A method for estimating initial values ​​of online replanning of thrust descent during glide phase of launch vehicle

Abstract

The present application provides a method for estimating initial values ​​of online replanning of thrust reduction across a glide segment of a launch vehicle, the method comprising: dividing the trajectory of a launch vehicle under a thrust reduction failure into a first active segment, a glide segment, and a second active segment; determining the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment based on the flight state of the launch vehicle; and estimating initial values ​​of replanning according to the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment. The present application divides the trajectory of a launch vehicle under a thrust reduction failure into a first active segment, a glide segment, and a second active segment, and simultaneously considers the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment to estimate initial values ​​of replanning, making the estimation process more reasonable, thereby improving the convergence and rapidity of numerical replanning.

Description

Technical Field

[0001] The present application relates to the field of launch vehicle control technology, and in particular to a method for estimating initial values ​​of online re-planning of thrust descent across a glide phase of a launch vehicle. Background Art

[0002] The process from the launch vehicle igniting and taking off from the launch pad to sending the spacecraft into the predetermined orbit is divided into an active phase and a glide phase.

[0003] The active segment is the orbital segment during which the rocket engine is operating, and the gliding segment is the orbital segment that does not generate thrust after the engine is shut down, that is, the inertial flight segment.

[0004] With regard to the mission profile of a launch vehicle with a glide phase in the orbital stage, how to improve the fault adaptability has become the focus of current research. Summary of the invention

[0005] In order to solve one of the above-mentioned technical defects, the present application provides an initial value estimation method for online re-planning of thrust descent of a launch vehicle across the glide phase.

[0006] In a first aspect, the present application provides a method for estimating initial values ​​of online replanning of thrust drop across a glide phase of a launch vehicle, the method comprising:

[0007] The trajectory of the launch vehicle under thrust reduction failure condition is divided into the first active segment, the glide segment, and the second active segment;

[0008] Determining a flight state sequence of the first active segment, a flight state sequence of the gliding segment, and a flight state sequence of the second active segment based on the flight state of the launch vehicle;

[0009] The replanning initial value estimation is performed according to the flight state sequence of the first active segment, the flight state sequence of the gliding segment, and the flight state sequence of the second active segment.

[0010] Optionally, the pitch program angle and the yaw program angle of the first active segment are both constant values;

[0011] The apogee height of the glide segment is the perigee height of the original target orbit;

[0012] In the second active segment, the thrust direction of the launch vehicle is always perpendicular to the direction of the distance from the center of the earth in the orbital plane;

[0013] Wherein, the carrier rocket enters the second active section when gliding to the apogee.

[0014] Optionally, determining the flight state sequence of the first active segment based on the flight state of the launch vehicle includes:

[0015] Constructing a first active segment replanning problem; the first active segment replanning problem includes: a motion equation, an initial state, a terminal state and a control quantity constraint;

[0016] A nonlinear programming method is used to solve the re-planning problem of the first active segment to obtain a flight state sequence X1 and a control quantity sequence U1 of the first active segment;

[0017] Among them, U1 is composed of the control quantities at each discrete point of the first active segment, and the control quantities at each discrete point of the first active segment are

[0018] is the pitch program angle of the first active segment at the time of failure, ψ0 is the yaw program angle of the first active segment at the time of failure, [] T is the transpose operation.

[0019] Optionally, the motion equation is:

[0020]

[0021] T=I sp dm;

[0022]

[0023] Where, for the first-order derivative operator, P is the position vector, V is the velocity vector, T is the engine thrust after the failure, m is the mass, u is the thrust direction, μ is the constant coefficient of the earth's gravity, r is the distance from the center of mass of the rocket to the center of the earth, r is the component of the distance from the center of the earth in the launch inertial coordinate system, dm is the flow rate per second after the failure, I sp is the engine specific impulse, is the pitch program angle of the first active segment, and ψ is the yaw program angle of the first active segment.

[0024] Optionally, in the launch inertial coordinate system, the origin O is at the launch point, the OX axis points to the launch direction in the horizontal plane, the OY axis is perpendicular to the local horizontal plane of the launch point and points to the sky, and the OZ axis satisfies the right-hand rule.

[0025] Optionally, the initial state is:

[0026] [P,V,m] T (t0) = [P0, V0, m0] T ;

[0027] Among them, P is the position vector, V is the velocity vector, m is the mass, t0 is the fault time, P0 is the position vector at the fault time, V0 is the velocity vector at the fault time, and m0 is the mass at the fault time.

