Whole-course guidance control joint design method for cross-atmosphere reentry vehicle

Abstract

The invention provides a whole-course guidance control joint design method for a cross-atmosphere reentry vehicle, and the method comprises the steps: 1, building a six-degree-of-freedom motion model of the cross-atmosphere reentry vehicle, and building a simulation model of the cross-atmosphere reentry vehicle based on the six-degree-of-freedom motion model; 2, defining constraint conditions in a reentry flight process of the cross-atmosphere reentry vehicle; and step 3, performing guidance control joint design of different flight sections based on the constraint conditions defined in the step and the established simulation model. According to the method, the guidance loop and the attitude control loop are integrally designed, so that the performance of the aircraft is improved, and the design cost is reduced.

Description

Technical Field

[0001] The present invention relates to the field of guidance and control technology, and in particular to a combined method for full-course guidance and control of a trans-atmospheric reentry vehicle. Background Art

[0002] A space re-entry hypersonic vehicle refers to a type of aircraft that relies on a booster or booster rocket to enter the outer space orbit of the atmosphere, brakes to re-enter the atmosphere after completing preset flight instructions or performing specific tasks in outer space, re-enters the atmosphere at a hypersonic speed and finally completes a horizontal landing or attacks ground targets.

[0003] Numerous experts and scholars, both domestically and internationally, have conducted research on hypersonic vehicle guidance and attitude control methods, achieving significant results both in theoretical and engineering applications. Near-space glide flight guidance primarily encompasses standard trajectory guidance, predictive-correction guidance, and quasi-balanced glide guidance. Dive guidance primarily encompasses two methods: dive guidance based on a three-dimensional target-vehicle relative motion model, and tracking guidance based on a preset standard trajectory. Unlike center-of-mass guidance, where the expected values of the controlled state variables must be determined based on different guidance strategies, center-of-mass attitude control requires no design of expected values or expected dynamic variations. The core of this design lies in solving multi-input, multi-output, high-order nonlinear control systems. Consequently, current research focuses on innovative applications of control theory and methods, as well as the implementation of multi-actuator composite control. These approaches include nonlinear dynamic inversion control, sliding-mode variable structure control, model predictive control, and attitude control methods based on disturbance estimation.

[0004] Hypersonic vehicles exhibit significant characteristics during flight, including rapid time-varying, strong coupling, nonlinearity, and uncertainty. High Mach number flight results in dramatic changes in vehicle states, intensifying the interaction and coupling between the guidance and attitude control subsystems. Aerodynamic forces and aerodynamic torque calculations are closely related to the state quantities of the guidance and attitude control systems. The nonlinearity of the dynamic and motion models is enhanced, and the uncertainty of the flight environment and aerodynamic model is significant, along with the presence of unknown disturbances. Therefore, the design of the guidance and control systems for hypersonic vehicles cannot simply follow the traditional separation of guidance and control systems used in ballistic vehicles. Instead, it is necessary to fully consider the coupling characteristics of the guidance and attitude control systems and the interaction between the center of mass motion and the motion around the center of mass. This approach should integrate the guidance and attitude control subsystems to avoid redundant tuning and costly adjustments. However, relatively little research has been conducted on six-degree-of-freedom guidance and control throughout the reentry phase, comprehensively considering both guidance and attitude control. Therefore, a joint design approach for guidance and control throughout the reentry phase is urgently needed, integrating the guidance and attitude control loops to improve vehicle performance and reduce design costs. Summary of the Invention

[0005] In view of the above shortcomings of the prior art, the present invention proposes a combined method for full-range guidance and control of a trans-atmospheric reentry vehicle, which is used to accurately predict the temperature distribution in the formation-wellbore system.

[0006] The technical solution adopted by the present invention comprises the following steps:

[0007] Step 1: Construct a six-degree-of-freedom motion model of the trans-atmospheric re-entry vehicle, and use the established six-degree-of-freedom motion model as a simulation model for the six-degree-of-freedom guidance and attitude control numerical simulation analysis of the trans-atmospheric re-entry vehicle in different flight phases;

[0008] Step 2: Establish auxiliary solution equations for solving some parameters in the simulation model established in step 1;

[0009] Step 3: Define the constraints for the auxiliary solution equations established in step 2;

[0010] Step 4: Solve the parameters of the simulation model based on the constraints defined in step 3 and the auxiliary solution equations in step 2, and perform joint guidance and control design for the trans-atmospheric reentry vehicle in different flight segments based on the obtained parameters.

[0011] Furthermore, the six-degree-of-freedom motion model established in step 1 includes:

[0012] 1) Center of mass dynamics model of a trans-atmospheric reentry vehicle in a ballistic coordinate system

[0013]

[0014] Among them, V is the velocity, θ is the velocity inclination, and σ is the track yaw angle. The three are calculated based on the ground coordinate system; g xh ,g yh ,g zh is the component of gravitational acceleration in the three axes of the ballistic coordinate system, R xh ,R yh ,R zh is the component of aerodynamic acceleration in the three axes of the ballistic coordinate system, a ex ,a ey ,a ez is the centrifugal inertial acceleration -ω e ×(ω e ×r) in the ballistic coordinate system, a kx ,a ky ,a kz is the Coriolis acceleration term -2ω e ×δr / δt components in the three axes of the ballistic coordinate system;

[0015] 2) Kinematic model of the center of mass of a trans-atmospheric reentry vehicle in the ground coordinate system

[0016]

[0017] Where X, Y, and Z are the three-axis components of the aircraft position in the ground coordinate system.

[0018] 3) Dynamic model of the trans-atmospheric reentry vehicle around the center of mass in the body coordinate system

[0019]

[0020] Where, is the derivative of the inertia tensor around the three axes of the system; ω Ix ,ω Iy ,ω Iz is the component of the absolute angular rate of rotation of the aircraft relative to the inertial system in the three axes of the body coordinate system; ρ is the atmospheric density; S ref is the aerodynamic reference area; L x , L y , L z is the component of lift acceleration L in the three axes of the body coordinate system; m x , m y , m z are the components of the vehicle mass m in the three axes of the body coordinate system;

[0021] 4) Kinematic model of the trans-atmospheric reentry vehicle around the center of mass in the body coordinate system

[0022]

[0023] Where, ψ,γ are the pitch, yaw, and roll angles of the aircraft calibrated relative to the ground coordinate system.

