A high-maneuver aircraft guidance and control method based on thrust vector

Abstract

The present invention provides a guidance and control method for a highly maneuverable aircraft based on thrust vector, which includes constructing a dynamic system of the highly maneuverable aircraft and a nonlinear model of the thrust vector object; designing an integrated controller for highly maneuverable guidance based on the coupling characteristics of the guidance loop and the control loop; using the designed integrated controller to control the highly maneuverable aircraft to fly on a predetermined reference profile, so as to achieve the tracking of speed and altitude. The present invention solves the problem that the guidance loop and the control loop cannot be considered together in the case of thrust vector of the highly maneuverable aircraft, and improves the control stability and maneuver agility of the highly maneuverable aircraft; through integrated control, the present invention reduces the influence of the thrust vector of the highly maneuverable aircraft on the centroid loop; thereby effectively reducing the altitude tracking error and adjustment time; the controller of the present invention can control the highly maneuverable aircraft to fly within a predetermined reference profile, so as to achieve the tracking of the speed and altitude of the highly maneuverable aircraft.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft guidance and control, and in particular to a guidance and control method for a highly maneuverable aircraft based on thrust vectoring. Background Art

[0002] As a vehicle capable of flying at hypersonic speed, compared with medium- and low-speed aircraft, a hypersonic vehicle has the capabilities of global precision strike and rapid maneuverability, can achieve flexible combat support, and is also a powerful weapon against nuclear threats. Therefore, it has great military and civilian value, and its design and development level represents the country's comprehensive scientific and technological strength, and is also attracting more and more attention.

[0003] However, due to its special structure and unique flight conditions, the hypersonic vehicle is extremely sensitive to aerodynamic parameters and has highly nonlinear dynamic characteristics. Due to these factors, the control design of the hypersonic vehicle is much more difficult than that of traditional aircraft.

[0004] The existing technology cannot take into account both the guidance loop and the control loop under the thrust condition of a highly maneuverable aircraft, resulting in poor control stability and maneuver agility of the highly maneuverable aircraft. The existing technology cannot well solve the above technical problems. Summary of the Invention

[0005] Aiming at the deficiencies of the existing technology, the present invention provides a guidance and control method for a highly maneuverable aircraft based on thrust vectoring. The present invention solves the problem that the guidance loop and the control loop cannot be considered together under the thrust vector condition of a highly maneuverable aircraft, and improves the control stability and maneuver agility of the highly maneuverable aircraft.

[0006] The technical solution of the present invention is: a guidance and control method for a highly maneuverable aircraft based on thrust vectoring, including the following steps:

[0007] S1), constructing a dynamic system of a highly maneuverable aircraft;

[0008] S2), constructing a nonlinear model of a thrust vector object;

[0009] S3), designing an integrated controller for highly maneuverable guidance based on the coupling characteristics of the guidance loop and the control loop;

[0010] S4), using the designed integrated controller to control the highly maneuverable aircraft to fly on a predetermined reference profile to achieve speed and altitude tracking.

[0011] Preferably, in step S1), the construction of the dynamic system of the highly maneuverable aircraft specifically includes the following steps:

[0012] S11), establish the translational dynamics vector equation of the center of mass of the highly maneuverable aircraft in the ground coordinate system:

[0013]

[0014] Among them, m is the instantaneous mass of the highly maneuverable aircraft, with the unit of kg; V is the velocity vector of the highly maneuverable aircraft, with the unit of m / s; R and G are the aerodynamic force and gravity respectively, with the unit of N;

[0015] S12), since the orbital coordinate system has displacement motion and rotational motion relative to the ground coordinate system; therefore, from the relationship between the relative derivative and the absolute derivative of the vector, we get:

[0016]

[0017] Among them, is the time derivative of the highly maneuverable aircraft relative to the orbital coordinate system, with the unit of m / s 2 ; Ω is the rotational angular velocity vector of the orbital coordinate system, with the unit of rad / s;

[0018] S13), according to the definition of the orbital coordinate system, the components of the center of mass velocity vector V of the highly maneuverable aircraft on the Ox2y2z2 axes are:

[0019]

[0020] In the formula, V x2 、V y2 、V z2 are the components of the center of mass velocity vector V of the highly maneuverable aircraft on the Ox2y2z2 axes respectively;

[0021] S14), apply the dynamic equation to obtain the scalar form of the dynamic equation of the translational motion of the center of mass of the highly maneuverable aircraft:

[0022]

[0023] Among them, t is the time; θ is the flight path angle; γ is the angle of attack; T, D, and L are the thrust, drag, and lift respectively.

