ESO and P gain-based direct-gas composite missile attitude control method

Abstract

The invention discloses a direct-gas composite missile attitude control method based on ESO and P gain, and the method comprises the steps: building a missile body model, and designing an ignition logic; an overload instruction is obtained, the normal / lateral overload ring adopts a PI controller, and meanwhile, the normal / lateral overload ring provides an attack angle / sideslip angle instruction for the attack angle / sideslip angle ring; acquiring an attack angle / sideslip angle instruction, designing a P gain control law based on ESO for an attack angle / sideslip angle ring based on the relationship between the attack angle / sideslip angle and the angular velocity, and providing a control instruction for an attack angle / sideslip angle velocity ring; a P gain control law based on ESO is designed for an attack angle / sideslip angle velocity loop, disturbance suppression and disturbance compensation of an impedance loop are achieved, an attack angle / sideslip angle velocity instruction is tracked, and attitude control is achieved. A hybrid control method is adopted, and the response speed is high; modeling is not carried out on specific disturbance, the disturbance is estimated through ESO, most of flight scenes can be effectively covered, and the universality is high.

Description

Technical Field

[0001] The invention belongs to the technical field of missile attitude control, and in particular relates to an attitude control method of a direct-air composite missile based on ESO and P gain. Background Art

[0002] Direct force control uses the reaction force of the engine jet to generate thrust, which acts on the controlled object to achieve the control task. It has the characteristics of fast response speed and is not affected by the atmospheric environment. It is suitable for flight environments such as near space and outside the atmosphere. However, when the direct force is turned on, the reaction force of the nozzle will interfere with the missile body. The existing direct force / aerodynamic composite missile attitude control method has poor response speed under the condition of direct force interference, which limits the maneuverability of the missile. Therefore, how to ensure the stability of the missile body under the aerodynamic interference of the environment and the direct force interference of the nozzle is a key problem that needs to be solved.

[0003] The existing ESO method is mainly combined with ADRC and sliding film control. ADRC has too many parameters, which makes it difficult to use in general, and requires extremely precise parameter adjustment for each scenario; while sliding film control introduces high-frequency adjustment signals to achieve fast sliding, which further aggravates the vibration of the missile. At the same time, the design of these two methods is relatively complex. Although they guarantee a certain degree of response speed, they are difficult to engineer.

[0004] Based on this, a direct-air composite missile attitude control method based on ESO and P gain is proposed in the present invention. Summary of the invention

[0005] In view of the above technical problems, the present invention provides a kind of.

[0006] The technical solution adopted by the present invention to solve the technical problem is:

[0007] A method for controlling a straight-air composite missile attitude based on ESO and P gain, the method comprising the following steps:

[0008] S100: Establish a missile body model and design an ignition logic; wherein the ignition logic is that when the aerodynamic overload is less than a certain range of the overload instruction, the side jet engine is ignited and turned on;

[0009] S200: obtaining a normal overload instruction of the pitch channel and a lateral overload instruction of the yaw channel, wherein the normal overload loop of the pitch channel adopts a PI controller, and the normal overload loop provides an angle of attack instruction for the angle of attack loop; the lateral overload loop of the yaw channel adopts a PI controller, and the lateral overload loop provides a sideslip angle instruction for the sideslip angle loop;

[0010] S300: Obtaining the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel, designing a P gain control law based on ESO for the angle of attack loop based on the relationship between the angle of attack and the angle of attack angular velocity, and providing a control command for the angle of attack angular velocity loop; designing a P gain control law based on ESO for the sideslip angle loop based on the relationship between the sideslip angle and the sideslip angular velocity, and providing a control command for the sideslip angular velocity loop;

[0011] S400: The design of the angle of attack velocity loop and the sideslip velocity loop is based on the P gain control law of ESO to achieve disturbance suppression and disturbance compensation of the impedance loop, and track the angle of attack velocity command and the sideslip velocity command to achieve attitude control.