[0024] Furthermore, the auxiliary solution equations established in step 2 include:

[0025] 1) The pitch, yaw, and roll angles of the aircraft attitude control loop, the velocity inclination and track yaw angle of the guidance loop, and the angle of attack, sideslip, and bank angle of the aerodynamically generated airflow angle are calculated as follows:

[0026]

[0027] Where, is the pitch angle, ψ is the yaw angle, γ is the roll angle, α is the angle of attack, β is the sideslip angle, υ is the roll angle, θ is the velocity inclination angle, and σ is the track yaw angle;

[0028] 2) The calculation formula of the geocentric radius r is:

[0029]

[0030] Where X, Y, and Z are the components of the aircraft's position in the three axes of the ground coordinate system, and R0 is the distance from the center of the earth at the intersection of the earth's center vector and the earth's surface at the initial moment;

[0031] 3) The calculation formula for geocentric latitude φ is:

[0032]

[0033] Among them, φ0 is the geocentric latitude of the launch point, and A0 is the launch azimuth;

[0034] 4) The calculation formula for the distance R from the center of the earth at the intersection of the center of the earth's radius vector and the earth's surface is:

[0035]

[0036] Where φ is the local latitude, a e is the semi-major axis of the Earth ellipsoid, b e is the semi-minor axis of the Earth ellipsoid;

[0037] 5) When the height of the aircraft from the ground is h = rR, the calculation formula for the geographical longitude λ at the intersection of the aircraft's geocentric radius and the ground is:

[0038]

[0039] Where λ0 is the geographical longitude of the aircraft's location at the initial moment;

[0040] 6) The calculation formula of the local velocity inclination angle Θ is:

[0041]

[0042] Where Θ is the local velocity inclination, V x ,V y ,V z The velocity vector V = [V x ,V y ,V z ] T Component form in the ground coordinate system.

[0043] Furthermore, the constraints defined in step 3 include:

[0044] 1) Center of mass state constraint

[0045] ①Speed angle constraint

[0046]

[0047] Where δ is the threshold for entering the quasi-balanced gliding handover; dr / dV is the first-order derivative of the distance from the center of the earth to the flight speed; (dr / dV) QEGCis the slope of the height-velocity profile of the quasi-equilibrium gliding trajectory;

[0048] ② Center of mass state constraint formula during hypersonic gliding flight

[0049]

[0050] Where h f ,λ f ,φ f is the target point height, longitude and geocentric latitude, V f is the terminal velocity, θ max is the maximum velocity inclination angle, ψ max is the maximum flight bearing error;

[0051] ③ Constraints on the center of mass state at the end of the dive flight

[0052] λ→λ T ,φ→φ T ,V→V f ,R→0

[0053] Where λ T ,φ T are the longitude and latitude of the target point respectively, and R is the terminal hit error;

[0054] 2) State constraints around the center of mass

[0055] ① Upper and lower bounds of the Euler angles and three-channel angular rates around the center of mass

[0056]

[0057] In the formula, the subscript min represents the lower limit of the corresponding variable, and max represents the upper limit of the corresponding variable;

[0058] ② The upper and lower bounds of the rate of change of the attitude Euler angle and the amplitude and rate of change of the angle of attack, sideslip angle and roll angle

[0059]

[0060] ③ Constraints on aircraft pneumatic actuators

[0061]

[0062] Where, δ e is the left elevator deflection, δ a is the right elevator deflection, δ r is the rudder deflection; the subscript min represents the minimum value of the variable, and the subscript max represents the maximum value of the variable;

[0063] 3) Process constraints

[0064] ① Constraints on heat flux, dynamic pressure and overload of trans-atmospheric reentry vehicles

[0065]

[0066] Where H(V) is the feasible altitude-speed flight profile of the aircraft, ρ0 is the reference ground atmospheric density, and h s =7110m, g0 is the standard acceleration of gravity;

[0067] The lower boundary H of the corresponding aircraft altitude-speed flight corridor Low (V) is:

[0068]

[0069] ② Attack angle constraint formula for the dive flight phase

[0070]

[0071] Where, α is the guidance instruction of the attack angle of the gliding flight phase, α pcon is the attack angle instruction corresponding to the process constraint, N max is the total overload upper limit, and g is the acceleration due to gravity.

[0072] Furthermore, the specific steps of step 4 include:

[0073] Step 4.1: Convert the heat flux, dynamic pressure, and total overload constraints into guidance command constraints for the control variables angle of attack and roll angle. Combined with the constraints around the center of mass, the guidance command for the angle of attack α in the gliding flight phase is obtained:

[0074]

[0075] Step 4.2: Based on the constraints and the established 6-DOF model, conduct joint design of guidance and control for the reentry glide flight phase.

[0076] Step 4.3: Based on the constraints and the established six-degree-of-freedom model, conduct joint design of guidance and control for the dive flight phase.

[0077] Furthermore, the specific steps of step 4.2 include:

[0078] Step 4.2.1: Design the roll angle command υ that satisfies longitudinal range control and azimuth error control C , perform re-entry glide flight center of mass guidance, bank angle command υ C for

[0079]

[0080] in, is the position sight angle ηLOS The first derivative of K1, K 01 ,K 02 is the parameter to be designed, θ ref is the reference velocity inclination value, is the sliding surface reaching law, θ g is the term related to the earth's gravity in the velocity and inclination differential equation, θ ωe is the term related to the earth's rotation in the velocity inclination differential equation, △σ is the lateral azimuth error, σ g ,σ ωe is the term related to the earth's gravity and rotation angular rate in the differential equation of the track yaw angle;

[0081] Step 4.2.2: Design the left and right elevator and rudder angle commands to perform backstepping attitude control around the center of mass during reentry glide. The right elevator, left elevator, and rudder angle commands are:

[0082]

[0083] Where, δ ac ,δ ec ,δ rc are the right elevator, left elevator and rudder angle commands; ω xc ,ω yc ,ω zc are virtual roll, yaw and pitch rate commands.

[0084] Furthermore, the specific steps of step 4.3 include:

[0085] Step 4.3.1: Design the desired roll angle command υ required for the vehicle to hit the target during the dive phase c and the aerodynamic lift command L c , to conduct the diving and downward pressure flight phase guidance, the υ c and L c Obtained by:

[0086]

[0087] Among them, K1,K2∈R 2×2 is the control feedback coefficient matrix to be designed;

[0088] Step 4.3.2: Based on the guidance attack angle and bank angle commands for the trans-atmospheric reentry vehicle during its dive, the desired roll moment, yaw moment, and pitch moment are obtained through the dynamic inverse control method. The dynamic inverse attitude control design for the high-speed dive phase is performed. The roll moment, yaw moment, and pitch moment are:

[0089]

[0090] Where M xc ,M yc ,M zc is the desired rolling moment, yaw moment and pitching moment, K1∈R 3×3 is the feedback coefficient matrix to be designed, S 11 ,S 12 ,S 13 is the element of the scalar sliding surface vector S1, is the control coefficient to be designed.

[0091] Therefore, the present invention adopts the above-mentioned joint design method for guidance and control throughout the reentry flight phase, which has the following beneficial effects:

[0092] First, the present invention establishes universal high-precision center-of-mass dynamic equations, center-of-mass kinematic equations, center-of-mass dynamic equations, center-of-mass kinematic equations, and auxiliary solution equations for reentry trajectory parameters for the entire reentry flight of a trans-atmospheric reentry vehicle, providing a standard high-precision mathematical model for six-degree-of-freedom trajectory and guidance control simulation.

[0093] Second, the present invention establishes a universal mathematical representation model for the various complex constraints imposed on a trans-atmospheric reentry vehicle during its entire reentry flight. By formula conversion, the flight process constraints are converted into constraints on the value of the center of mass and the state parameters around the center of mass. The constraints on the value of the state parameters around the center of mass are designed in conjunction with the guidance and control tasks and integrated into the entire process of the joint design of the reentry flight guidance and control.

