A method for agile turning control of a missile

Abstract

The application provides a novel missile agile turning control method, a deceleration parachute is introduced on the basis of a missile to provide flexible resistance for the missile; the missile turns to a target direction under the action of multiple source forces such as aerodynamic force provided by a parachute, a lateral RCS and an aerodynamic rudder and main engine thrust, the action of the deceleration parachute causes the speed of the missile to decrease, the change rate of the attack angle of the missile is significantly increased, the rapid increase of the trajectory inclination angle of the missile is embodied in the agile turning process, so that the turning radius is smaller than that of a traditional agile missile, and therefore the maneuverability is stronger. The application divides the agile turning process into a deceleration section and a maneuvering turning section according to the release and separation time of the deceleration parachute, force analysis is carried out on each section, a dynamics model of the novel missile is established, and a segmented sliding mode controller is constructed based on the dynamics model, the segmented sliding mode controller is used for real-time control of the lateral RCS and the aerodynamic rudder, is suitable for the agile turning process under different attack angles, and can avoid the singularity phenomenon of the sliding mode control.

Description

Technical Field

[0001] This invention relates to the field of new concept weapon technology, specifically to a method for agile turning control of a missile. Background Technology

[0002] To adapt to modern warfare, the all-aspect attack capability of missiles has gradually become an important tactical indicator. Its core lies in the missile's ability to strike rear areas, requiring it to rapidly achieve agile turns at high angles. During this process, the missile maneuvers at extremely high angles of attack, entering a severe stall phase. The aerodynamic forces provided by the control surfaces are insufficient to generate enough control torque to achieve agile turns. Therefore, direct lateral force devices or thrust vectoring devices are typically introduced in engineering to supplement the control force, employing this composite control method to achieve agile missile turns. Furthermore, at high angles of attack, the missile's dynamics exhibit strong nonlinearity, rapid time-varying characteristics, and high uncertainty, with severe coupling between systems, which brings enormous difficulty and challenges to agile missile turn control.

[0003] However, conventional missile maneuvers often employ direct lateral force devices to provide the force and torque required for turning. But due to limitations in lateral pulse engine thrust and fuel availability, there is still significant room for improvement in missile maneuverability. Furthermore, while relying on direct lateral force to control the missile's attitude and achieve agile turns, the missile's speed remains relatively high, while its trajectory angle changes slowly. This results in a large turning radius, implying that there is room for further reduction in its turning radius.

[0004] Therefore, there is an urgent need to study a missile agile turning technology to further enhance missile maneuverability and significantly reduce the missile's turning radius. Summary of the Invention

[0005] In view of this, the present invention provides a missile agile turning control method, which can improve the missile's maneuverability and significantly reduce the missile's turning radius.

[0006] To achieve the above-mentioned objectives, the technical solution of this invention is as follows:

[0007] A method for agile turn control of a missile, the missile comprising a lateral reaction jet device and a missile body with aerodynamic rudders and a main engine, and a deceleration chute mounted on the missile; the agile turn control method includes:

[0008] The missile's agile turning process is divided into a deceleration phase and a maneuvering turning phase, depending on whether the drag chute is deployed. During the deceleration phase, the drag chute is deployed, while during the maneuvering turning phase, the drag chute has detached.

[0009] Based on the control effects of the drag chute, lateral reaction jet, aerodynamic rudder, and main engine on the missile, dynamic equations for pitch angle, angle of attack, and pitch acceleration are constructed. Specifically, during the deceleration phase, the missile decelerates and turns under the control of the drag chute, aerodynamic rudder, and lateral reaction jet; during the maneuver control phase, the missile turns under the combined control of the lateral reaction jet and the aerodynamic forces of the control surfaces. In both phases, if the pitch angle reaches 120°, the main engine is activated and remains activated.

[0010] A segmented sliding mode controller is constructed based on the dynamic equations, parameters are set, and control commands for the aerodynamic rudder and the lateral reaction jet device are calculated.

[0011] Furthermore, it also includes: the missile changes its pitch angle under the control commands of the aerodynamic rudder and the lateral reaction jet device; when the pitch angle reaches the target value, the control process ends; otherwise, the parameters of the segmented sliding mode controller are changed, and control commands are continued to be calculated and applied to the missile until the pitch angle reaches the target value.