[0012] Preferably, S100 specifically includes:

[0013] The relationship between the thrust of the attitude engine and time is:

[0014]

[0015] Among them, F a is the thrust of the attitude control engine, T is the working cycle of the attitude control engine, F max is the thrust of the attitude control engine in steady state, τ is the opening and closing time of the attitude control engine;

[0016] The direct forces and moments generated by the attitude motor are:

[0017]

[0018]

[0019] Among them, F x1 、F y1 、F z1 O generated by the attitude engine x1 , O y1 With O z1 Direct force on the shaft, M x1 、M y1 、M z1 O generated by the attitude engine x1 , O y1 With O z1 The moment of the shaft, l a is the average distance of the attitude control engine relative to the missile's center of mass; N y 、N z Respectively represent O y1 With O z1 The number of attitude control engines turned on in the axial direction is determined, and the attitude control engine partitions that need to be turned on are distinguished based on the positive and negative signs.

[0020] Preferably, the attitude engine is a solid rocket engine in the form of a trapezoidal pulse, using 48 side-jet engines, 12 in each circle, for a total of 4 circles, evenly distributed in two directions of the pitch channel and two directions of the yaw channel, forming a negative pitch control area, a positive yaw control area, a positive pitch control area and a negative yaw control area respectively.

[0021] Preferably, S200 is specifically:

[0022] The overload loop of the pitch channel is:

[0023] α c =k p (n yc -n y )+k i ∫(n yc -n y )dt

[0024] The overload loop of the yaw channel is:

[0025] β c =k p (n zc -n z )+k i ∫(n zc -n z )dt

[0026] Among them, α c and β c are the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel, k p Used to reflect α and n y and β and n z The coefficient relationship, k i Used to eliminate steady-state error, n yc and n zc is the normal overload command of the pitch channel and the lateral overload command of the yaw channel, n y and n z They are the overload of the pitch channel and the yaw channel respectively.

[0027] Preferably, when it is a pitch channel, S300 is specifically:

[0028] The model of the angle of attack loop of the pitch channel is:

[0029]

[0030] in,

[0031] The model of the pitch channel angle of attack loop can be rewritten as:

[0032]

[0033] in, is the known part of the model, is the external disturbance and model uncertainty, is the unknown part, where u 11 =ω z is the control quantity of the angular velocity loop; g is the gravitational acceleration, c y is the lift coefficient, s is the characteristic area, δ z is the lateral rudder angle, ρ is the atmospheric density, V is the missile speed;

[0034] For the above formula, the designed ESO is:

[0035]

[0036] Among them, z 11 Tracking system state α, z 12 Tracking external disturbances and model uncertainty f2, β 01 , β 02 It is an ADRC parameter, which needs to be set according to the task situation. The fall function is defined as follows:

[0037]

[0038] The above ESO can be written as:

[0039]

[0040] Among them, u 10 =f1+z 12 +u 11 ;

[0041] For the above formula, the P gain control law based on ESO is designed as:

[0042]

[0043] Among them, α c is the input command of the angle of attack loop, which is provided by the control output of the overload loop. The final control law of the angle of attack loop is:

[0044] u 11 (t) = u 10 (t)-z 12 -f1.

[0045] Preferably, when it is a yaw channel, S300 is specifically:

[0046] The model of the sideslip angle loop β of the yaw channel is:

[0047]

[0048] in,

[0049] The model of the sideslip angle loop β of the yaw channel is rewritten as:

[0050]

[0051] in, is the known part of the model, is the external disturbance and model uncertainty, which is the unknown part, b1=cosα, u 21 =ω y is the control quantity of the sideslip angle loop;

[0052] For the above formula, the designed ESO is:

[0053]

[0054] Among them, z 21 Tracking system state β,z 22 Tracking external disturbances and model uncertainty f4, β 03 , β 04 It is an ADRC parameter and needs to be set according to the task situation;

[0055] The above ESO can be written as:

[0056]

[0057] Among them, u 21 =f3+z 22 +b1u 21 ;

[0058] For the above formula, the P gain control law based on ESO is designed as:

[0059] e2=β c -z 21

[0060] u4(t)=β4e 41

[0061] Among them, β c is the input command of the sideslip angle loop, which is provided by the control output of the lateral overload loop. The final control law of the sideslip angle loop is:

[0062] u 21 (t) = u 20 (t)-z 22 -f3.