[0094] Third, the present invention flexibly uses sliding mode control, backstepping control, dynamic surface control and dynamic inverse control technologies to integrate the design of the center-of-mass guidance law and attitude controller around the center of mass in the re-entry gliding phase, the guidance law and attitude control law in the dive-down flight phase, and uses numerical methods to set appropriate handover point parameters to complete the joint design of guidance and control for re-entry flight gliding and high-speed dive-down.

[0095] In summary, the method proposed in the present invention fully considers the coupling characteristics of the aircraft guidance system and the attitude control system and the interaction between the center of mass motion and the motion around the center of mass, and integrates the guidance loop and the attitude control loop into an integrated design, thereby improving the performance of the aircraft and reducing the design cost.

[0096] The technical solution of the present invention is further described in detail below through the accompanying drawings and embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0097] Figure 1 These are the left and right elevator deflection and rudder deflection curves in the embodiment of the present invention.

[0098] Figure 2 : is the attack angle change curve in the embodiment of the present invention.

[0099] Figure 3 is a sideslip angle variation curve in an embodiment of the present invention.

[0100] Figure 4 is a roll angle variation curve in an embodiment of the present invention.

[0101] Figure 5 These are the pitch, yaw, and roll angle curves in the embodiment of the present invention.

[0102] Figure 6 1 is a three-channel angular rate curve in an embodiment of the present invention.

[0103] Figure 7 It is the track yaw angle, velocity inclination angle and velocity curve in the embodiment of the present invention.

[0104] Figure 8 It is the longitude, geocentric latitude and altitude curve in the embodiment of the present invention.

[0105] Figure 9 This is the full three-dimensional flight trajectory of the trans-atmospheric reentry vehicle in an embodiment of the present invention. DETAILED DESCRIPTION

[0106] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, these embodiments are only used to illustrate the present invention and are not intended to limit the scope of the present invention. In addition, it should be understood that after reading the content taught by the present invention, those skilled in the art will make various changes or modifications to the present invention, and these equivalent forms also fall within the scope defined by the appended claims of the application.

[0107] The present invention proposes a joint design method for the whole-process guidance and control of a reentry vehicle, which comprises the following steps:

[0108] Step 1: Construct a six-degree-of-freedom motion model of the trans-atmospheric re-entry vehicle, and based on it, establish a simulation model of the trans-atmospheric re-entry vehicle;

[0109] Step 2: Define the constraints during the re-entry flight of the trans-atmospheric re-entry vehicle;

[0110] Step 3: Based on the constraints defined in the step and the established simulation model, conduct joint design of guidance and control for different flight segments.

[0111] Specifically, the specific steps of step 1 to establish a six-degree-of-freedom motion model include:

[0112] Step 1.1: Establish the center-of-mass dynamic equations of the trans-atmospheric reentry vehicle in the ballistic coordinate system and the center-of-mass motion model in the ground coordinate system;

[0113] Step 1.2: Establish the dynamics and motion model around the center of mass in the aircraft body coordinate system.

[0114] The following is an introduction to the six-degree-of-freedom model of the trans-atmospheric reentry vehicle, which includes:

[0115] 1. Center of mass dynamics and kinematics model

[0116] In the semi-velocity coordinate system / ballistic coordinate system O B -X H Y H Z H The vector form of the dynamic equation of the trans-atmospheric reentry vehicle can be described as:

[0117]

[0118] Where r is the radius vector of the aircraft relative to the center of the earth, g is the gravitational acceleration vector, R is the aerodynamic acceleration vector, ω e is the Earth's rotation angular velocity vector.

[0119] Decomposing the vector variables in the above equation into the ballistic coordinate system, we have:

[0120]

[0121] Among them, V is the velocity, θ is the velocity inclination, and σ is the track yaw angle;

[0122] Considering the J2 ellipsoid term, decompose the gravitational acceleration vector g along the direction of the earth's center radial vector and the direction of the earth's axis, and obtain the components of the gravitational acceleration in the three axes of the ground coordinate system:

[0123]

[0124] Where g x ,g y ,g z is the component of gravitational acceleration in the three axes of the ground coordinate system, X, Y, Z are the components of the aircraft position in the three axes of the ground coordinate system, φ0 is the geocentric latitude of the aircraft's ground projection point at the initial moment, A0 is the initial azimuth, r is the size of the geocentric radius vector, and R0 is the distance from the center of the earth at the intersection of the geocentric radius vector and the earth's surface at the initial moment;

[0125]

[0126] Where a e is the semi-major axis of the Earth ellipsoid, b e is the semi-minor axis of the Earth ellipsoid;

[0127] The component g of the gravitational acceleration vector g in the direction of the earth's center vector r and its component on the Earth's axis The specific form is:

[0128]

[0129] Where μ is the Earth's gravitational constant, r is the diameter of the Earth's center, φ is the current Earth's center latitude, and J2 = 1.08263 × 10 -3 is the J2 coefficient of the Earth ellipsoid model.

[0130] The conversion matrix T from the ground coordinate system to the velocity coordinate system disclosed in the prior art is used. VO And the transformation matrix T from the velocity coordinate system to the ballistic coordinate system HV The component of gravitational acceleration in the ballistic coordinate system can be obtained:

[0131]

[0132] Where g xh ,g yh ,g zh are the components of gravitational acceleration in the three axes of the ballistic coordinate system.

[0133] The aerodynamic acceleration R acting on the aircraft is decomposed into drag acceleration D, lift acceleration L and lateral acceleration N in the velocity coordinate system. Its components in the ballistic coordinate system are:

[0134]

[0135] Where R xh ,R yh ,R zh are the components of aerodynamic acceleration in the three axes of the ballistic coordinate system. The specific forms of D, L, and N are:

[0136]

[0137] Where ρ is the atmospheric density, V is the velocity, S ref is the aerodynamic reference area, C D ,C L ,C N are the aerodynamic drag, lift and side force coefficients, and m is the mass of the aircraft.

[0138] Centrifugal inertia acceleration-ω e ×(ω e ×r) in the ballistic coordinate system ex ,a ey ,a ez for:

[0139]

[0140] In the formula, the matrix element T ij , i=1,2,3,j=1,2,3, and the specific form of the matrix elements is:

[0141]

[0142] Among them, ω e is the angular velocity of the Earth's rotation, φ0 is the geocentric latitude of the launch point, and A0 is the launch azimuth;

[0143] Coriolis acceleration term -2ω e ×δr / δt components a in the three axes of the ballistic coordinate system kx ,a ky ,a kz for:

[0144]

[0145] The combined equations (1)-(11) can be used to calibrate the center of mass dynamics model of velocity, velocity inclination and track yaw angle based on the ground coordinate system:

[0146]

[0147] Formula (12) is the dynamic equation of the trans-atmospheric reentry vehicle in the ballistic coordinate system. In this formula, the velocity inclination angle and the track yaw angle are calculated based on the ground coordinate system. The velocity inclination angle is used to measure the direction of the vehicle's velocity relative to the ground coordinate system O. O X O Z O The tilt angle of the plane, the track yaw angle / heading angle is used to measure the deflection angle of the aircraft relative to the initial shooting plane.