[0012] Furthermore, during the deceleration phase, the forces that slow down the missile include the flexible drag generated after the deceleration parachute is released and the aerodynamic drag experienced by the missile itself during flight; the forces that turn the missile include the direct control force generated by the lateral reaction jet under the control command of the lateral reaction jet and the aerodynamic force generated by the aerodynamic rudder under the control command of the aerodynamic rudder.

[0013] During maneuvering turns, the forces that cause the missile to turn include the direct control force generated by the lateral reaction jet under the control command of the lateral reaction jet and the aerodynamic force generated by the aerodynamic rudder under the control command of the aerodynamic rudder; the forces that accelerate the missile toward the target include the thrust generated by the main engine according to the change in pitch angle.

[0014] When the missile's angle of attack α > 40°, the aerodynamic force on the control surface is 0; when the missile's angle of attack α ≤ 40°, the aerodynamic force on the control surface is not 0.

[0015] When the missile's pitch angle When the main engine's thrust changes according to the pitch angle, the thrust is 0; when the missile's pitch angle changes... At that time, the thrust generated by the main engine according to the change in pitch angle is not zero.

[0016] Furthermore, the expression for the missile's dynamic equations during the deceleration phase is as follows:

[0017]

[0018] Where V is the velocity of the missile's center of mass; Q is the dynamic pressure; m is the mass of the missile; S is the characteristic area; L is the characteristic length; C L C is the drag coefficient;D The drag coefficient; u P The main engine control switch; X P P represents the flexible drag of the deceleration parachute; P represents the engine thrust of the missile's main engine; u P The main engine control switch; α and θ represent the pitch angle, angle of attack, and trajectory inclination angle, respectively; ω z ω is the pitch angular velocity; z U is the pitch angular velocity; δ1 The rudder control effect during the deceleration phase is given by f, where f is a known quantity in the segmented sliding mode controller during the deceleration phase, and u is the rudder control effect during the deceleration phase. rcs For the control effect of the lateral reaction jet device; T rcs The maximum thrust of the lateral reaction jet; δ rcs For the control commands of the lateral reaction jet system; J z Let be the moment of inertia of the missile about the z-axis of the missile body coordinate system. The missile body coordinate system has its origin O at the missile's center of mass, the x-axis is the missile's flight direction, the y-axis is the missile's pitch direction, and the z-axis is obtained by the right-hand screw law.

[0019] The expression for the missile's dynamic equations during the maneuvering turn is as follows:

[0020]

[0021] Where f is a known quantity in the segmented sliding mode controller during the maneuvering turning segment.

[0022] Furthermore, the main engine control switch u P Satisfying constraints when At that time, the missile is in the deceleration phase, so the main engine is not activated; when At that time, the missile was in a maneuvering turn, and the main engine provided thrust, causing the missile to accelerate in the -X direction.

[0023] Furthermore, when the absolute value of the pitch angle tracking error signal is greater than or equal to 1, the expression for the sliding mode controller s1 is:

[0024] s1=x2+β 1 / v sign(x1)|x1| w |x1|≥1

[0025] The corresponding expression for the aerodynamic rudder control command is:

[0026]

[0027] The corresponding expression for the lateral reaction jet control command is:

[0028] δ rcs1 =-βv-1 sign(x2)|x2| 2-v -f-ks1-δsign(s1)-u δ

[0029] in, For pitch angle error, The target value for the pitch angle; x2 = ω z y is the pitch angular velocity; k, β, v, w are preset parameters of the segmented sliding mode controller; sign(·) is the sign function, Q is the dynamic pressure, S is the characteristic area, L is the characteristic length, C mδ J is the first derivative of the pitching moment coefficient with respect to the aerodynamic deflection angle δ. Z Let u be the moment of inertia of the missile about the Z-axis of the missile's coordinate system. δ For the rudder control effect, f1 is the known quantity in the segmented sliding mode controller during the deceleration phase, and δ is the aerodynamic rudder deflection angle; the missile body coordinate system takes the missile's center of mass as the origin O, the x-axis is the missile's flight direction, the positive y-axis is the missile's pitch direction, and the z-axis is obtained by the right-hand screw law.

[0030] When the absolute value of the pitch angle tracking error signal is less than 1, the expression for the non-singular sliding mode controller s2 is:

[0031]

[0032] The corresponding aerodynamic control command δ c2 The expression is:

[0033]

[0034] The corresponding lateral reaction jet control command δ rcs2 The expression is:

[0035] δ rcs2 =-wβ 1 / v |x1| w-1 x2-f-ks2-δsign(s2)

[0036] Where f is a known quantity in the segmented sliding mode controller.