[0063] Preferably, when it is a pitch channel, S400 is specifically:

[0064] Angular velocity loop ω of the pitch channel z The model is:

[0065]

[0066] Rewrite the above formula as:

[0067]

[0068] in, f6=M dz is the model uncertainty and external disturbance, which is the unknown part, and u 31 =δ z ;M z is the component of the moment of all external forces acting on the missile on the center of mass on each axis of the missile body coordinate system, J z is the moment of inertia, a1, a2, a3 are coefficients;

[0069] For the above formula, design the following ESO:

[0070]

[0071] Among them, z 31 Tracking system statusω z , z 32 Tracking external disturbances and model uncertainty f6, β 05 , β 06 It is an ADRC parameter and needs to be set according to the task situation;

[0072] The above ESO can be written as:

[0073]

[0074] In the formula, u3=(f5+z 32 ) / b2+u 31 , where b2 is the nominal value of b;

[0075] The above equation is a linear system, and the P gain control law based on ESO is designed as:

[0076] e3=ω zc -z 31

[0077] u3(t)=β3e3

[0078] Among them, ω zc is the pitch angular velocity loop input command, which is provided by the attack angle loop control output, so the final angular velocity loop control law is:

[0079] u 30 (t) = u3(t)-[z 32 +f3(α,ω z )] / b2.

[0080] Preferably, when it is a yaw channel, S400 is specifically:

[0081] Angular velocity loop of the yaw channel The model is:

[0082]

[0083] Rewrite the above formula as:

[0084]

[0085] in, is the known part of the model, f8=M dy is the model uncertainty and external disturbance, which is the unknown part, u 41 =δ y ;M y is the component of the moment of all external forces acting on the missile on the center of mass on each axis of the missile body coordinate system, J y is the moment of inertia;

[0086] Design ESO to:

[0087]

[0088] Among them, z 41 Tracking system statusω y , z 42 Tracking external disturbances and model uncertainties f8, β 07 , β 08 It is an ADRC parameter and needs to be set according to the task situation;

[0089] The above ESO can be written as:

[0090]

[0091] In the formula, u 40 =(f7+z 42 ) / b3+u 41 ;

[0092] Design the ESO-based P gain control as:

[0093] e4=ω yc -z 41

[0094] u4(t)=β4e4

[0095] Among them, ω yc is the input command of the yaw rate loop, which is provided by the sideslip angle loop control output. The final yaw rate loop control law is:

[0096] u4(t)=u 40 -[z 42 +f7(β,ω y )] / b3.

[0097] The above-mentioned direct-air composite missile attitude control method based on ESO and P gain adopts a hybrid control method of ESO and P gain, which has high engineering value and excellent response speed; the direct force model targeted by the scheme is discrete direct force, which is consistent with most nozzle types in practical applications; the scheme does not model specific disturbances, but uses ESO to estimate disturbances, so it can effectively cover most flight scenarios, has high versatility, and is relatively convenient to deploy on small cruise missiles. BRIEF DESCRIPTION OF THE DRAWINGS

[0098] Figure 1 It is a flow chart of a method for controlling a straight-air composite missile attitude based on ESO and P gain in one embodiment of the present invention;

[0099] Figure 2 A structural diagram of a single-gain composite control system based on ESO in one embodiment of the present invention;

[0100] Figure 3 A schematic diagram of a thrust curve of an attitude control engine in one embodiment of the present invention;

[0101] Figure 4 This is a simplified circumferential expansion diagram of the attitude control engine group in one embodiment of the present invention. DETAILED DESCRIPTION

[0102] In order to enable those skilled in the art to better understand the technical solution of the present invention, the present invention is further described in detail below in conjunction with the accompanying drawings.