[0148] Based on formula (12), the center of mass kinematic model of the trans-atmospheric reentry vehicle described in the ground coordinate system is obtained:

[0149]

[0150] Where X, Y, and Z are the three-axis components of the aircraft position in the ground coordinate system, V is the velocity, θ is the velocity inclination, and σ is the track yaw angle. The three components of the velocity vector are all calibrated relative to the ground coordinate system.

[0151] 2. Dynamic model and kinematic model around the center of mass

[0152] The dynamic equations around the center of mass are used to characterize the relationship between the three-channel angular rates of the aircraft and its resultant torque, and the motion model is used to characterize the relationship between the change in the attitude angle of the aircraft measured relative to the ground coordinate system and the three-channel angular rates.

[0153] The vector form of the center-of-mass dynamics model of the trans-atmospheric reentry vehicle is:

[0154]

[0155] Where J is the aircraft inertia matrix, ωI is the absolute angular velocity of the aircraft relative to the inertial coordinate system, M is the aerodynamic torque vector acting on the aircraft, and M k is the additional Coriolis moment term;

[0156] In order to simplify the problem analysis process, the vector form of the dynamic equation around the center of mass is projected into the aircraft body coordinate system.

[0157] It is known that the aircraft has an axisymmetric structure and its inertia product J xy =J xz =J yz =0, then (14) is equivalent to:

[0158]

[0159] Where, ω I is the angular velocity of the aircraft relative to the inertial system;

[0160] And the angular velocity of the aircraft relative to the ground coordinate system ω, the angular velocity of the earth's rotation ω e and ω I Satisfy between:

[0161] ω=ω I -ω e

[0162] Projecting Equation (15) into the aircraft body coordinate system, we obtain the center-of-mass dynamics model in this coordinate system:

[0163]

[0164] Where ρ is the atmospheric density, S ref is the aerodynamic reference area, ω Ix ,ω Iy ,ω Iz is the component of the absolute angular velocity of the aircraft relative to the inertial system in the three axes of the body coordinate system, is the derivative of the inertia tensor around the three axes of the body coordinate system; L x , L y , L z is the component of lift acceleration L in the three axes of the body coordinate system; m x , m y , m z are the components of the vehicle mass m in the three axes of the body coordinate system.

[0165] Equation (16) establishes the final trans-atmospheric reentry vehicle dynamic model around the center of mass, and Equation (17) is the kinematic model around the center of mass in the body coordinate system:

[0166]

[0167] Where, ψ,γ are the pitch, yaw, and roll angles of the aircraft calibrated relative to the ground coordinate system.

[0168] At this point, a dynamic and kinematic model of the trans-atmospheric reentry vehicle around the center of mass has been established. This model will be used in the design of the attitude control algorithm. At the same time, the center of mass dynamic model and the center of mass kinematic model constructed above are combined as simulation models for the numerical simulation analysis of the six-degree-of-freedom guidance and attitude control of the trans-atmospheric reentry vehicle in different flight phases. Some parameters in this simulation model need to be calculated by the following auxiliary solution equations.

[0169] 3. Assisted equation solving

[0170] The pitch angle, yaw angle, and roll angle of the aircraft attitude control loop, the velocity inclination angle and track yaw angle of the guidance loop, and the angle of attack, sideslip angle, and bank angle of the aerodynamic force-generated airflow angle satisfy the following equations:

[0171]

[0172] Where, is the pitch angle, ψ is the yaw angle, γ is the roll angle, α is the angle of attack, β is the sideslip angle, υ is the roll angle, θ is the velocity inclination angle, and σ is the track yaw angle;

[0173] The calculation formula of the geocentric radius r is:

[0174]

[0175] Where X, Y, and Z are the three-axis components of the aircraft position in the ground coordinate system;

[0176] The calculation formula for geocentric latitude φ is:

[0177]

[0178] The calculation formula for the distance R from the center of the earth at the intersection of the earth's center vector and the earth's surface is:

[0179]

[0180] Where φ is the local latitude;

[0181] The height of the aircraft from the ground is h = rR, and the calculation formula for the geographical longitude λ at the intersection of the aircraft's geocentric radius and the ground is:

[0182]

[0183] Where λ0 is the geographical longitude of the aircraft's location at the initial moment;

[0184] If the velocity vector in the ground coordinate system is in the form of V = [V x ,V y ,V z ] T , then the calculation formula of the local velocity inclination angle Θ is:

[0185]

[0186] Where Θ is the local velocity inclination angle.

[0187] 4. Constraints on the re-entry process of a trans-atmospheric re-entry vehicle

[0188] During the re-entry flight of a trans-atmospheric re-entry vehicle, there are many constraints that require constraints on the values of the auxiliary solution equations established above. These constraints include aerodynamic and thermal constraints, control capability constraints, guidance task constraints, and flight state constraints around the center of mass. The above constraints are divided into the following three categories:

[0189] (1) Center of mass state constraint

[0190] The center of mass state constraint is closely related to the aircraft guidance mission. Different flight phases correspond to different guidance missions and, at the same time, different center of mass state constraints. The guidance mission of the initial descent flight segment is to accurately enter the quasi-equilibrium glide condition, that is, to find the quasi-equilibrium glide matching value of altitude and speed. The terminal velocity, track yaw angle, and terminal position of this flight segment are not constrained. Only the velocity inclination angle is required to satisfy the following formula:

[0191]

[0192] Where δ is the threshold for entering the quasi-balanced gliding shift, which is usually a small positive constant; dr / dV is the first-order derivative of the distance from the center of the earth with respect to the flight speed, obtained according to the center of mass dynamics equation; (dr / dV) QEGC The slope of the altitude-velocity profile of the quasi-equilibrium glide trajectory can be obtained by taking the first-order derivative with respect to velocity on both sides of the equation dθ / dt=0.

[0193] For the hypersonic gliding flight phase, the center of mass guidance task is to guide the aircraft to a specific target area. Its center of mass state constraint can be expressed as:

[0194]

[0195] Where h f ,λ f ,φ f is the target point height, longitude and geocentric latitude, V f is the terminal velocity, θ max is the maximum velocity inclination angle, ψ maxis the maximum flight orientation error. This formula can be used as the most comprehensive guidance task description for the gliding flight phase, and its respective constraint values are given from the perspective of the center of mass state.

[0196] The core of the guidance during the dive phase is to hit the preset target. For the dive phase, the center of mass state constraints at the terminal (i.e., the end of the flight) can be summarized as follows:

[0197] λ→λ T ,φ→φ T ,V→V f ,R→0 (25)

[0198] Where λ T ,φ T is the longitude and latitude of the target point, V f is the terminal velocity (which may not be controlled), and R is the terminal hit error (referring to the relative distance between the aircraft and the target at the end of the flight, that is, the hit accuracy of the strike). The simulation ends when the aircraft touches the ground. At this time, R is the main guidance indicator of concern.