[0037] Furthermore, the drag chute, aerodynamic rudders, and main engine are all mounted at the tail of the missile body, while the lateral reaction jet is mounted at the front of the missile body.

[0038] Beneficial effects:

[0039] 1. This invention proposes a missile agile turning control method. A deceleration parachute is installed on the missile to provide flexible drag. Under the combined forces of the parachute, lateral RCS, aerodynamic rudders, and main engine thrust, the missile turns towards the target. The deceleration parachute reduces the missile's velocity, significantly increasing the rate of change of the missile's angle of attack. During the agile turn, this manifests as a rapid increase in the missile's trajectory inclination angle, resulting in a smaller turning radius compared to traditional agile missiles, thus enhancing maneuverability. This invention divides the agile turn process into a deceleration phase and a maneuvering phase based on the angle of attack. Force analysis is performed on each phase to establish a missile dynamic model. Based on this dynamic model, a segmented sliding mode controller is constructed. This segmented sliding mode controller is used to control the lateral reaction jet and aerodynamic rudders in real time, applicable to agile turns at different angles of attack.

[0040] 2. This invention adds drag to the original aerodynamic drag by using a deceleration parachute. According to the dynamic equation, the magnitude of the drag on the missile directly affects the speed change. In the deceleration phase, thanks to the large aerodynamic drag of the deceleration parachute, the missile acceleration is less than zero and the angle of attack is also smaller. Compared with the traditional scheme without a deceleration parachute, the missile speed decreases faster.

[0041] 3. The main engine used in this invention is... The main engine is not activated at this time because the missile's pitch angle is relatively small, and activating the main engine would be detrimental to the missile's trajectory. When the missile is in a maneuvering turn, the main engine provides thrust, causing the missile to accelerate in the -X direction. This reasonable thrust distribution can improve the efficiency of agile turns.

[0042] 4. This invention employs different sliding mode controllers to control the missile's pitch angle during the deceleration and maneuvering turns, thereby obtaining control commands for the aerodynamic forces and direct control forces of the control surfaces at small and large angles of attack, avoiding peculiar phenomena that could affect the accuracy of missile control. Attached Figure Description

[0043] Figure 1 This is a flowchart of the method of the present invention.

[0044] Figure 2 This is a schematic diagram of the missile's structure.

[0045] Figure 3 This is a schematic diagram of the missile's agile turning process.

[0046] Among them, 1-missile, 2-deceleration parachute, 3-aerodynamic rudder, 4-lateral reaction jet device. Detailed Implementation

[0047] The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0048] like Figure 1 As shown, the present invention provides a method for agile turning control of a missile, the steps of which include:

[0049] Step 1: Install a drag chute on the missile to construct its geometry.

[0050] like Figure 2 As shown, the missile includes a missile body 1, a deceleration parachute 2, aerodynamic rudders 3, and a lateral reaction jet device 4. Its coordinate system has the missile's center of mass as the origin O, the x-axis as the missile's flight direction (positive direction is forward relative to the initial position, negative direction is backward relative to the initial position), the y-axis as the missile's pitch direction (positive direction is upward relative to the ground, negative direction is downward relative to the ground), and the z-axis is obtained using the right-hand screw law. The missile body 1 used in this invention is a wingless missile with a tail rudder, its tail rudder being an aerodynamic rudder 3, and the main engine is equipped at the tail. The deceleration parachute 2 is installed at the tail of the missile body 1, and the lateral reaction jet device 4 is installed in front of the missile's center of mass. The lateral reaction jet device 4 used in this invention is a lateral pulse engine capable of generating continuous thrust, with a thrust magnitude T. now At maximum thrust T rcs Continuous variation within a range. The control command for the lateral reaction jet is defined as δ. rcs The opening degree of the jet valve, representing the lateral reaction jet device 4, is expressed as:

[0051]

[0052] Among them, valve opening δ rcs The value range is [-1, 1]. When it is 1, the valve is fully open and jets air in the -y direction. When it is -1, the valve is fully open and jets air in the +y direction. In the initial stage of agile turning, the valve is set to a positive value to provide lateral force in the +y direction and reduce the turning radius. The valve opening command is given by the segmented sliding mode controller.

