the limitations of the actual actuator of the guided projectile, limit the rudder deflection angle:

[0229]

[0230] The relevant parameters in the simulation are set as:

[0231] χ 11 = χ 12 = χ 13 = χ 21 = χ22 = χ 23 = η 11 = η 12 = η 13 = η 21 = η 22 = η 23 = diag(0.01, 0.01),

[0232] m 1 = n 1 = μ 1 = ν 1 = 2 - 7 / 9, m 2 = n 2 = μ 2 = ν 2 = 7 / 9, β 1 = 100, β 2 = 20。

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

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

[0235] Figure 3 It shows the simulation results of the angular rate loop tracking curve. Among them, ω z , ω y are the actual angular velocities obtained from the simulation, and ω zref , ω yref are the reference angular velocities obtained from the angle loop. It can be seen from the figure that both the pitch angular rate and the yaw angular rate can quickly track their corresponding virtual control commands, and the adjustment time is less than 0.5 s.

[0236] Figure 4 It shows the variation curves of the equivalent pitch rudder deflection angle and the equivalent yaw rudder deflection angle. It can be seen from the figure that the amplitude is always within the limit range of the maximum rudder deflection angle, and the equivalent input control command curve is smooth, meeting the engineering requirements.

[0237] Figure 5 It shows the simulation results of the tracking error curve. Among them, E1(1) is the error between the actual value and the reference value of the angle of attack, E2(1) is the error between the actual value and the reference value of the pitch angular rate, E1(2) is the error between the actual value and the reference value of the sideslip angle, and E2(2) is the error between the actual value and the reference value of the yaw angular rate. It can be seen from the figure that in the presence of external disturbances, the error can quickly converge.

[0238] Figure 6 、 7 shows the estimation effect of the state observer on the disturbances d 1 and d 2 , where d 1 (1) is the total disturbance component 1 of the angle loop, d 1 (1) esm is the estimated value of the total disturbance component 1 of the angle loop, d 1 (2) is the total disturbance component 2 of the angle loop, d 1 (2) esm is the estimated value of the total disturbance component 2 of the angle loop, d 2 (1) is the total disturbance component 1 of the angular rate loop, d 2 (1) esm is the estimated value of the total disturbance component 1 of the angular rate loop, d 2 (2) is the total disturbance component 2 of the angular rate loop, d 2 (2) esm is the estimated value of the total disturbance component 2 of the angular rate loop. It can be clearly seen from the figure that the reduced-order state observer can quickly and accurately estimate the disturbance value, improving the control accuracy of the closed-loop system.

[0239] In the description of the present invention, it should be noted that the orientation or positional relationship indicated by the terms "upper", "lower", "inner", "outer", "front", "rear", etc. is the orientation or positional relationship based on the working state of the present invention. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation to the present invention. In addition, the terms "first", "second", "third", "fourth" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance.

[0240] In the description of the present invention, it should be noted that unless otherwise clearly specified and defined, the terms "installed", "connected", "connected" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium, and it can be the communication inside two elements. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific situations.

[0241] The above has described the present invention in combination with preferred embodiments. However, these embodiments are only exemplary and only serve an illustrative role. On this basis, various substitutions and improvements can be made to the present invention, and these all fall within the protection scope of the present invention.

Claims

1. A design method for the attitude autopilot of a new type of guided vehicle, Characterized in that, It includes the following steps: Construct a model of the attitude control system of the guided vehicle; Based on the attitude control system model of a guided vehicle, an angle loop and an angular rate loop are constructed. The angle loop takes the reference angle as the input signal to obtain the virtual control rate The angular rate loop takes the virtual control rate as the input signal to obtain the attitude control command U; The guided vehicle changes its attitude according to the attitude control command U.

2. The roll vehicle attitude control method according to claim 1, Characterized in that, The attitude control system model of the guided vehicle is expressed as: Wherein, α represents the angle of attack, β represents the sideslip angle, θ represents the pitch angle, ω x represents the roll rate, ω y represents the yaw rate, ω z represents the pitch rate, δ y represents the yaw rudder deflection angle, δ z represents the pitch rudder deflection angle, d 1 is the total disturbance in the angle loop, d 2 is the total disturbance in the angular rate loop, ρ represents the air density, V represents the vehicle speed, S ref represents the characteristic area of the vehicle, m represents the mass of the vehicle, represents the partial derivative of the lift coefficient with respect to the angle of attack, represents the partial derivative of the lift coefficient with respect to the sideslip angle, represents the partial derivative of the pitch moment coefficient with respect to the angle of attack, represents the partial derivative of the yaw moment coefficient with respect to the sideslip angle, represents the partial derivative of the pitch damping moment coefficient with respect to the pitch rate, represents the partial derivative of the yaw damping moment coefficient with respect to the yaw rate, represents the partial derivative of the pitch moment coefficient with respect to the rudder deflection angle δ z of, represents the partial derivative of the pitch moment coefficient with respect to the rudder deflection angle δ y of, l represents the characteristic length of the vehicle, J x represents the moment of inertia about the x - direction, J y represents the moment of inertia about the y - direction, J z represents the moment of inertia about the z - direction.

3. The roll vehicle attitude control method according to claim 2, Characterized in that, The reference angle is set to where α ref is the reference angle of attack, and β ref is the reference sideslip angle, both provided by the guidance command; The angle loop includes the following sub-steps: S201. Based on the reference angle and the attitude control system of the guided vehicle, obtain the nominal angular rate, and the nominal angular rate is expressed as: S202. Set the angle loop control subsystem, expressed as: Among them, E 1 is the angle loop tracking error; S203. Set a state observer to estimate the total disturbance in the angle loop control subsystem; S204. Obtain the control rate of the angle loop control subsystem by using sliding mode control S205. Based on the nominal angular rate and the control rate of the angle loop control subsystem obtain the virtual control rate which is expressed as:

4. The roll vehicle attitude control method according to claim 3, Characterized in that, In S203, the state observer is a reduced-order state observer, preferably set as: Among them, represents the estimated value of d 1 , p 1 represents the auxiliary variable of the state observer, β 1 represents the observer gain of the state observer.

5. The roll vehicle attitude control method according to claim 3, Characterized in that, In S204, a sliding mode surface S is set 1 as follows: where τ represents the time constant, and E 1 (0) represents the initial state of E 1 , is the nominal value of.

6. The roll vehicle attitude control method according to claim 2, Characterized in that, The angular rate loop includes the following sub-steps: S301. Set the nominal control rate S302. Set the angle loop control subsystem, expressed as: where E 2 is the tracking error of the angular rate loop; S303. Set a state observer to estimate the total disturbance in the angle loop control subsystem; S304. Obtain the control rate of the angular rate loop control subsystem by using 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 command U, expressed as:

7. The roll vehicle attitude control method according to claim 6, Characterized in that, In S302, the angle loop tracking error E 1 is set to: Among them, Obtained by measuring with an aircraft sensor.

8. The roll vehicle attitude control method according to claim 6, Characterized in that, In S303, the state observer is a reduced-order state observer, preferably set as: wherein, represents the estimated value of d 2 , p 2 represents the auxiliary variable of the state observer, and β 2 represents the observer gain of the state observer.

9. The roll vehicle attitude control method according to claim 6, Characterized in that, In S304, a sliding mode surface S is set 2 It is: Among them, E 2 (0) represents the initial state of E 2 , is the nominal value of.

10. The roll vehicle attitude control method according to claim 6, Characterized in that, Control rate of the obtained angular rate loop control subsystem It is expressed as: Among them, η 21 , η 22 , η 23 are coefficient matrices, and v 1 , v 2 are constants.

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