[0199] (2) State constraints around the center of mass

[0200] The state around the center of mass is related to the attitude stability of the aircraft, and is directly related to the spatial generation of the angle of attack, sideslip angle, and roll angle. In order to ensure the stability and controllability of the aircraft around the center of mass, the stability is usually ensured by setting the upper and lower limits of the Euler angles around the center of mass and the three-channel angular rate. The Euler angles of the center of mass attitude (pitch angle) The upper and lower limits of the yaw angle ψ, roll angle γ) and the three-channel angular rate are:

[0201]

[0202] In the formula, the subscript min represents the lower limit of the corresponding variable, and max represents the upper limit of the corresponding variable.

[0203] At the same time, set the change rate of the attitude Euler angle (i.e. the first-order derivative of each attitude angle) and the upper and lower bounds of the amplitude and change rate of the angle of attack, sideslip angle and roll angle:

[0204]

[0205] The above two equations limit the amplitude and angle change rate of the attitude Euler angle and airflow angle respectively, and also limit the angular rate of the three channels.

[0206] The amplitude and speed limits of the attitude control variables around the center of mass are mainly used to ensure that the aircraft can complete the tracking of the guidance instructions in a smooth attitude change process, and to ensure the attitude stability of the aircraft from both quantitative and qualitative perspectives. The constraints of the aircraft's pneumatic actuators are as follows:

[0207]

[0208] Where, δ e is the left elevator deflection, δ a is the right elevator deflection, δ r is the rudder deflection; the subscript min represents the minimum value of the variable, and the subscript max represents the maximum value of the variable.

[0209] (3) Process constraints

[0210] During the flight of a trans-atmospheric re-entry vehicle, it is subject to constraints such as dynamic pressure, heat flow, and overload. Dynamic pressure and heat flow are directly related to the vehicle material, while overload is related to parameters such as the vehicle control system and the vehicle fuselage. Process constraints are actually an equivalent description of the maximum harsh flight environment that the vehicle can withstand. The main constraint in the initial descent phase is the heat flux density constraint. At this time, the vehicle has a very high speed, and the vehicle should maintain the maximum angle of attack to minimize the speed and reduce the heat flow. At this time, the aircraft's roll angle command can be set to a constant value. During long-term gliding flight, the heat flow, dynamic pressure, and overload constraints need to be considered comprehensively. If the maximum allowable heat flux density, dynamic pressure, and overload of the vehicle are expressed as q max N max , then the following formula holds:

[0211]

[0212] Where H(V) is the feasible altitude-speed flight profile of the aircraft determined by multiple process constraints, ρ0 is the reference ground atmospheric density, and h s =7110m, g0 is the standard acceleration of gravity.

[0213] The above formula converts the stagnation point heat flux density, dynamic pressure and total overload constraints into their corresponding height boundary values. The lower boundary H of the corresponding aircraft height-speed flight corridor is Low (V) The specific form can be expressed as:

[0214]

[0215] At the same time, the upper bound of the aircraft's altitude-velocity profile is generally the equilibrium glide altitude corresponding to a constant bank angle and a preset angle of attack profile. However, the guidance and control system for the equilibrium glide phase is designed based on quasi-equilibrium glide conditions, so the upper bound of altitude is not considered here.

[0216] From the balanced gliding condition dθ / dt=0, we know that in the quasi-balanced gliding flight mode, when the bank angle command υ* is given, only two of the three variables, the distance from the center of the earth r, the speed V, and the angle of attack α, are independent. Therefore, at any given speed V during the gliding flight, the distance from the center of the earth and the angle of attack are one-to-one corresponding. We can then use the lower boundary of the altitude-speed corridor H Low (V) Get the lower boundary of the aircraft's angle of attack α Low (V). If the flight angle of attack of the aircraft is guaranteed to satisfy α≥α Low (V), it can ensure that the aircraft altitude is greater than the lower boundary of the altitude corresponding to the process constraint condition, thereby satisfying various process constraints.

[0217] The stagnation point heat flux, dynamic pressure, and total overload constraints during the gliding flight process are converted into guidance command constraints for the control variables angle of attack and roll angle. Combined with the constraints around the center of mass state, the guidance command for the angle of attack during the gliding flight phase can be finally expressed as:

[0218]

[0219] Where α* is the guidance instruction of the angle of attack obtained by the guidance solution. Through the processing of formula (31), the center of mass state of the aircraft is correspondingly restricted, thereby meeting the requirements of stagnation point heat flux density, dynamic pressure and total overload.

[0220] The dive flight phase mainly realizes ballistic downward pressure under ultra-high-speed flight conditions. The main constraint of this flight process is the overload constraint. The following formula is used to convert the overload constraint into the angle of attack constraint:

[0221]

[0222] Where, α pcon is the attack angle instruction corresponding to the process constraint, N max is the total overload limit, and g is the acceleration due to gravity. Considering that the angle of attack determines the generation of aerodynamic forces, while the roll angle only distributes aerodynamic forces, only the angle of attack of the aircraft is constrained here.

[0223] 5. Joint design of guidance and control for different flight phases of a trans-atmospheric reentry vehicle

[0224] Next, the guidance and control are jointly designed for the re-entry gliding flight phase and the dive flight phase of the aircraft.

[0225] (1) Joint design of guidance and control for the reentry glide phase

[0226] ① Reentry glide flight center of mass guidance

[0227] Ignoring altitude differences, the actual great-circle distance between the aircraft and the target can be calculated using the following formula using the current latitude and longitude of the aircraft and the latitude and longitude of the target point:

[0228] L real =acos(sinφ T sinφ+cosφ T cosφcos(λ T -λ)) (33)

[0229] Where, L real is the actual great circle distance between the aircraft and the target point.

[0230] The reference speed and inclination angle that enables the aircraft to reach the preset longitude and latitude is obtained by the following formula:

[0231]

[0232] Where θ ref is the reference speed inclination value.

[0233] In the longitudinal plane, by tracking θ ref To achieve the goal of controlling the flight range of the aircraft, the following sliding surface is designed:

[0234] S0=K 01 (θ-θ ref )+K 02 ∫(θ-θ ref )dt (35)

[0235] Where S0 is the sliding surface, K 01 ,K 02 are the parameters to be designed.

[0236] Select the sliding surface reaching law shown below

[0237]

[0238] Where K 03 ,β0,ν0 are the parameters to be designed, |·| is the absolute value function, and sgn(·) is the sign function.

[0239] Combined with the differential formula of aircraft velocity and inclination angle, the following formula can be obtained:

[0240]

[0241] Where q = 0.5ρV 2 is the dynamic pressure, ρ is the atmospheric density, S ref is the aerodynamic reference area, m is the mass of the aircraft, C L.C ,υ C are the desired aerodynamic lift coefficient and bank angle, θ g is the term related to the earth's gravity in the velocity and inclination differential formula, and its specific form is:

[0242]

[0243] Where μ is the Earth's gravitational coefficient, J = 1.5J2, J2 = 1.08263×10 -3 is the band harmonic coefficient, a e is the semi-major axis of the Earth's ellipsoid.

[0244] θ in formula (37) ωe is the term related to the Earth's rotation in the velocity and inclination differential equation, and its specific form is:

[0245]

[0246] Where, ω e is the angular rate of the Earth's rotation.

[0247] The lateral plane roll angle command is generated by using the azimuth error control method. Position sight angle η LOS The specific form is:

[0248]

[0249] Where λ T ,φ T is the longitude and latitude of the target point.