[0024] Preferably, in step S2), the thrust vector object includes the total pitch thrust vector deflection angle δ z 、the total yaw thrust vector deflection angle δ y 、the roll thrust vector deflection angle δ x .

[0025] Preferably, in step S2), the total pitch thrust vector deflection angle δ z is expressed as:

[0026] δ z =(δzr +δ zl ) / 2;

[0027] Wherein, δ zi (i = l, r) is the pitch thrust vector deflection angle, l and r respectively represent the left and right nozzles, and the downward deflection of the vector nozzle axis is positive.

[0028] Preferably, in step S2), the total yaw thrust vector deflection angle δ y is expressed as:

[0029] δ y =(δ yr +δ yl ) / 2;

[0030] Wherein, δ yi is the yaw thrust vector deflection angle; the left deflection of the vector nozzle axis is positive.

[0031] Preferably, in step S2), half of the difference between the pitch thrust vector deflection angles of the left and right nozzles is defined as the roll thrust vector deflection angle δ x , that is:

[0032] δ x =(δ xr -δ xl ) / 2;

[0033] δ x is positive when generating a moment that makes the aircraft roll to the left.

[0034] Preferably, in step S2), it also includes establishing an influence model of the thrust vector object on the longitudinal speed loop of the highly maneuverable aircraft.

[0035] Preferably, in step S2), the expression of the dynamic equation of the thrust vector object on the longitudinal speed loop of the highly maneuverable aircraft is:

[0036]

[0037] In the formula, δ T is the thrust vector deflection angle; T is the thrust term; T real is the true thrust term; α is the angle of attack; D is the drag; V is the velocity vector of the highly maneuverable aircraft.

[0038] Preferably, in step S2), for the highly maneuverable aircraft during the altitude loop control process, the thrust vector and the aerodynamic rudder are distributed in a chain-like manner, that is, the thrust vector deflection is triggered when the aerodynamic rudder cannot meet the maneuverability requirements; and it is assumed that the thrust coefficient k Φ satisfies:

[0039]

[0040] In the formula, K and are respectively the maximum and minimum values of the thrust coefficient; k Φ is the thrust coefficient; δ T is the thrust vector deflection angle.

[0041] Preferably, in step S2), the thrust term T and the fuel ratio Φ satisfy a linear relationship:

[0042]

[0043] In the formula, k Φ is the thrust coefficient, and its value corresponds to the ratio of the maximum thrust to the maximum fuel ratio.

[0044] Preferably, in step S3), the integrated controller adopts a full-state joint design of the guidance subsystem and the attitude control subsystem based on the spectrum separation principle, regards the guidance subsystem as the outer-loop slow control loop, and the attitude control subsystem as the inner-loop fast control loop; the guidance airflow angle command is realized by adjusting the attitude of the aircraft.

[0045] Preferably, in step S3), the construction of the integrated controller for high-maneuver guidance specifically includes the following steps:

[0046] S31), calculate the speed and altitude errors of the high-maneuver aircraft; wherein, the speed error is the difference between the current speed value and the speed value of the nominal trajectory; the altitude error is the difference between the current altitude value and the nominal altitude value;

[0047] S32), design the control system for the altitude error and the speed error in a manner that meets the requirements of the Lyapunov function;

[0048] S33), connect the designed control system to the output end; the output end includes the control surface and the engine fuel ratio.

[0049] Preferably, in step S32), the expression of the integrated controller for high-maneuver guidance is:

[0050]

[0051] In the formula, is the derivative of the Lyapunov function; e is the speed error, e = V - V r , where V is the current speed value; V r is the speed value of the given nominal trajectory; k Φ is the thrust coefficient; Φ is the fuel ratio; δ T is the thrust vector deflection angle; D is the drag; m is the command; k is the controller parameter; is the derivative of the current speed; is the derivative of the reference speed.