[0053] Step 2, as follows Figure 3 As shown, based on the missile's geometric configuration, the missile's agile turning process is divided into a deceleration phase with the deceleration parachute deployed and a maneuvering turning phase with the deceleration parachute detached, depending on whether the deceleration parachute is active. The parachute deployment time and duration can be selected within the deceleration phase as needed.

[0054] Step 3: Analyze the forces acting on the missile during the deceleration and maneuvering turns, and construct the corresponding dynamic models for the deceleration and maneuvering turns:

[0055] To simplify the analysis process, this invention proposes three preconditions: This invention only studies the forces acting on the missile in the pitch direction, therefore the direct control force provided by the lateral reaction jet device only acts on the y-axis; since the agile turning process is relatively rapid, it is assumed that the mass and moment of inertia of the missile remain almost unchanged during this process, and the change in the position of the missile's center of gravity is not considered; during the deceleration phase of the missile, the deceleration parachute 2 is considered to be released and opened instantaneously, and the detachment of the deceleration parachute 2 occurs instantaneously, so its impact on the missile is negligible.

[0056] Based on three preconditions, this invention analyzes the force situation of the missile during the deceleration phase. During the deceleration phase, the missile is subjected to multiple control forces, including the flexible force generated after the deceleration parachute 2 is released, the aerodynamic force of the control surface generated by the aerodynamic rudder 3 under the action of the aerodynamic rudder control command, and the direct control force generated by the lateral reaction jet device 4. Under the multiple control forces, the missile completes the deceleration and partial agile turning process.

[0057] Construct the dynamic model of the deceleration phase, with the following expression:

[0058]

[0059] Where ρ is the atmospheric density; V is the velocity of the missile's center of mass; For dynamic pressure; m is mass; S is characteristic area; L is characteristic length; C L and C D These represent the lift coefficient and drag coefficient, respectively; P is the engine thrust of the main engine. α and θ represent the pitch angle, angle of attack, and trajectory inclination angle, respectively; ω z Mz is the pitch angular velocity; Mz is the aerodynamic torque. When α < 40°, Mz = QSLC mα When α >= 40°, f = QSLC N C mα and C mδ L represents the first derivative of the pitching moment coefficient with respect to the angle of attack and the aerodynamic control deflection, respectively; rcs The distance between the lateral reaction jet and the missile's center of mass; m zd δ represents the sum of aerodynamic uncertainties and the impact of external disturbances on the pitching moment coefficient; δ is the aerodynamic rudder deflection angle; δ c τ is the aerodynamic control command; τ is the time constant of the aerodynamic control element; X and Y represent the horizontal and vertical coordinates of the missile's center of mass, respectively; J z Let u be the moment of inertia of the missile about the z-axis of the missile's coordinate system; P The main engine's control switch satisfies the constraints. when At that time, the missile was still far from the target with a pitch angle of 180°, and needed to continuously decelerate, so the main engine was not activated; when At that time, the missile needs the main engine to provide thrust to fly in the -X direction, therefore the main engine is activated; X P This is the flexible drag of the deceleration parachute.

[0060] During the maneuvering turn phase, the drag chute has disengaged. Therefore, at this stage, the missile is no longer affected by the aerodynamic drag of the drag chute, and its attitude is controlled by direct lateral forces and control surface aerodynamic forces. The expression for the dynamic model becomes:

[0061]

[0062] Among them, L s Let Mz be the distance from the center of pressure to the center of mass of the missile, and Mz be the aerodynamic torque. When α < 40°, Mz = QSLC. mα When α >= 40°, f = QSLC N C N This is the normal force coefficient.

[0063] During the deceleration and maneuvering turns, the presence of aerodynamic control surfaces depends on whether the missile's angle of attack is greater than 40°. If the missile's angle of attack is greater than 40°, the aerodynamic control surfaces will not function; otherwise, they will function. Once the missile's pitch angle is greater than or equal to 120°, the main engine will ignite and begin to function. The main engine will not be shut down after it is started, and the missile will continue to turn and accelerate toward the target under the action of the main engine.

[0064] Step 4: Based on the two-stage dynamic model, and according to whether the absolute value of the pitch angle tracking error is greater than 1, a segmented non-singular sliding mode controller is constructed. This controller operates during the missile deceleration phase and the maneuvering turn phase. The segmented sliding mode controller outputs different control commands at different stages to control the entire agile turn process.