[0250] The basic goal of lateral control is to make the lateral position error △σ gradually approach zero. If the control method shown is used, the following formula is established:

[0251]

[0252] Where K1 is the parameter to be designed, the first-order derivative of the position sight angle is for:

[0253]

[0254] The specific form of F1 is:

[0255]

[0256] based on The differential equation of the track yaw angle σ in equation (12) can be obtained as follows:

[0257]

[0258] Where σ g ,σ ωe are the terms related to the earth's gravity and rotation angular rate in the differential equation of the track yaw angle, and their specific forms are:

[0259]

[0260] Based on equations (44) and (37), the roll angle command υ that satisfies longitudinal range control and azimuth error control is obtained C :

[0261]

[0262] At this point, the angle of attack and roll angle commands can be calculated.

[0263] ② Reentry glide backstepping attitude control around the center of mass

[0264] The goal of the attitude controller is to design left and right elevator and rudder angle commands to ensure that the attitude control subsystem can track the guidance commands as required and ensure that the various state variables of the aircraft around the center of mass are stable and controllable.

[0265] The guidance loop can provide the angle of attack command α c , sideslip angle command β c and the roll angle command υ c , where the roll angle command υ is obtained from equation (46) c , substitute it into equation (37) or equation (44) to obtain the angle of attack command α c , sideslip angle command β c is zero;

[0266] Using the airflow angle commands (i.e., angle of attack, sideslip angle, and bank angle commands) provided by the guidance loop, the following sliding surface vector is designed:

[0267]

[0268] Where, e α 、e β 、e υ They represent the tracking errors of variables α, β, and υ respectively, and S represents the sliding surface;

[0269] It should be noted that the exponential approaching law has a varying approaching speed during the approaching process, which has the advantage of both accelerating the approaching speed and weakening chattering. Combining equations (17) and (47), we obtain:

[0270]

[0271] Where K 01 ,K 02 ∈R 3×3 is the feedback coefficient matrix to be designed, sign(·) is the sign function, is the first-order time derivative of the angle of attack, sideslip angle and roll angle commands, ω xc ,ω yc ,ωzc are virtual roll, yaw and pitch rate commands.

[0272] The following sliding surface vectors are designed using three-channel virtual control instructions:

[0273]

[0274] Where S1∈R 3×1 is the sliding surface vector, which is also the tracking error vector of the three-channel angular rate;

[0275] Using the same sliding mode reaching law as airflow angle tracking:

[0276]

[0277] Where K 11 ,K 12 ∈R 3×3 is the feedback coefficient matrix to be designed.

[0278] Based on formula (50), the desired rudder angle command is obtained as:

[0279]

[0280] Where, δ ac ,δ ec ,δ rc Right elevator, left elevator and rudder angle commands.

[0281] (2) Joint design of guidance and control for the dive flight phase

[0282] ① Guidance during the dive and downward flight phase

[0283] It is known that the transfer matrix from the ground coordinate system to the line of sight coordinate system is T SO , the subscript SO indicates the transfer from the ground coordinate system to the line of sight coordinate system, O represents the ground coordinate system, and S represents the line of sight coordinate system. The transfer matrix from the velocity coordinate system to the ground coordinate system is T OV , subscript V represents the velocity coordinate system. Acceleration vector a HV The components in the line of sight coordinate system and the components in the velocity coordinate system satisfy the following formula:

[0284]

[0285] In the formula is the component of the aircraft acceleration in the line of sight coordinate system, is the component of the aircraft acceleration in the velocity coordinate system;

[0286] According to the previous analysis, the It is not used as a control input for dive guidance. At this time:

[0287]

[0288] Based on formula (53), the center of mass dynamics equation of the aircraft in the velocity coordinate system is obtained by substituting formula (53) into the center of mass dynamics model. According to this equation, we can know:

[0289]

[0290] By controlling the angular rate of the line of sight inclination and line of sight deflection of the dive relative motion to gradually approach zero, the trans-atmospheric reentry vehicle can be guided to fly towards the preset target. To complete this control task, the following sliding surface vectors are designed:

[0291]

[0292] Where S∈R 2×1 The sliding mode surface vector of the line of sight angular rate for dive guidance is controlled by using the reaching law of the sliding mode index shown in the following formula:

[0293]

[0294] Where K1, K2∈R 2×2 is the control feedback coefficient matrix to be designed.

[0295] Combining equations (54) and (56), we can obtain:

[0296]

[0297] The above formula can be used to obtain the desired angle of attack command, roll angle command, and aerodynamic lift command L required for the aircraft to hit the target. c .

[0298] ② Dynamic reverse attitude control design during high-speed dive

[0299] Based on the guidance attack angle and bank angle commands of the trans-atmospheric reentry vehicle dive flight calculated by equation (57), the dynamic inverse control method is used to complete the design of the dive phase attitude control system.

[0300] Based on the idea of dynamic inversion attitude control method, the differential formula of the three-channel angular rate of the aircraft is combined, that is, formula (17), and we can get:

[0301]

[0302] By designing the desired pitch, yaw, and roll angle commands, the required torque required to track the preset attitude angle command can be obtained using the above formula.

[0303] First, the airflow angle command can be converted into the attitude angle command of the aircraft through formula (18). Then, the pitch, yaw and roll angle commands obtained by the solution are used to design the sliding surface vector of the following formula:

[0304]

[0305] Where S1∈R 3×1 is the sliding surface vector, S 11 ,S 12 ,S 13 is the element of the scalar sliding mode face, is the control coefficient to be designed; e ψ ,e γ are the tracking errors of pitch, yaw and roll angles, is the first-order time derivative of the Euler angle tracking error of the aircraft attitude.

[0306] Taking into account the fast convergence and easy implementation of the sliding surface vector tracking error, the sliding surface vector reaching law shown in the following formula is selected:

[0307]

[0308] Where K1∈R 3×3 is the feedback coefficient matrix to be designed, κ1,κ2,κ3>0,ν1,ν2,ν3>0 are the attitude controller coefficients to be designed, and sgn(·) is the sign function.

[0309] In summary, we can get:

[0310]

[0311] Where M xc ,M yc ,M zc are the desired rolling, yaw, and pitching moments.

[0312] Example

[0313] To more systematically analyze the effectiveness of the guidance and control separation design approach for the guidance and attitude control technology of a trans-atmospheric reentry vehicle (TRAV), and to more realistically demonstrate the impact of the guidance and control separation design on the TRAV, this section conducts a comprehensive guidance and control joint simulation throughout the reentry flight. A typical TRAV reentry flight scenario is set up, and a six-degree-of-freedom guidance and control numerical simulation is conducted. A TRAV can utilize a combined air-breathing engine and rocket motor hybrid system to achieve a preset altitude and speed in a horizontal takeoff mode, or it can utilize a rocket boost for a vertical launch to a preset altitude and speed. The vehicle then reenters at a speed of approximately Mach 20. This long-duration, long-range reentry glide flight can serve as the primary flight phase for a high-lift-to-drag ratio gliding attack warhead, as well as a typical flight phase for a long-distance rapid Earth transportation system.