[0052] Preferably, in step S4), an integrated controller is used to control the flight of the highly maneuverable aircraft, which specifically includes the following steps:

[0053] S41) Obtain the reference altitude and reference speed profile of the constant-altitude maneuverable aircraft according to the combat mission of the highly maneuverable aircraft;

[0054] S42) For the speed loop, input the speed error between the current speed and the reference speed into the integrated controller; the integrated controller outputs the fuel ratio according to the input speed error; adjust the speed of the highly maneuverable aircraft by controlling the output of the fuel ratio; automatically adjust the fuel ratio output of the aircraft by inputting the speed error; perform acceleration or deceleration operations on the highly maneuverable aircraft to reduce the error value between the aircraft speed and the speed profile; finally, achieve the tracking of the flight speed by the highly maneuverable aircraft;

[0055] S43) For the altitude loop, input the altitude error between the altitude and the reference altitude into the integrated controller; the integrated controller obtains the elevator deflection angle according to the input altitude error; and obtains the control moment according to the elevator deflection angle; control the flight altitude of the highly maneuverable aircraft according to the control moment term.

[0056] The beneficial effects of the present invention are as follows:

[0057] 1. The present invention solves the problem that the guidance loop and the control loop cannot be considered together in the case of thrust vector of the highly maneuverable aircraft, and improves the control stability and maneuver agility of the highly maneuverable aircraft;

[0058] 2. The present invention reduces the influence of the thrust vector of the highly maneuverable aircraft on the centroid loop through integrated control; thus, the altitude tracking error and adjustment time can be effectively reduced;

[0059] 3. The controller of the present invention can control the highly maneuverable aircraft to fly within a predetermined reference profile, thereby achieving the speed and altitude tracking of the highly maneuverable aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0060] Figure 1 is a schematic flow chart of the control method according to an embodiment of the present invention;

[0061] Figure 2 is a diagram of the integrated control tracking error situation of the highly maneuverable aircraft according to an embodiment of the present invention. DETAILED DESCRIPTION OF THE INVENTION

[0062] The following further describes the specific embodiments of the present invention with reference to the drawings:

[0063] As Figure 1As shown in the figure, this embodiment provides a guidance and control method for a highly maneuverable aircraft based on thrust vector, including the following steps:

[0064] S1). Construct the dynamic system of the highly maneuverable aircraft; specifically including the following steps:

[0065] S11). In the ground coordinate system, establish the translational dynamics vector equation of the center of mass of the highly maneuverable aircraft:

[0066]

[0067] where m is the instantaneous mass of the highly maneuverable aircraft, with the unit of kg; V is the velocity vector of the highly maneuverable aircraft, with the unit of m / s; R and G are the aerodynamic force and gravitational force respectively, with the unit of N;

[0068] S12). Since the orbital coordinate system has displacement motion and rotational motion relative to the ground coordinate system; therefore, from the relationship between the relative derivative and the absolute derivative of the vector, we get:

[0069]

[0070] where, is the time derivative of the highly maneuverable aircraft relative to the orbital coordinate system, with the unit of m / s 2 ; Ω is the rotational angular velocity vector of the orbital coordinate system, with the unit of rad / s;

[0071] S13). According to the definition of the orbital coordinate system, the components of the center of mass velocity vector V of the highly maneuverable aircraft on the Ox2y2z2 axes are:

[0072]

[0073] In the formula, V x2 、V y2 、V z2 are the components of the center of mass velocity vector V of the highly maneuverable aircraft on the Ox2y2z2 axes respectively;

[0074] S14). Apply the dynamic equation to obtain the scalar form of the dynamic equation of the translational motion of the center of mass of the highly maneuverable aircraft:

[0075]

[0076] where t is the time; θ is the flight path angle; γ is the angle of attack; T, D, and L are the thrust, drag, and lift respectively.

[0077] S2). Construct the nonlinear model of the thrust vector object; the thrust vector object includes the total pitch thrust vector deflection angle δ z 、the total yaw thrust vector deflection angle δ y 、the roll thrust vector deflection angle δ x。

[0078] Among them, the total pitch thrust vector deflection angle δ z is expressed as:

[0079] δ z =(δ zr +δ zl ) / 2;

[0080] In the formula, δ zi (i = l, r) is the pitch thrust vector deflection angle, l and r respectively represent the left and right nozzles, and the downward deflection of the vector nozzle axis is positive.

[0081] The total yaw thrust vector deflection angle δ y is expressed as:

[0082] δ y =(δ yr +δ yl ) / 2;

[0083] In the formula, δ yi is the yaw thrust vector deflection angle; the left deflection of the vector nozzle axis is positive.

[0084] Half of the difference between the pitch thrust vector deflection angles of the left and right nozzles is defined as the roll thrust vector deflection angle δ x , that is:

[0085] δ x =(δ xr -δ xl ) / 2;

[0086] δ x is positive when generating a moment that makes the aircraft roll to the left.