[0065] This invention considers an extreme case of agile turning, where the missile's flight direction is completely reversed from its initial trajectory during an agile turn, resulting in a pitch angle of 180° after the agile turn. The target value for the preset pitch angle is... Considering the control effect of multi-source control forces, based on angular acceleration in the dynamic model From the expression, we construct the missile's pitch angle error equation, which is expressed as follows:

[0066]

[0067] in, For pitch angle error, x2=ω z The pitch angular velocity is given by f, where f is a known quantity in the segmented sliding mode controller, and u is the pitch angular velocity. δ For the control effect of the aerodynamic rudder 3, u rcs This refers to the control effect of the lateral reaction jet device 4.

[0068] Based on the dynamic model, a piecewise sliding mode controller s is constructed to control the pitch angle error x1, and its expression is:

[0069]

[0070] Where β, v, and w are the parameters of the piecewise sliding mode controller s (β > 0, 1 < v < 2, w > 1), and sign(·) is the sign function. From the expression of the piecewise sliding mode controller s, it can be seen that when |x1| ≥ 1, the pitch angle error is far from the equilibrium point, and a traditional sliding mode controller is used, resulting in fast convergence. When |x1| < 1, a non-singular sliding mode controller is used, thus avoiding singular phenomena.

[0071] Because the forces acting on the missile differ during the deceleration and maneuvering turns, the control effects of each control force also differ. This invention achieves agile turning by controlling the missile's attitude, and obtains f and u based on the missile attitude control dynamics model. δ and u rcs Piecewise expressions:

[0072]

[0073]

[0074]

[0075] Where f1 is a known quantity in the segmented sliding mode controller when α < 40°, u δ1 f1 represents the control effect of the aerodynamic rudder when α < 40°; f2 represents the known quantity in the segmented sliding mode controller when α > = 40°; u δ2 The control effect of the aerodynamic rudder when α>=40°.

[0076] Set the parameters k, β, v, w (k > 0) of the segmented sliding mode controller s, respectively, and calculate the aerodynamic rudder control command δ when |x1| >= 1. c1 and lateral reaction jet control command δ rcs1 And calculate the aerodynamic rudder control command δ when |x1|<1. c2 and lateral reaction jet control command δ rcs2 Its expression is:

[0077]

[0078]

[0079] Where, δ c For agile turning, δ rcsControl commands for the lateral reaction jet system during agile turning.

[0080] Step 5: Convert the control command δ output by the segmented sliding mode controller s c δ rcs The control process is applied to the missile at the current moment to measure the pitch angle. When the pitch angle reaches the target value of 180°, the control flow ends; otherwise, the parameters of the segmented sliding mode controller are changed, and the control command δ is calculated again. c δ rcs And apply it to the missile at that moment until the pitch angle reaches the target value of 180°.

[0081] In summary, the above are merely preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for agile turning control of a missile, the missile comprising a lateral reaction jet device and a missile body with aerodynamic rudders and a main engine, characterized in that, The missile is equipped with a deceleration chute; the agile turning control method includes: The missile's agile turning process is divided into a deceleration phase and a maneuvering turning phase based on whether the drag chute is deployed. During the deceleration phase, the drag chute is deployed, and during the maneuvering turning phase, the drag chute has detached. Based on the control effects of the drag chute, lateral reaction jet, aerodynamic rudder, and main engine on the missile, dynamic equations for pitch angle, angle of attack, and pitch acceleration are constructed. Specifically, during the deceleration phase, the missile decelerates and turns under the control of the drag chute, aerodynamic rudder, and lateral reaction jet; during the maneuvering and turning phase, the missile turns under the combined control of the lateral reaction jet and the aerodynamic forces of the control surfaces. In both phases, if the pitch angle reaches 120°, the main engine is activated and remains operational. A segmented sliding mode controller is constructed based on the dynamic equations, parameters are set, and control commands for the aerodynamic rudder and the lateral reaction jet device are calculated. When the absolute value of the pitch angle tracking error signal is greater than or equal to 1, the expression for the sliding mode controller s1 is: s1=x2+β 1 / v sign(x1)|x1| w |x1|≥1 The corresponding expression for the aerodynamic rudder control command is: The corresponding expression for the lateral reaction jet control command is: δ rcs1 =-βv -1 sign(x2)|x2| 2-v -f-ks1-δsign(s1)-u δ in, For pitch angle error, The target value for the pitch angle; x2 = ω z y is the pitch angular velocity; k, β, v, w are preset parameters of the segmented sliding mode controller; sign(·) is the sign function, Q is the dynamic pressure, S is the characteristic area, L is the characteristic length, C mδ J is the first derivative of the pitching moment coefficient with respect to the aerodynamic deflection angle δ. Z Let u be the moment of inertia of the missile about the Z-axis of the missile's coordinate system. δ For the rudder control effect, f1 is the known quantity in the segmented sliding mode controller during the deceleration phase, and δ is the aerodynamic rudder deflection angle; the missile body coordinate system takes the missile's center of mass as the origin O, the x-axis is the missile's flight direction, the positive y-axis is the missile's pitch direction, and the z-axis is obtained by the right-hand screw law; When the absolute value of the pitch angle tracking error signal is less than 1, the expression for the non-singular sliding mode controller s2 is: The corresponding aerodynamic control command δ c2 The expression is: The corresponding lateral reaction jet control command δ rcs2 The expression is: δ rcs2 =-wβ 1 / v |x1| w-1 x2-f-ks2-δsign(s2) Where f is a known quantity in the segmented sliding mode controller.