[0314] The reentry process of a trans-atmospheric reentry vehicle can be divided into three flight phases with different ballistic characteristics: an initial descent adjustment phase, a long glide phase, and a dive phase. The initial descent phase is the unpowered reentry phase after the vehicle's reentry attitude adjustment is completed, and the vehicle relies on the vehicle's control actuators to complete a quasi-balanced glide phase. This phase covers a longitudinal range of hundreds of kilometers. From its starting altitude, the vehicle gradually descends into the dense atmosphere where it can generate sufficient lift to offset the effects of longitudinal gravity, completing the initial descent phase and the quasi-balanced glide phase. The near-space hypersonic glide phase is the main flight phase. During this phase, the vehicle relies on its high lift-to-drag ratio aerodynamic characteristics to achieve force balance in the longitudinal plane, maintaining near-constant altitude flight. The vehicle's attitude and guidance control tasks are completely dependent on the vehicle's aerodynamic control mechanisms. Due to the relatively low atmospheric density at near-space altitudes, the vehicle's speed loss during the glide phase is relatively limited. This characteristic also facilitates the vehicle's range extension, making it very feasible for strategic weapons with a range of 20,000 kilometers or circumnavigation of the Earth. The dive / landing phase is the final flight phase. As a weapon, it's used to execute a dive attack, and as a transport system, it's used to complete a horizontal landing. For vehicles with a high lift-to-drag ratio, the vehicle's trajectory is lowered primarily by the ground-facing aerodynamic lift generated by a 180-degree bank, and terminal velocity is typically limited. The above only qualitatively describes the trajectory characteristics of a trans-atmospheric reentry vehicle throughout its reentry flight. Below are numerical simulation results for the six-degree-of-freedom center of mass and around the center of mass.

[0315] The left and right elevator and rudder deflections of the aircraft are as follows Figure 1As shown, during the initial descent phase, the attitude control system tracks the constant angle of attack and bank angle commands. However, the atmospheric density is very low at this time, and the aerodynamic torque of the aircraft is very limited. Although the aircraft has used the maximum amplitude of the elevator for control, the tracking of the command is still not effective in the absence of insufficient dynamic pressure. As the altitude decreases, the atmospheric density increases, and the aerodynamic control capability of the aircraft becomes stronger and stronger. After the aircraft completes the precise adaptation of the six degrees of freedom of quasi-balanced gliding at a suitable flight altitude, the aerodynamic control mechanism of the aircraft enters a smooth stage until the aircraft enters the dive flight stage. During the entire reentry flight, the rudder deflection angle meets the amplitude limit and the rudder deflection rate limit.

[0316] The changes of angle of attack, sideslip angle and roll angle are as follows: Figure 2 、 Figure 3 and Figure 4 As shown. During the initial descent flight, the aircraft attitude controller cannot ideally track the guidance constant angle command due to insufficient dynamic pressure. However, the oscillation of the airflow angle is a phenomenon of finite-time oscillation and decrease. The initial descent flight segment ends when the difference between the velocity inclination angle change rate and the quasi-balanced gliding velocity change rate is less than the set constant value. The aircraft enters the quasi-balanced gliding guidance control phase. At the beginning of this phase, the changes in the aircraft's various state variables are relatively drastic. This is a normal phenomenon for the quasi-balanced gliding handover point, and it also reflects the aircraft's own attempt to accurately match the six-degree-of-freedom state variables of the quasi-balanced gliding handover conditions. After achieving the precise matching conditions, the aircraft enters a stable quasi-balanced gliding flight phase.

[0317] Pitch, yaw and roll angles and their angular rates change as follows Figure 5 and Figure 6 As shown in the figure, the attitude angle change is basically consistent with the trend of the angle of attack, sideslip angle and roll angle. Among the three-channel angular rates as the fastest variable, the roll angular rate has the largest change amplitude, because the aircraft needs to tilt and flip to distribute the aerodynamic force to ensure the force balance in the longitudinal plane. The track yaw angle, velocity tilt angle and velocity change curves are shown in the figure. Figure 7 As shown in the figure, the track yaw angle is related to the flight azimuth error, and its change is consistent with the goal of the azimuth error angle gradually approaching zero. The changes in the velocity inclination angle during the initial descent and quasi-balanced gliding are relatively small. After reaching the dive stage, the aircraft performs ballistic downward pressure to complete the dive attack mission, and a large negative value appears. The initial speed of the aircraft is 6000m / s, and the terminal speed is 2107m / s. During the ultra-high-speed flight, the aircraft can complete the entire flight mission well by relying only on the aerodynamic control mechanism. The changes in longitude, geocentric latitude and altitude, as well as the three-dimensional flight trajectory of the entire re-entry are shown as follows: Figure 8 and Figure 9 shown.

[0318] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention rather than to limit the same. Although the present invention has been described in detail with reference to the preferred embodiments, those skilled in the art should understand that they can still modify or replace the technical solutions of the present invention with equivalents, and these modifications or equivalent replacements cannot cause the modified technical solutions to deviate from the spirit and scope of the technical solutions of the present invention.

Claims

1. A joint design method for the full-range guidance and control of a trans-atmospheric reentry vehicle, characterized by: The following steps are involved: Step 1: Construct a six-degree-of-freedom motion model of the trans-atmospheric re-entry vehicle, and use the established six-degree-of-freedom motion model as a simulation model for the six-degree-of-freedom guidance and attitude control numerical simulation analysis of the trans-atmospheric re-entry vehicle in different flight phases; Step 2: Establish auxiliary solution equations for solving some parameters in the simulation model established in step 1; Step 3: Define the constraints for the auxiliary solution equations established in step 2; Step 4: Solve the parameters of the simulation model based on the constraints defined in step 3 and the auxiliary solution equations in step 2, and perform joint guidance and control design for the trans-atmospheric reentry vehicle in different flight segments based on the obtained parameters.

2. The method for joint design of full-range guidance and control of a trans-atmospheric reentry vehicle according to claim 1, characterized in that: The six-degree-of-freedom motion model established in step 1 includes: 1) Center of mass dynamics model of a trans-atmospheric reentry vehicle in a ballistic coordinate system Among them, V is the velocity, θ is the velocity inclination, and σ is the track yaw angle. The three are calculated based on the ground coordinate system; g xh ,g yh ,g zh is the component of gravitational acceleration in the three axes of the ballistic coordinate system, R xh ,R yh ,R zh is the component of aerodynamic acceleration in the three axes of the ballistic coordinate system, a ex ,a ey ,a ez is the centrifugal inertial acceleration -ω e ×(ω e ×r) in the ballistic coordinate system, a kx ,a ky ,a kz is the Coriolis acceleration term -2ω e ×δr / δt components in the three axes of the ballistic coordinate system; 2) Kinematic model of the center of mass of a trans-atmospheric reentry vehicle in the ground coordinate system Where X, Y, and Z are the three-axis components of the aircraft position in the ground coordinate system; 3) Dynamic model of the trans-atmospheric reentry vehicle around the center of mass in the body coordinate system Where, is the derivative of the inertia tensor around the three axes of the system; ω Ix ,ω Iy ,ω Iz is the component of the absolute angular rate of rotation of the aircraft relative to the inertial system in the three axes of the body coordinate system; ρ is the atmospheric density; S ref is the aerodynamic reference area; L x , L y , L z is the component of lift acceleration L in the three axes of the body coordinate system; m x , m y , m z are the components of the vehicle mass m in the three axes of the body coordinate system; 4) Kinematic model of the trans-atmospheric reentry vehicle around the center of mass in the body coordinate system Where, The pitch, yaw, and roll angles of the aircraft are calibrated relative to the ground coordinate system.