[0087] Establish an influence model of the thrust vector object on the longitudinal speed loop of a highly maneuverable aircraft. The expression of the dynamic equation of the thrust vector object on the longitudinal speed loop of a highly maneuverable aircraft is:

[0088]

[0089] In the formula, δ T is the thrust vector deflection angle; T is the thrust term; T real is the true thrust term; α is the angle of attack; D is the drag; V is the speed vector of the highly maneuverable aircraft.

[0090] In this embodiment, for a highly maneuverable aircraft during the altitude loop control process, the thrust vector and the aerodynamic rudder are distributed in a chain-like manner, that is, the thrust vector deflection is triggered when the aerodynamic rudder cannot meet the maneuverability requirements; and it is assumed that the thrust coefficient k Φ satisfies:

[0091]

[0092] In the formula, K, are respectively the maximum and minimum values of the thrust coefficient; k Φ is the thrust coefficient; δ T is the thrust vector deflection angle.

[0093] The thrust term T and the fuel ratio Φ satisfy a linear relationship:

[0094]

[0095] In the formula, k Φ is the thrust coefficient, and its value corresponds to the ratio of the maximum thrust to the maximum fuel ratio.

[0096] S3) Design an integrated controller for high-maneuver guidance based on the coupling characteristics of the guidance loop and the control loop;

[0097] In this embodiment, the integrated controller adopts a full-state joint design of the guidance subsystem and the attitude control subsystem based on the spectrum separation principle, regards the guidance subsystem as the outer-loop slow control loop, and the attitude control subsystem as the inner-loop fast control loop; realizes the guidance air flow angle command by adjusting the attitude of the aircraft.

[0098] In this embodiment, the full-state joint design of guidance and attitude control based on the spectrum separation principle means that when analyzing and designing the control system, the guidance control problem is decomposed into two parts: the guidance subsystem and the attitude control subsystem, and it is assumed that these two sub-systems are separated in the frequency domain, that is, they have no overlap or interference in the frequency response.

[0099] Specifically, in the guidance control system of the aircraft, the guidance subsystem is mainly responsible for determining the optimal path of the aircraft from the current position to the target, that is, trajectory planning, while the attitude control subsystem is responsible for adjusting the attitude of the aircraft in real time to achieve the guidance air flow angle command. And the guidance subsystem and the attitude control subsystem adopt the spectrum separation principle, that is, it is assumed that the guidance subsystem is a slow control loop with relatively slow dynamic characteristics, mainly affecting the performance of the aircraft in the low-frequency region; while the attitude control subsystem is a fast control loop with relatively fast dynamic characteristics, mainly affecting the performance of the aircraft in the high-frequency region.

[0100] In the aircraft control system, guidance and attitude control are two key components, and they achieve precise control of the aircraft through full-state joint design. Specifically:

[0101] Guidance subsystem: The guidance subsystem is regarded as the outer-loop slow control loop. Its main task is to determine the optimal path for the aircraft from the current position to the specified target, that is, the "trajectory". This process not only includes path planning but also generating desired control commands such as speed, angular velocity, and acceleration according to the planned path to guide the aircraft to fly along the predetermined trajectory.

[0102] Attitude control subsystem: The attitude control subsystem serves as the inner-loop fast control loop, responsible for adjusting the attitude of the aircraft in real time to ensure the realization of the guidance airflow angle command. This involves the control of attitude angles such as the angle of attack, bank angle, and sideslip angle of the aircraft, as well as the control of angular velocities such as roll, pitch, and yaw.

[0103] The core idea of the full-state joint design is to comprehensively consider the state variables of the guidance and attitude control subsystems, improve their integration, and thus optimize the flight performance of the aircraft. This design method not only reduces the number of iterations in controller design and the number of required sensors but also takes into account the coupling between subsystems, ensuring the stability and reliability of the system.

[0104] The construction of the integrated controller for high-maneuver guidance specifically includes the following steps:

[0105] S31): Calculate the speed and altitude errors of the high-maneuver aircraft; where the speed error is the difference between the current speed value and the speed value of the nominal trajectory; the altitude error is the difference between the current altitude value and the nominal altitude value.

[0106] S32): Design the control system according to the requirements of the Lyapunov function for the altitude error and speed error.