[0028] Optionally, the terminal state is:

[0029] [a f1 , e f1 ,i f1 ,Ω f1 ,w f1 ,f f1 ] T =Fun orbit (P(t f1 ),V(t f1 ));

[0030]

[0031]

[0032] |i f1 -i ref |≤ε i ,|Ω f1 -Ω ref |≤ε Ω ;

[0033] Among them, a f1 is the semi-major axis of the orbit at the terminal moment of the first active segment, e f1 is the orbital eccentricity at the terminal moment of the first active segment, i f1 is the orbital inclination at the terminal moment of the first active segment, Ω f1 is the longitude of the ascending node of the first active segment at the terminal time, w f1 is the orbital perigee argument at the terminal moment of the first active segment, f f1 is the true anomaly at the terminal moment of the first active segment, Fun orbit ( ) is the conversion function between the orbital elements of the first active segment and the position and velocity in the launch inertial coordinate system, t f1 is the terminal time of the first active segment, P(t f1 ) is the position vector of the first active segment at the terminal moment, V(t f1 ) is the velocity vector at the terminal moment of the first active segment, ha f1 is the orbital apogee height at the terminal moment, r B is the telecentricity of the sliding track, R e is the radius of the Earth, n T is the variable to be solved, m0 is the mass at the time of failure, m s is the structural mass, m load is the payload mass, V B is the perigee velocity of the original target orbit, μ is the constant coefficient of the earth's gravity, I sp is the engine specific impulse, i ref is the orbital inclination of the target orbit, εi is the maximum value of the orbital inclination deviation, Ω ref is the longitude of the ascending node of the target orbit, ε Ω is the maximum value of the longitude deviation of the ascending node.

[0034] Optionally,

[0035] Among them, r A is the pericentric distance of the sliding track.

[0036] Optionally, the control quantity constraint is:

[0037]

[0038] Wherein, t is any moment of the first active segment, is the pitch program angle of the first active segment at time t, and ψ(t) is the yaw program angle of the first active segment at time t.

[0039] Optionally, determining the flight state sequence of the glide segment based on the flight state of the launch vehicle includes:

[0040] Determining the number of orbital elements of the glide segment according to the terminal state of the launch vehicle at the moment of entering the glide orbit;

[0041] Determine the velocity vector of each discrete point of the sliding segment based on the number of track elements of the sliding segment and position vector Wherein, j is the identifier of the discrete point of the glide segment, j=1,…,N+1, N+1 is the total number of discrete points of the glide segment, and N is the number of intervals bisected by the true anomaly angle of the glide segment;

[0042] Determining the time of each discrete point of the sliding segment;

[0043] According to the velocity vector of each discrete point of the sliding segment and position vector The time of each discrete point of the glide segment obtains the flight state sequence X of the glide segment c and control quantity sequence U c ;

[0044] Among them, U c It is composed of control quantities at each discrete point of the sliding segment, and the control quantities at each discrete point of the sliding segment are all 0.

[0045] Optionally, the number of rails in the sliding section is

[0046] in, is the semi-major axis of the orbit of the glide segment, is the track eccentricity of the sliding segment, is the track inclination of the sliding segment, is the longitude of the orbit ascending node of the glide segment, is the orbital perigee argument of the glide segment, is the true anomaly of the orbit of the glide segment.

[0047] Optionally, the velocity vector of each discrete point in the sliding segment and position vector Satisfies the following relationship:

[0048]

[0049] Among them, Fun PV ( ) is the conversion function between the orbital elements of the glide segment and the position and velocity in the launch inertial coordinate system,

[0050] Optionally, the time of each discrete point in the sliding segment satisfies the following relationship:

[0051]

[0052]

[0053] Among them, t f1 is the terminal time of the first active segment, j′ is the non-first discrete point identifier of the sliding segment, j′=2,…,N+1, is the eccentric anomaly angle of the discrete point j′ of the sliding segment, and μ is the constant coefficient of the earth's gravity.

[0054] Optionally, determining the flight state sequence of the second active segment based on the flight state of the launch vehicle includes:

[0055] Determining the time interval and initial state of the second active segment according to the remaining flight time of the carrier rocket;

[0056] Based on the time interval and initial state of the second active segment, the flight state quantity sequence X2 and the control quantity sequence U2 of the second active segment are calculated by numerical integration.

[0057] Optionally, the time interval of the second active segment

[0058] Among them, t f2 is the remaining flight time of the launch vehicle, and M is the total number of discrete points in the second active segment - 1.