[0103] In one embodiment, Figure 1 As shown, a straight-air composite missile attitude control method based on ESO and P gain, the method comprising the following steps:

[0104] S100: Establish a missile body model and design an ignition logic; wherein, the ignition logic is that when the aerodynamic overload is less than a certain range of the overload instruction, the side jet engine is ignited and turned on; further, when the aerodynamic overload does not reach the overload instruction, it is necessary to ignite to generate direct force to make up for the difference between the overload instruction and the aerodynamic overload;

[0105] S200: obtaining a normal overload instruction of the pitch channel and a lateral overload instruction of the yaw channel, wherein the normal overload loop of the pitch channel adopts a PI controller, and the normal overload loop provides an angle of attack instruction for the angle of attack loop; the lateral overload loop of the yaw channel adopts a PI controller, and the lateral overload loop provides a sideslip angle instruction for the sideslip angle loop;

[0106] S300: Obtaining the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel, designing a P gain control law based on ESO for the angle of attack loop based on the relationship between the angle of attack and the angle of attack angular velocity, and providing a control command for the angle of attack angular velocity loop; designing a P gain control law based on ESO for the sideslip angle loop based on the relationship between the sideslip angle and the sideslip angular velocity, and providing a control command for the sideslip angular velocity loop;

[0107] S400: The design of the angle of attack velocity loop and the sideslip velocity loop is based on the P gain control law of ESO to achieve disturbance suppression and disturbance compensation of the impedance loop, and track the angle of attack velocity command and the sideslip velocity command to achieve attitude control.

[0108] Specifically, angular velocity is directly related to the adjustment of the projectile's attitude, because changes in angular velocity will cause changes in the angle of attack and sideslip angle, which will in turn cause overload; attitude control is actually to control the attack angle alpha, sideslip angle beta and roll angle gamma. Here, the roll channel is assumed to be non-deflected, that is, the roll angle is 0, and only the other two channels are considered.

[0109] Furthermore, the single-gain composite control system based on ESO is Figure 2 As shown, the pitch channel (upper half) and the yaw channel (lower half) have similar structures, and the two sides are completely symmetrical. For the pitch channel and the yaw channel, according to the characteristics of the model and the time scale analysis results, the pitch channel in the present invention selects ω z , α and n y As a state variable, the yaw channel selects ω y , β and n z As state variables, three loops are designed respectively. The inner loop (angular velocity loop) and the middle loop (angle of attack and sideslip angle loop) are designed with ESO, mainly considering that the inner loop has a large disturbance torque, which is difficult to accurately model and control. The disturbance torque of the inner loop directly affects the output of the middle loop, and the middle loop itself also has some unmodeled dynamics, such as the uncertainty of the direct lateral force itself. The outer loop (overload loop) is designed with PI, considering the design method of linear system progressive tracking. When the system output is a combination of states, a PI controller needs to be added to make the steady-state error zero, so the relationship between overload and angle of attack and sideslip angle is manifested through PI. P is used to realize the proportional coefficient between overload and angle of attack, and I is used to eliminate the steady-state error of overload. Here, the inner loop is required to have good control and a response speed that is more than one order of magnitude faster than the middle loop. When designing the middle loop controller, the inner loop can be regarded as a direct path. In this way, the design of a combined control strategy based on ADRC and PI is realized.

[0110] The above-mentioned direct-air composite missile attitude control method based on ESO and P gain adopts a hybrid control method of ESO and P gain, which has high engineering value and excellent response speed; the direct force model targeted by the scheme is discrete direct force, which is consistent with most nozzle types in practical applications; the scheme does not model specific disturbances, but uses ESO to estimate disturbances, so it can effectively cover most flight scenarios, has high versatility, and is relatively convenient to deploy on small cruise missiles.