[0107] S33): Connect the designed control system to the output end; the output end includes the control surface and the engine fuel ratio. The expression of the integrated controller for high-maneuver guidance is as follows:

[0108]

[0109] In the formula, is the derivative of the Lyapunov function; e is the speed error, e = V - V r , where V is the current speed value; V r is the speed value of the given nominal trajectory; k Φ is the thrust coefficient; Φ is the fuel ratio; δ T is the thrust vector deflection angle; D is the drag; m is the command; k is the controller parameter; is the derivative of the current speed; is the derivative of the reference speed.

[0110] S4). Use the designed integrated controller to control the highly maneuverable aircraft to fly on a predetermined reference profile, achieving the tracking of speed and altitude. The specific steps are as follows:

[0111] S41). Obtain the reference altitude and reference speed profile of the constant altitude maneuverable aircraft according to the combat mission of the highly maneuverable aircraft;

[0112] S42). For the speed loop, input the speed error between the current speed and the reference speed into the integrated controller; the integrated controller outputs the fuel ratio according to the input speed error; adjust the speed of the highly maneuverable aircraft by controlling the output of the fuel ratio; automatically adjust the output of the aircraft's fuel ratio by inputting the speed error; perform acceleration or deceleration operations on the highly maneuverable aircraft to reduce the error value between the aircraft speed and the speed profile; finally achieve the tracking of the flight speed of the highly maneuverable aircraft;

[0113] S43). For the altitude loop, input the altitude error between the altitude and the reference altitude into the integrated controller; the integrated controller obtains the elevator deflection angle according to the input altitude error; and obtains the control moment according to the elevator deflection angle; control the flight altitude of the highly maneuverable aircraft according to the control moment term.

[0114] In this embodiment, the designed controller of this embodiment and the existing separately designed control are used to simulate the maneuver mission of the highly maneuverable aircraft diving from 6000 m to 1000 m altitude. It can be seen from Figure 2 this that, compared with the existing controller, the upper bound index of the altitude tracking error of the controller of this embodiment is reduced by 61.03%; and the altitude tracking adjustment time index is reduced by 22.76%. Thus, it is proved that the integrated controller of the present invention can well control the highly maneuverable aircraft to fly within a predetermined reference profile.

[0115] The above embodiments and the descriptions in the specification only illustrate the principles and the best embodiments of the present invention. Without departing from the spirit and scope of the present invention, the present invention will have various changes and improvements, and these changes and improvements all fall within the scope of the present invention claimed.

Claims

1. A high maneuverability aircraft guidance and control method based on thrust vectoring, characterized in that: The following steps are involved: S1), constructing a high maneuverability aircraft dynamics system; S2), constructing a nonlinear model of the thrust vector object; S3) Design an integrated controller for high maneuverability guidance based on the coupling characteristics of the guidance loop and the control loop; The integrated controller adopts a full-state joint design of the guidance subsystem and the attitude control subsystem based on the spectrum separation principle, and regards the guidance subsystem as an outer slow control loop and the attitude control subsystem as an inner fast control loop; the guidance airflow angle instruction is realized by adjusting the attitude of the aircraft; The construction of the integrated controller for high mobility guidance specifically includes the following steps: S31), calculating the speed error and altitude error of the high maneuverability aircraft; wherein the speed error is the difference between the current speed value and the speed value of the nominal trajectory; the altitude error is the difference between the current altitude value and the nominal altitude value; S32), design the control system by using the height error and speed error in a manner that meets the requirements of the Lyapunov function; its expression is: In the formula, is the derivative of the Lyapunov function; e is the velocity error, e = VV r , where V is the current speed; V r is the velocity value of a given nominal trajectory; k Φ is the thrust coefficient; Φ is the fuel ratio; δ T is the thrust vector deflection angle; D is the resistance; m is the command; k is the controller parameter; is the derivative of the current speed; is the derivative of the reference speed; S33), connecting the designed control system to the output end; the output end includes the ratio of the rudder surface to the engine fuel; S4), using the designed integrated controller to control the high maneuverability aircraft to fly in a predetermined reference profile to achieve speed and altitude tracking; specifically including the following steps: S41), obtaining a reference altitude and a reference speed profile of a high-altitude maneuverable aircraft according to a high-altitude maneuverable aircraft combat mission; S42), for the speed loop, input the speed error between the current speed and the reference speed into the integrated controller; the integrated controller outputs the fuel ratio according to the input speed error; the speed of the high maneuverable aircraft is adjusted by controlling the output of the fuel ratio; the fuel ratio output of the aircraft is automatically adjusted by inputting the speed error; the high maneuverable aircraft is accelerated or decelerated to reduce the error value between the aircraft speed and the speed profile; and finally the high maneuverable aircraft is tracked to the flight speed; S43), for the altitude loop, inputting the altitude error between the altitude and the reference altitude into the integrated controller; the integrated controller obtains the elevator angle according to the input altitude error; and obtains the control torque according to the elevator angle; and controls the flight altitude of the high maneuverable aircraft according to the control torque term.