2. The method as described in claim 1, characterized in that, Also includes: The missile changes its pitch angle under the control commands of the aerodynamic rudder and the lateral reaction jet device. When the pitch angle reaches the target value, the control process ends; otherwise, the parameters of the segmented sliding mode controller are changed, and control commands are continued to be calculated and applied to the missile until the pitch angle reaches the target value.

3. The method as described in claim 1, characterized in that, During the deceleration phase, the forces that slow down the missile include the flexible drag generated after the deceleration parachute is released and the aerodynamic drag experienced by the missile itself during flight; the forces that turn the missile include the direct control force generated by the lateral reaction jet under the control command of the lateral reaction jet and the aerodynamic force generated by the aerodynamic rudder under the control command of the aerodynamic rudder. During the maneuvering turn, the forces that cause the missile to turn include the direct control force generated by the lateral reaction jet under the control command of the lateral reaction jet and the aerodynamic force generated by the aerodynamic rudder under the control command of the aerodynamic rudder; the forces that accelerate the missile toward the target include the thrust generated by the main engine according to the change of pitch angle. When the missile's angle of attack α > 40°, the aerodynamic force of the control surface is 0; when the missile's angle of attack α ≤ 40°, the aerodynamic force of the control surface is not 0. When the missile's pitch angle When the main engine's thrust changes according to the pitch angle, the thrust is 0; when the missile's pitch angle changes... At that time, the thrust generated by the main engine according to the change in pitch angle is not zero.

4. The method as described in claim 1 or 3, characterized in that, The expression for the missile's dynamic equations during the deceleration phase is as follows: Where V is the velocity of the missile's center of mass; Q is the dynamic pressure; m is the mass of the missile; S is the characteristic area; L is the characteristic length; C L C is the drag coefficient; D The drag coefficient; u P The main engine control switch; X P P represents the flexible drag of the deceleration parachute; P represents the engine thrust of the missile's main engine. α and θ represent the pitch angle, angle of attack, and trajectory inclination angle, respectively; ω z U is the pitch angular velocity; δ1 The rudder control effect during the deceleration phase is given by f, where f is a known quantity in the segmented sliding mode controller during the deceleration phase, and u is the rudder control effect during the deceleration phase. rcs For the control effect of the lateral reaction jet device; T rcs The maximum thrust of the lateral reaction jet; δ rcs For the control commands of the lateral reaction jet system; J z Let be the moment of inertia of the missile about the z-axis of the missile body coordinate system. The missile body coordinate system has its origin O at the missile's center of mass, the x-axis is the missile's flight direction, the y-axis is the missile's pitch direction, and the z-axis is obtained by the right-hand screw law. The expression for the missile's dynamic equations during the maneuvering turn is as follows: Where f is a known quantity in the segmented sliding mode controller during the maneuvering turning segment.

5. The method as described in claim 3, characterized in that, Main engine control switch u P Satisfying constraints when At that time, the missile is in the deceleration phase, so the main engine is not activated; when At that time, the missile was in a maneuvering turn, and the main engine provided thrust, causing the missile to accelerate in the -X direction.

6. The method as described in claim 1, 2, 3 or 5, characterized in that, The deceleration parachute, the aerodynamic rudder, and the main engine are all mounted at the tail of the missile body, while the lateral reaction jet device is mounted at the front of the missile body.

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