[0111] In one embodiment, S100 specifically includes:

[0112] The relationship between the thrust of the attitude engine and time is:

[0113]

[0114] Among them, F a is the thrust of the attitude control engine, T is the working cycle of the attitude control engine, F max is the thrust of the attitude control engine in steady state, τ is the opening and closing time of the attitude control engine;

[0115] The direct forces and moments generated by the attitude motor are:

[0116]

[0117]

[0118] Among them, F x1 、F y1 、F z1 O generated by the attitude engine x1 , O y1 With O z1 Direct force on the shaft, M x1 、M y1 、M z1 O generated by the attitude engine x1 , O y1 With O z1 The moment of the shaft, l a is the average distance of the attitude control engine relative to the missile's center of mass; N y 、N z Respectively represent O y1 With O z1 The number of attitude control engines turned on in the axial direction is determined, and the attitude control engine partitions that need to be turned on are distinguished based on the positive and negative signs.

[0119] In one embodiment, the attitude engine is a solid rocket engine in the form of a trapezoidal pulse, using 48 side-jet engines, 12 in each circle, for a total of 4 circles, evenly distributed in two directions of the pitch channel and two directions of the yaw channel, forming a negative pitch control area, a positive yaw control area, a positive pitch control area and a negative yaw control area respectively.

[0120] Specifically, the attitude control engine is a solid rocket engine in trapezoidal pulse form. Ignoring its ignition delay and considering the power on / off delay, the thrust curve of the attitude control engine is as follows: Figure 3 shown.

[0121] Furthermore, considering that the missile attitude control engines are distributed in different positions, have fixed working time and thrust, and cannot be turned off once turned on, the attitude control engine allocation method is crucial. The rationality of the allocation method design will directly determine the utilization rate of the attitude control engine and the performance of the controller. The system uses 48 side-jet engines, 12 in each circle, for a total of 4 circles, evenly distributed in four directions.

[0122] In order to facilitate the design of the controller, the attitude control engine is equivalent to the corresponding missile body axis. At the same time, it is assumed that the nominal thrust generated by attitude control engines of different circle numbers at different times is the same relative to the center of mass. The attitude control engine is simplified as follows Figure 4 The simplified distribution model of attitude control engines is shown.

[0123] In one embodiment, S200 specifically includes:

[0124] The overload loop of the pitch channel is:

[0125] α c =k p (n yc -n y )+k i ∫(n yc -n y )dt

[0126] The overload loop of the yaw channel is:

[0127] β c =k p (n zc -n z )+k i ∫(n zc -n z )dt

[0128] Among them, α c and β c are the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel, k p Used to react α and n y and β and nz The coefficient relationship, k i Used to eliminate steady-state error, n yc and n zc is the normal overload command of the pitch channel and the lateral overload command of the yaw channel, n y and n z They are the overload of the pitch channel and the yaw channel respectively.

[0129] Specifically, the design goals of the normal overload loop and the lateral overload loop are to track the overload instruction as quickly as possible and eliminate the overload tracking deviation.

[0130] In one embodiment, when it is a pitch channel, S300 is specifically:

[0131] The model of the angle of attack loop of the pitch channel is:

[0132]

[0133] in,

[0134] The model of the pitch channel angle of attack loop can be rewritten as:

[0135]

[0136] in, is the known part of the model, is the external disturbance and model uncertainty, is the unknown part, where u 11 =ω z is the control quantity of the angular velocity loop; g is the gravitational acceleration, c y is the lift coefficient, s is the characteristic area, deltaz is the lateral rudder deflection angle, ρ is the atmospheric density, V is the missile speed;

[0137] For the above formula, the designed ESO is:

[0138]

[0139] Among them, z 11 Tracking system state α, z 12 Tracking external disturbances and model uncertainty f2, β 01 , β 02 It is an ADRC parameter, which needs to be set according to the task situation. The fall function is defined as follows:

[0140]

[0141] The above ESO can be written as:

[0142]

[0143] Among them, u 10 =f1+z 12 +u 11 ;

[0144] For the above formula, the P gain control law based on ESO is designed as:

[0145]

[0146] Among them, α c is the input command of the angle of attack loop, which is provided by the control output of the overload loop. The final control law of the angle of attack loop is:

[0147] u 11 (t) = u 10 (t)-z 12 -f1.