[0059] Optionally, the initial state of the second active segment is the terminal state of the gliding segment.

[0060] In a second aspect of the present application, an electronic device is provided, comprising:

[0061] Memory;

[0062] Processor; and

[0063] Computer programs;

[0064] The computer program is stored in the memory and is configured to be executed by the processor to implement the method as described in the first aspect above.

[0065] According to a third aspect of the present application, a computer-readable storage medium is provided, on which a computer program is stored; the computer program is executed by a processor to implement the method described in the first aspect above.

[0066] The present application provides a method for estimating initial values ​​of online replanning of thrust reduction across a glide segment of a launch vehicle, the method comprising: dividing the trajectory of a launch vehicle under a thrust reduction failure into a first active segment, a glide segment, and a second active segment; determining the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment based on the flight state of the launch vehicle; and estimating initial values ​​of replanning according to the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment. The present application divides the trajectory of a launch vehicle under a thrust reduction failure into a first active segment, a glide segment, and a second active segment, and simultaneously considers the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment to estimate initial values ​​of replanning, making the estimation process more reasonable, thereby improving the convergence and rapidity of numerical replanning. BRIEF DESCRIPTION OF THE DRAWINGS

[0067] The drawings described herein are used to provide a further understanding of the present application and constitute a part of the present application. The illustrative embodiments of the present application and their descriptions are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0068] Figure 1 A flow chart of an initial value estimation method for online re-planning of thrust drop across a glide phase of a launch vehicle provided in an embodiment of the present application;

[0069] Figure 2 A flowchart of another method for estimating initial values ​​of online re-planning of thrust descent across a glide phase of a launch vehicle provided in an embodiment of the present application. DETAILED DESCRIPTION

[0070] In order to make the technical solutions and advantages in the embodiments of the present application more clearly understood, the exemplary embodiments of the present application are further described in detail below in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present application, rather than an exhaustive list of all the embodiments. It should be noted that the embodiments in the present application and the features in the embodiments can be combined with each other without conflict.

[0071] In the process of realizing the present application, the inventors found that, for the flight mission profile of a launch vehicle with a glide segment in the orbital stage, how to improve the fault adaptability has become the focus of current research.

[0072] In view of the above problems, an embodiment of the present application provides an initial value estimation method for online replanning of thrust reduction across the glide segment of a launch vehicle, the method comprising: dividing the trajectory of the launch vehicle in the case of thrust reduction failure into a first active segment, a glide segment, and a second active segment; determining the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment based on the flight state of the launch vehicle; and performing replanning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment. The present application divides the trajectory of the launch vehicle in the case of thrust reduction failure into a first active segment, a glide segment, and a second active segment, and simultaneously considers the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment to perform replanning initial value estimation, so that the estimation process is more reasonable, thereby improving the convergence and rapidity of numerical replanning.

[0073] See also Figure 1 The implementation process of the method for estimating the initial value of the online re-planning of the thrust drop across the glide phase of a launch vehicle provided in this embodiment is as follows:

[0074] 101, the trajectory of the launch vehicle in the case of thrust reduction failure is divided into the first active segment, the glide segment, and the second active segment.

[0075] Among them, the pitch program angle and the yaw program angle of the first active segment are both constant values.

[0076] The apogee altitude of the glide segment is the perigee altitude of the original target orbit.

[0077] In the second active segment, the thrust direction of the launch vehicle is always perpendicular to the direction of the Earth's center in the orbital plane.

[0078] Moreover, the launch vehicle enters the second active phase when it glides to the apogee.

[0079] That is to say, in the method of this embodiment, it is assumed that the pitch and yaw program angles of the first active segment are constant values, the sliding track surface to be solved is close to the original sliding track surface, and the apogee height of the sliding segment is considered to be the perigee height of the original target track.

[0080] 102. Determine a flight state sequence of a first active segment, a flight state sequence of a glide segment, and a flight state sequence of a second active segment based on the flight state of the launch vehicle.

[0081] In this step, the flight state sequence of the first active segment is determined based on the flight state of the launch vehicle, the flight state sequence of the glide segment is determined based on the flight state of the launch vehicle, and the flight state sequence of the second active segment is determined based on the flight state of the launch vehicle.

[0082] 1. The implementation process of determining the flight state sequence of the first active segment based on the flight state of the launch vehicle

[0083] 1) Construct the first active segment replanning problem.

