A method of attitude control for a straight-air composite missile based on ESO and P gain

Abstract

The present invention discloses a method for attitude control of a direct-air composite missile based on electro-optical speed (ESO) and P-gain. The method comprises the following steps: establishing a missile body model and designing ignition logic; obtaining overload commands, using PI controllers for the normal / lateral overload loops, and simultaneously providing angle of attack / sideslip angle commands for the normal / lateral overload loops; obtaining the angle of attack / sideslip angle commands, and designing an ESO-based P-gain control law for the angle of attack / sideslip angle loop based on the relationship between the angle of attack / sideslip angle and angular velocity, thereby providing control commands for the angle of attack / sideslip angle velocity loop; designing an ESO-based P-gain control law for the angle of attack / sideslip angle velocity loop to achieve disturbance suppression and disturbance compensation in the impedance loop, and tracking the angle of attack / sideslip angle velocity commands to achieve attitude control. The method utilizes a hybrid control method, resulting in a fast response speed. Instead of modeling specific disturbances, the ESO is used to estimate disturbances, effectively covering most flight scenarios and offering high versatility.

Description

Technical Field

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

[0002] Direct force control utilizes the reaction force of the engine's jet to generate thrust, which acts on the controlled object to achieve control. This system features fast response and is unaffected by the atmospheric environment, making it suitable for flight environments such as near-space and the exoatmosphere. However, when direct force is activated, the reaction force of the nozzle interferes with the missile's trajectory. Existing direct force / aerodynamic hybrid missile attitude control methods exhibit poor response speed under direct force interference, limiting the missile's maneuverability. Therefore, ensuring missile stability under both environmental aerodynamic interference and direct force interference from the nozzle is a key issue that needs to be addressed.

[0003] Existing ESO methods primarily combine ADRC and sliding film control. ADRC's excessive number of parameters makes it difficult to apply universally, requiring extremely precise tuning for each scenario. Meanwhile, sliding film control, in order to achieve rapid sliding, introduces high-frequency modulation signals, which further exacerbates the vibration of the missile. Furthermore, both methods are complex in design. While they maintain a certain level of response speed, they are difficult to implement in engineering.

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

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

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

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

[0008] S100: Establish a missile body model and design the ignition logic. The ignition logic is to ignite the side jet engine when the aerodynamic overload is less than a certain range of the overload instruction.

[0009] S200: Obtaining a normal overload command for the pitch channel and a lateral overload command for the yaw channel. The normal overload loop of the pitch channel adopts a PI controller, and the normal overload loop provides an angle of attack command 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 command for the sideslip angle loop.

[0010] S300: Obtain the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel. Based on the relationship between the angle of attack and the angle of attack angular velocity, design an ESO-based P gain control law for the angle of attack loop to provide control commands for the angle of attack angular velocity loop. Based on the relationship between the sideslip angle and the sideslip angular velocity, design an ESO-based P gain control law for the sideslip angle loop to provide control commands 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 between the attitude control engine and 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, 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 specifically includes:

[0022] The overload ring 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 are the overloads of the pitch channel and 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, which is the unknown part, where u 11 =ω z is the control quantity of the angular velocity loop; g is the acceleration of gravity, 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 and needs to be set according to the task situation. The fal 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 the yaw channel is used, S300 is specifically as follows:

[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 sideslip angle loop control law 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 status 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. The P gain control law based on ESO is designed as follows:

[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 the yaw channel is used, S400 is specifically as follows:

[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 status y , z 42 Tracking external disturbances and model uncertainty 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] Where u 40 =(f7+z 42 ) / b3+u 41 ;

[0092] The P gain control based on ESO is designed 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 good response speed; the direct force model targeted by the scheme is discrete direct force, which is consistent with most nozzle types in actual 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 This is a flow chart of a method for attitude control of a direct-air composite missile based on ESO and P gain in one embodiment of the present invention;

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

[0100] Figure 3 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 with reference to the accompanying drawings.

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

[0104] S100: Establish a missile model and design the ignition logic. The ignition logic is such that when the aerodynamic overload is less than a certain range of the overload command, the side jet engine is ignited. Furthermore, when the aerodynamic overload does not reach the overload command, ignition is required to generate direct force to compensate for the difference between the overload command and the aerodynamic overload.

[0105] S200: Obtaining a normal overload command for the pitch channel and a lateral overload command for the yaw channel. The normal overload loop of the pitch channel adopts a PI controller, and the normal overload loop provides an angle of attack command 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 command for the sideslip angle loop.