[0148] In one embodiment, when the yaw channel is used, S300 is specifically as follows:

[0149] The model of the sideslip angle loop β of the yaw channel is:

[0150]

[0151] in,

[0152] The model of the sideslip angle loop β of the yaw channel is rewritten as:

[0153]

[0154] in, is the known part of the model, is the external disturbance and model uncertainty, which is the unknown part, b1=cosα, u 21 =ω y is the control quantity of the sideslip angle loop;

[0155] For the above formula, the designed ESO is:

[0156]

[0157] Among them, z 21 Tracking system state β,z 22 Tracking external disturbances and model uncertainty f4, β 03 , β 04 It is an ADRC parameter and needs to be set according to the task situation;

[0158] The above ESO can be written as:

[0159]

[0160] Among them, u 21=f3+z 22 +b1u 21 ;

[0161] For the above formula, the P gain control law based on ESO is designed as:

[0162] e2=β c -z 21

[0163] u4(t)=β4e 41

[0164] Among them, β c is the input command of the sideslip angle loop, which is provided by the control output of the lateral overload loop. The final control law of the sideslip angle loop is:

[0165] u 21 (t) = u 20 (t)-z 22 -f3.

[0166] The above realizes the design of sideslip angle loop and angle of attack loop of pitch channel and yaw channel based on ADRC. The sideslip ring design is the basis of the lateral overload ring design and should have faster response speed and better robustness.

[0167] In one embodiment, when it is a pitch channel, S400 is specifically:

[0168] Angular velocity loop ω of the pitch channel z The model is:

[0169]

[0170] Rewrite the above formula as:

[0171]

[0172] in, f6=M dz is the model uncertainty and external disturbance, which is the unknown part, and u 31 =δ z ;M z is the component of the moment of all external forces acting on the missile on the center of mass on each axis of the missile body coordinate system, J z is the moment of inertia, a1, a2, a3 are coefficients;

[0173] For the above formula, design the following ESO:

[0174]

[0175] Among them, z 31 Tracking system statusωz , z 32 Tracking external disturbances and model uncertainty f6, β 05 , β 06 It is an ADRC parameter and needs to be set according to the task situation;

[0176] The above ESO can be written as:

[0177]

[0178] In the formula, u3=(f5+z 32 ) / b2+u 31 , where b2 is the nominal value of b;

[0179] The above equation is a linear system, and the P gain control law based on ESO is designed as:

[0180] e3=ψ zc -z 31

[0181] u3(t)=β3e3

[0182] Among them, ω zc is the pitch angular velocity loop input command, which is provided by the attack angle loop control output, so the final angular velocity loop control law is:

[0183] u 30 (t) = u3(t)-[z 32 +f3(α,ω z )] / b2.

[0184] In one embodiment, when the yaw channel is used, S400 is specifically:

[0185] Angular velocity loop of the yaw channel The model is:

[0186]

[0187] The above formula can be rewritten as:

[0188]

[0189] in, is the known part of the model, f8=M dy is the model uncertainty and external disturbance, which is the unknown part, u 41 =δ y ;M y is the component of the moment of all external forces acting on the missile on the center of mass on each axis of the missile body coordinate system, J y is the moment of inertia;

[0190] Design ESO to:

[0191]

[0192] Among them, z 41 Tracking system statusω y , z 42 Tracking external disturbances and model uncertainties f8, β 07 , β 08 It is an ADRC parameter and needs to be set according to the task situation;

[0193] The above ESO can be written as:

[0194]

[0195] In the formula, u 40 =(f7+z 42 ) / b3+u 41 ;

[0196] Design the ESO-based P gain control as:

[0197] e4=ω yc -z 41

[0198] u4(t)=β4e4

[0199] Among them, ω yc is the input command of the yaw rate loop, which is provided by the sideslip angle loop control output. The final yaw rate loop control law is:

[0200] u4(t)=u 40 -[z 42 +f7(β,ω y )] / b3.

[0201] The above realizes the design of the inner loop based on ADRC, namely the angular velocity loop. The yaw angular velocity loop design is the basis for the sideslip angle loop and the lateral overload loop design, and should have a very fast response speed and good robustness.