3. The method for joint design of full-range guidance and control of a trans-atmospheric reentry vehicle according to claim 1, characterized in that: The auxiliary solution equations established in step 2 include: 1) The pitch, yaw, and roll angles of the aircraft attitude control loop, the velocity inclination and track yaw angle of the guidance loop, and the angle of attack, sideslip, and bank angle of the aerodynamically generated airflow angle are calculated as follows: Where, is the pitch angle, ψ is the yaw angle, γ is the roll angle, α is the angle of attack, β is the sideslip angle, υ is the roll angle, θ is the velocity inclination angle, and σ is the track yaw angle; 2) The calculation formula of the geocentric radius r is: Where X, Y, and Z are the components of the aircraft's position in the three axes of the ground coordinate system, and R0 is the distance from the center of the earth at the intersection of the earth's center vector and the earth's surface at the initial moment; 3) The calculation formula for geocentric latitude φ is: Among them, φ0 is the geocentric latitude of the launch point, and A0 is the launch azimuth; 4) The calculation formula for the distance R from the center of the earth at the intersection of the center of the earth's radius vector and the earth's surface is: Where φ is the local latitude, a e is the semi-major axis of the Earth ellipsoid, b e is the semi-minor axis of the Earth ellipsoid; 5) When the height of the aircraft from the ground is h = rR, the calculation formula for the geographical longitude λ at the intersection of the aircraft's geocentric radius and the ground is: Where λ0 is the geographical longitude of the aircraft's location at the initial moment; 6) The calculation formula of the local velocity inclination angle Θ is: Where Θ is the local velocity inclination, V x ,V y ,V z The velocity vector V = [V x ,V y ,V z ] T Component form in the ground coordinate system.

4. The method for joint design of full-range guidance and control of a trans-atmospheric reentry vehicle according to claim 3, characterized in that: The constraints defined in step 3 include: 1) Center of mass state constraint ①Speed angle constraint Where δ is the threshold for entering the quasi-balanced gliding handover; dr / dV is the first-order derivative of the distance from the center of the earth to the flight speed; (dr / dV) QEGC is the slope of the height-velocity profile of the quasi-equilibrium gliding trajectory; ② Center of mass state constraint formula during hypersonic gliding flight Where h f ,λ f ,φ f is the target point height, longitude and geocentric latitude, V f is the terminal velocity, θ max is the maximum velocity inclination angle, ψ max is the maximum flight bearing error; ③ Constraints on the center of mass state at the end of the dive flight λ→λ T ,φ→φ T ,V→V f ,R→0 Where λ T ,φ T are the longitude and latitude of the target point respectively, and R is the terminal hit error; 2) State constraints around the center of mass ① Upper and lower bounds of the Euler angles and three-channel angular rates around the center of mass In the formula, the subscript min represents the lower limit of the corresponding variable, and max represents the upper limit of the corresponding variable; ② The upper and lower bounds of the rate of change of the attitude Euler angle and the amplitude and rate of change of the angle of attack, sideslip angle and roll angle ③ Constraints on aircraft pneumatic actuators Where, δ e is the left elevator deflection, δ a is the right elevator deflection, δ r is the rudder deflection; the subscript min represents the minimum value of the variable, and the subscript max represents the maximum value of the variable; 3) Process constraints ① Constraints on heat flux, dynamic pressure and overload of trans-atmospheric reentry vehicles Where H(V) is the feasible altitude-speed flight profile of the aircraft, ρ0 is the reference ground atmospheric density, and h s =7110m, g0 is the standard acceleration of gravity; The lower boundary H of the corresponding aircraft altitude-speed flight corridor Low (V) is: ② Attack angle constraint formula for the dive flight phase Where, α is the guidance instruction of the attack angle of the gliding flight phase, α pcon is the attack angle instruction corresponding to the process constraint, N max is the total overload upper limit, and g is the acceleration due to gravity.

5. The method for joint design of full-range guidance and control of a trans-atmospheric reentry vehicle according to claim 4, characterized in that: The specific steps of step 4 include: Step 4.1: Convert the heat flux, dynamic pressure, and total overload constraints into guidance command constraints for the control variables angle of attack and roll angle. Combined with the constraints around the center of mass, the guidance command for the angle of attack α in the gliding flight phase is obtained: Step 4.2: Based on the constraints and the established 6-DOF model, conduct joint design of guidance and control for the reentry glide flight phase. Step 4.3: Based on the constraints and the established six-degree-of-freedom model, conduct joint design of guidance and control for the dive flight phase.

6. The method for joint design of full-range guidance and control of a trans-atmospheric reentry vehicle according to claim 5, characterized in that: The specific steps of step 4.2 include: Step 4.2.1: Design the roll angle command υ that satisfies longitudinal range control and azimuth error control C , perform re-entry glide flight center of mass guidance, bank angle command υ C for in, is the position sight angle η LOS The first derivative of K1, K 01 ,K 02 is the parameter to be designed, θ ref is the reference velocity inclination value, is the sliding surface reaching law, θ g is the term related to the earth's gravity in the velocity and inclination differential equation, θ ωe is the term related to the earth's rotation in the velocity inclination differential equation, △σ is the lateral azimuth error, σ g ,σ ωe is the term related to the earth's gravity and rotation angular rate in the differential equation of the track yaw angle; Step 4.2.2: Design the left and right elevator and rudder angle commands to perform backstepping attitude control around the center of mass during reentry glide. The right elevator, left elevator, and rudder angle commands are: Where, δ ac ,δ ec ,δ rc are the right elevator, left elevator and rudder angle commands; ω xc ,ω yc ,ω zc are virtual roll, yaw and pitch rate commands.

7. The method for joint design of full-range guidance and control of a trans-atmospheric reentry vehicle according to claim 6, characterized in that: The specific steps of step 4.3 include: Step 4.3.1: Design the desired roll angle command υ required for the vehicle to hit the target during the dive phase c and the aerodynamic lift command L c , to conduct the diving and downward pressure flight phase guidance, the υ c and L c Obtained by: Among them, K1,K2∈R 2×2 is the control feedback coefficient matrix to be designed; Step 4.3.2: Based on the guidance attack angle and bank angle commands for the trans-atmospheric reentry vehicle during its dive, the desired roll moment, yaw moment, and pitch moment are obtained through the dynamic inverse control method. The dynamic inverse attitude control design for the high-speed dive phase is performed. The roll moment, yaw moment, and pitch moment are: Where M xc ,M yc ,M zc is the desired rolling moment, yaw moment and pitching moment, K1∈R 3×3 is the feedback coefficient matrix to be designed, S 11 ,S 12 ,S 13 is the element of the scalar sliding surface vector S1, is the control coefficient to be designed.

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