2. The high maneuverability aircraft guidance and control method based on thrust vectoring according to claim 1, characterized in that: In step S1), the construction of the high maneuverability aircraft dynamics system specifically includes the following steps: S11), in the ground coordinate system, establish the center of mass translation dynamics vector equation of the high maneuverability aircraft: Wherein, m is the instantaneous mass of the high maneuverable aircraft, in kg; V is the velocity vector of the high maneuverable aircraft, in m / s; R and G are aerodynamic force and gravity, in N, respectively; S12), since the orbital coordinate system has displacement motion and rotation motion relative to the ground coordinate system; therefore, the relationship between the relative derivative and the absolute derivative of the vector is obtained: in, is the time derivative of the highly maneuverable aircraft relative to the orbital coordinate system, in m / s 2 ;Ω is the rotational angular velocity vector of the orbital coordinate system, in rad / s; S13), according to the definition of the orbital coordinate system, the components of the high maneuverability aircraft center of mass velocity vector V on the Ox2y2z2 axes are: Where V x2 、V y2 、V z2 are the components of the high maneuverable aircraft's center-of-mass velocity vector V on the Ox2y2z2 axes respectively; S14), applying the dynamics equation to obtain the scalar form of the dynamics equation of the center of mass translation of the high maneuverability aircraft: Where, t is time; θ is the track angle; γ is the angle of attack; T, D, and L are thrust, drag, and lift, respectively.

3. The high maneuverability aircraft guidance and control method based on thrust vectoring according to claim 1, characterized in that: In step S2), the thrust vector object includes the total pitch thrust vector deflection angle δ z 、Total yaw thrust vector angle δ y 、Roll thrust vector deflection angle δ x .

4. The high maneuverability aircraft guidance and control method based on thrust vectoring according to claim 3 is characterized in that: In step S2), the total pitch thrust vector deflection angle δ z It is expressed as: d z =(δ zr +d zl ) / 2; Among them, δ zi (i=l,r) is the pitch thrust vector deflection angle, l and r represent the left and right nozzles respectively, and the downward deflection of the vector nozzle axis is positive.

5. The high maneuverability aircraft guidance and control method based on thrust vectoring according to claim 3 is characterized in that: In step S2), the total yaw thrust vector deflection angle δ y It is expressed as: d y =(δ yr +d yl ) / 2; Among them, δ yi is the yaw thrust vector deflection angle; the left deviation of the vector nozzle axis is positive.

6. The high maneuverability aircraft guidance and control method based on thrust vectoring according to claim 3 is characterized in that: In step S2), half of the difference between the pitch thrust vector deflection angles of the left and right nozzles is defined as the roll thrust vector deflection angle δ x ,Right now: d x =(δ xr -d xl ) / 2; When the moment to roll the aircraft to the left is generated, x Is positive.

7. The high maneuverability aircraft guidance and control method based on thrust vectoring according to claim 3 is characterized in that: Step S2) also includes establishing an influence model of the thrust vector object on the longitudinal velocity loop of the high maneuverability aircraft; the expression of the dynamic equation of the thrust vector object on the longitudinal velocity loop of the high maneuverability aircraft is: In the formula, δ T is the thrust vector deflection angle; T is the thrust term; T real is the real thrust term; α is the angle of attack; D is the drag; V is the velocity vector of the high maneuverable aircraft; For high maneuverability aircraft, during the altitude control process, the thrust vector and the aerodynamic rudder are distributed in a chain manner, that is, the thrust vector deflection is triggered when the aerodynamic rudder cannot meet the maneuverability requirements; and assuming that the thrust coefficient k Φ satisfy: In the formula, K, are the maximum and minimum values ​​of the thrust coefficient respectively; k Φ is the thrust coefficient; δ T is the thrust vector deflection angle; The thrust term T and the fuel ratio Φ satisfy a linear relationship: In the formula, k Φ is the thrust coefficient, and its value corresponds to the ratio of maximum thrust to maximum fuel ratio.

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