[0202] The above-mentioned direct-air composite missile attitude control method based on ESO and P gain can stabilize the attitude of the missile body under various disturbances. The method takes into account the discrete direct force in the design, which is closer to practical application; the method takes into account the response speed of ESO and the engineering capability of P gain, so it is more convenient to deploy on small cruise missiles to improve missile technology.

[0203] The above is a detailed introduction to the straight-air composite missile attitude control method based on ESO and P gain provided by the present invention. This article uses specific examples to illustrate the principles and implementation methods of the present invention. The description of the above embodiments is only used to help understand the core idea of ​​the present invention. It should be pointed out that for ordinary technicians in this technical field, without departing from the principle of the present invention, the present invention can also be improved and modified, and these improvements and modifications also fall within the scope of protection of the claims of the present invention.

Claims

1. A straight-air composite missile attitude control method based on ESO and P gain, characterized in that: The method comprises the following steps: S100: Establish a missile body model and design an ignition logic; wherein the ignition logic is that when the aerodynamic overload is less than a certain range of the overload instruction, the side jet engine is ignited and turned on; S200: obtaining a normal overload instruction of the pitch channel and a lateral overload instruction of the yaw channel, wherein the normal overload loop of the pitch channel adopts a PI controller, and the normal overload loop provides an angle of attack instruction for the angle of attack loop; the lateral overload loop of the yaw channel adopts a PI controller, and the lateral overload loop provides a sideslip angle instruction for the sideslip angle loop; S300: Obtaining the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel, designing a P gain control law based on ESO for the angle of attack loop based on the relationship between the angle of attack and the angle of attack angular velocity, and providing a control command for the angle of attack angular velocity loop; designing a P gain control law based on ESO for the sideslip angle loop based on the relationship between the sideslip angle and the sideslip angular velocity, and providing a control command for the sideslip angular velocity loop; S400: The design of the angle of attack velocity loop and the sideslip velocity loop is based on the P gain control law of ESO to achieve disturbance suppression and disturbance compensation of the impedance loop, and track the angle of attack velocity command and the sideslip velocity command to achieve attitude control.

2. The method according to claim 1, characterized in that S100 is specifically: The relationship between the thrust of the attitude engine and time is: Among them, F a is the thrust of the attitude control engine, T is the working cycle of the attitude control engine, F max is the thrust of the attitude control engine in steady state, τ is the opening and closing time of the attitude control engine; The direct forces and moments generated by the attitude motor are: Among them, F x1 、F y1 、F z1 O generated by the attitude engine x1 , O y1 With O z1 Direct force on the shaft, M x1 、M y1 、M z1 O generated by the attitude engine x1 , O y1 With O z1 The moment of the shaft, l a is the average distance of the attitude control engine relative to the missile's center of mass; N y 、N z Respectively represent O y1 With O z1 The number of attitude control engines turned on in the axial direction is determined, and the attitude control engine partitions that need to be turned on are distinguished based on the positive and negative signs.

3. The method according to claim 2, characterized in that The attitude engine is a solid rocket engine in the form of a trapezoidal pulse, using 48 side-jet engines, 12 in each circle, for a total of 4 circles, evenly distributed in two directions of the pitch channel and two directions of the yaw channel, forming a negative pitch control area, a positive yaw control area, a positive pitch control area and a negative yaw control area respectively.

4. The method according to claim 1, characterized in that: S200 is specifically: The overload loop of the pitch channel is: α c =k p (n yc -n y )+k i ∫(n yc -n y )dt The overload loop of the yaw channel is: β c =k p (n zc -n z )+k i ∫(n zc -n z )dt Among them, α c and β c are the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel, k p Used to reflect α and n y and β and n z The coefficient relationship, k i Used to eliminate steady-state error, n yc and n zc is the normal overload command of the pitch channel and the lateral overload command of the yaw channel, n y and n z They are the overload of the pitch channel and the yaw channel respectively.