[0106] S300: Obtain the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel. Based on the relationship between the angle of attack and the angle of attack angular velocity, design an ESO-based P gain control law for the angle of attack loop to provide control commands for the angle of attack angular velocity loop. Based on the relationship between the sideslip angle and the sideslip angular velocity, design an ESO-based P gain control law for the sideslip angle loop to provide control commands 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 in turn will cause overload; attitude control is actually to control the angle of attack 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 ESO-based single-gain composite control system is Figure 2 As shown, the pitch channel (upper half) and the yaw channel (lower half) have similar structures and are completely symmetrical on both sides. For the pitch channel and 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 the state variable, the yaw channel selects ω y , β, and n z As state variables, three loops are designed separately. The inner loop (angular velocity loop) and the middle loop (angle of attack and sideslip angle loop) utilize ESO design, primarily due to the large perturbation torque of the inner loop, which is difficult to accurately model and control. The perturbation 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. The outer loop (overload loop) utilizes PI design, drawing on the design method of asymptotic tracking for linear systems. When the system output is a combination of states, a PI controller is required to achieve zero steady-state error. Therefore, the relationship between overload and angle of attack and sideslip angle is expressed through PI. P is used to implement the proportionality coefficient between overload and angle of attack, and I is used to eliminate the steady-state error caused by overload. The inner loop is required to have good control and a response speed at least an order of magnitude faster than that of the middle loop. When designing the middle loop controller, the inner loop can be treated as a straight-through path. This achieves the design of a combined control strategy based on ADRC and PI.

[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 good response speed; the direct force model targeted by the scheme is discrete direct force, which is consistent with most nozzle types in actual 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 for 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 between the attitude control engine and 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, 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 the form of a trapezoidal pulse. 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's attitude control engines are distributed in different positions, have fixed working hours and thrust, and cannot be turned off once turned on, the attitude control engine distribution method is crucial. The rationality of the distribution 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 torque generated by the attitude control engine 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 the attitude control engine is shown.

[0123] In one embodiment, S200 specifically includes:

[0124] The overload ring 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 reflect 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 are the overloads of the pitch channel and 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 the channel is a pitch channel, S300 specifically includes:

[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, which is the unknown part, where u 11 =ω z is the control quantity of the angular velocity loop; g is the acceleration of gravity, c y is the lift coefficient, s is the characteristic area, deltaz is the 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 and needs to be set according to the task situation. The fal 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 sideslip angle loop control law 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 loop design is the basis of lateral overload loop design and should have faster response speed and better robustness.

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

[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 statusz , 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. The P gain control law based on ESO is designed as follows:

[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 as follows:

[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 status y , z 42 Tracking external disturbances and model uncertainty 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] Where u 40 =(f7+z 42 ) / b3+u 41 ;

[0196] The P gain control based on ESO is designed 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 lateral overload loop design, and should have a very fast response speed and good robustness.

[0202] The aforementioned attitude control method for a direct-air hybrid missile based on ESO and P-gain enables the missile to maintain stable attitude under various disturbances. The method's design incorporates discrete direct forces, making it more practical. It also balances the response speed of ESO with the engineering capabilities of P-gain, making it suitable for deployment on small cruise missiles, thereby advancing missile technology.

[0203] The above describes in detail the attitude control method for a direct-air composite missile 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 above examples are intended only to facilitate understanding of the core concepts of the present invention. It should be noted that those skilled in the art will be able to make various improvements and modifications to the present invention without departing from the principles of the present invention, and such improvements and modifications are also within the scope of protection of the claims.

Claims

1. A method for attitude control of a direct-air composite missile based on ESO and P gain, characterized in that: The method comprises the following steps: S100: Establish a missile body model and design the ignition logic. The ignition logic is to ignite the side jet engine when the aerodynamic overload is less than a certain range of the overload instruction. S200: Obtaining a normal overload command for the pitch channel and a lateral overload command for the yaw channel. The normal overload loop of the pitch channel adopts a PI controller, and the normal overload loop provides an angle of attack command 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 command for the sideslip angle loop. S300: Obtain the angle of attack command of the pitch channel and the sideslip angle command of the yaw channel. Based on the relationship between the angle of attack and the angle of attack angular velocity, design an ESO-based P gain control law for the angle of attack loop to provide control commands for the angle of attack angular velocity loop. Based on the relationship between the sideslip angle and the sideslip angular velocity, design an ESO-based P gain control law for the sideslip angle loop to provide control commands 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 between the attitude control engine and 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, 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 trapezoidal pulses, using 48 side-jet engines, 12 in each circle, for a total of 4 circles, evenly distributed in the two directions of the pitch channel and the 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, wherein S200 is specifically: The overload ring 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 are the overloads of the pitch channel and yaw channel respectively.

5. The method according to claim 4, characterized in that When it is the 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, which is the unknown part, where u 11 =ω z is the control quantity of the angular velocity loop; g is the acceleration of gravity, 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 and needs to be set according to the task situation. The fal 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 attack angle 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)=β4e 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 sideslip angle loop control law is: u 21 (t)=u 20 (t)-z 22 -f3。 7. The method according to claim 6, characterized in that When it is the 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 status 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. The P gain control law based on ESO is designed as follows: 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 status y , z 42 Tracking external disturbances and model uncertainty 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 ; The P gain control based on ESO is designed 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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