5. The method according to claim 4, characterized in that When it is a pitch channel, S300 is specifically: The model of the pitch channel angle of attack loop is: in, The model of the pitch channel angle of attack loop can be rewritten as: in, is the known part of the model, is the external disturbance and model uncertainty, is the unknown part, where u 11 =ω z is the control quantity of the angular velocity loop; g is the gravitational acceleration, c y is the lift coefficient, s is the characteristic area, δ z is the lateral rudder angle, ρ is the atmospheric density, V is the missile speed; For the above formula, the designed ESO is: Among them, z 11 Tracking system state α, z 12 Tracking external disturbances and model uncertainty f2, β 01 , β 02 It is an ADRC parameter, which needs to be set according to the task situation. The fall function is defined as follows: The above ESO can be written as: Among them, u 10 = f1 + z 12 +u 11 ; For the above formula, the P gain control law based on ESO is designed as: Among them, α c is the input command of the angle of attack loop, which is provided by the control output of the overload loop. The final control law of the angle of attack loop is: u 11 (t)=u 10 (t)-z 12 -f1。 6. The method according to claim 5, characterized in that When it is the yaw channel, S300 is specifically: The model of the sideslip angle loop β of the yaw channel is: in, The model of the sideslip angle loop β of the yaw channel is rewritten as: in, is the known part of the model, is the external disturbance and model uncertainty, which is the unknown part, b1=cosα, u 21 =ω y is the control quantity of the angle of attack loop; For the above formula, the designed ESO is: Among them, z 21 Tracking system state β,z 22 Tracking external disturbances and model uncertainty f4, β 03 , β 04 It is an ADRC parameter and needs to be set according to the task situation; The above ESO can be written as: Among them, u 21 =f3+z 22 +b1u 21 ; For the above formula, the P gain control law based on ESO is designed as: e2=β c -With 21 u4(t)β=4e 41 Among them, β c is the input command of the sideslip angle loop, which is provided by the control output of the lateral overload loop. The final control law of the sideslip angle loop is: u 21 (t)=u 20 (t)-z 22 -f3。 7. The method according to claim 6, characterized in that When it is a pitch channel, S400 is specifically: Angular velocity loop ω of the pitch channel z The model is: Rewrite the above formula as: in, f6=M dz is the model uncertainty and external disturbance, which is the unknown part, and u 31 =δ z ;M z is the component of the moment of all external forces acting on the missile on the center of mass on each axis of the missile body coordinate system, J z is the moment of inertia, a1, a2, a3 are the aerodynamic coefficients; For the above formula, design the following ESO: Among them, z 31 Tracking system statusω z , z 32 Tracking external disturbances and model uncertainty f6, β 05 , β 06 It is an ADRC parameter and needs to be set according to the task situation; The above ESO can be written as: In the formula, u3=(f5+z 32 ) / b2+u 31 , where b2 is the nominal value of b; The above equation is a linear system, and the P gain control law based on ESO is designed as: e3=ω zc -z 31 u3(t)=β3e3 Among them, ω zc is the pitch angular velocity loop input command, which is provided by the attack angle loop control output, so the final angular velocity loop control law is: u 30 (t)=u3(t)-[z 32 +f3(α,ω z )] / b2。 8. The method according to claim 7, characterized in that When it is the yaw channel, S400 is specifically: Angular velocity loop of the yaw channel The model is: Rewrite the above formula as: in, is the known part of the model, f8=M dy is the model uncertainty and external disturbance, which is the unknown part, u 41 =δ y ;M y is the component of the moment of all external forces acting on the missile on the center of mass on each axis of the missile body coordinate system, J y is the moment of inertia; Design ESO to: Among them, z 41 Tracking system statusω y , z 42 Tracking external disturbances and model uncertainties f8, β 07 , β 08 It is an ADRC parameter and needs to be set according to the task situation; The above ESO can be written as: In the formula, u 40 = (f7+z 42 ) / b3+u 41 ; Design the ESO-based P gain control as: e4=ω yc -z 41 u4(t)=β4e4 Among them, ω yc is the input command of the yaw rate loop, which is provided by the sideslip angle loop control output. The final yaw rate loop control law is: u4(t)=u 40 -[z 42 +f7(β,ω y )] / b3.

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