[0084] The replanning of the first active segment after a fault needs to consider the constraints of the motion equation, the initial state constraints at the time of the fault, the terminal constraints of the sliding track, and the control quantity constraints. Therefore, the replanning problem of the first active segment includes: motion equation, initial state, terminal state, and control quantity constraints.

[0085] (1) Equation of motion

[0086] That is, the motion equation constraint, which is the state equation describing the center of mass motion of the rocket in the launch inertial coordinate system. Among them, the origin O of the launch inertial coordinate system (abbreviated as launch inertial system) is at the launch point, the OX axis points to the launch direction in the horizontal plane, the OY axis points to the sky perpendicular to the local horizontal plane at the launch point, and the OZ axis satisfies the right-hand rule.

[0087] Specifically, the equation of motion is:

[0088]

[0089] T=I sp dm.

[0090]

[0091] Where, for the first-order derivative operator, P is the position vector, V is the velocity vector, T is the engine thrust after the failure, m is the mass, u is the thrust direction, μ is the constant coefficient of the earth's gravity, r is the distance from the center of mass of the rocket to the center of the earth, r is the component of the distance from the center of the earth in the launch inertial coordinate system, dm is the flow rate per second after the failure, I sp is the engine specific impulse, is the pitch program angle of the first active segment, and ψ is the yaw program angle of the first active segment.

[0092] Therefore, the state quantity is [P, V, m] T , the control quantity is

[0093] (2) Initial state

[0094] That is, the initial state constraint at the time of failure.

[0095] If the fault time is defined as t0, the corresponding initial state is [P0, V0, m0] T , then the initial state is:

[0096] [P,V,m] T (t0) = [P0, V0, m0] T .

[0097] Among them, P is the position vector, V is the velocity vector, m is the mass, P0 is the position vector at the time of fault, V0 is the velocity vector at the time of fault, and m0 is the mass at the time of fault.

[0098] (3) Terminal status

[0099] That is, the terminal constraint of the sliding track.

[0100] If the terminal time of the first active segment is defined as t f1 , the number of sliding track elements corresponding to the terminal velocity and position must satisfy the following constraints,

[0101] [a f1 ,e f1 ,i f1 ,Ω f1 ,w f1 ,f f1 ] T =Fun orbit (P(t f1 ),V(t f1 )).

[0102]

[0103] Among them, a f1 is the semi-major axis of the orbit at the terminal moment of the first active segment, e f1 is the orbital eccentricity at the terminal moment of the first active segment, i f1 is the orbital inclination at the terminal moment of the first active segment, Ω f1 is the longitude of the ascending node of the first active segment at the terminal time, w f1 is the orbital perigee argument at the terminal moment of the first active segment, f f1 is the true anomaly at the terminal moment of the first active segment, Fun orbit ( ) is the conversion function between the orbital elements of the first active segment and the position and velocity in the launch inertial coordinate system, t f1 is the terminal time of the first active segment, P(t f1) is the position vector of the first active segment at the terminal moment, V(t f1 ) is the velocity vector at the terminal moment of the first active segment, ha f1 is the orbital apogee height at the terminal moment, r B is the telecentricity of the sliding track, R e is the radius of the Earth, n T is the variable to be solved.

[0104] If the pericentric distance of the sliding track is defined as r A ,but

[0105] In addition, if the perigee velocity of the original target orbit is defined as V B The speed of the launch vehicle when it glides to the apogee of the glide track is V TB ,but

[0106] Then, the velocity increment ΔV of the launch vehicle when implementing the second active segment orbit change at the apogee of the glide segment is B =V B -V TB ,

[0107] Remaining flight time of the first active segment

[0108] Among them, m0 is the mass at the time of failure, m s is the structural mass, m load is the payload mass, V B is the perigee velocity of the original target orbit, μ is the constant coefficient of the earth's gravity, t f2 It is the flight duration of the second active segment.

[0109] Assume that the speed increment ΔV required for the second active segment track change is B Provided by the engine thrust, the flight time of the second active segment is

[0110] Therefore, the terminal constraint can be expressed as n T The function of , that is, the terminal state is:

[0111] [a f1 ,e f1 ,i f1 ,Ω f1 ,w f1 ,f f1 ] T =Fun orbit (P(t f1 ),V(t f1 )).

[0112]

[0113]

[0114] |i f1 -i ref |≤ε i ,|Ω f1 -Ω ref |≤ε Ω .

[0115] Among them, a f1 is the semi-major axis of the orbit at the terminal moment of the first active segment, e f1 is the orbital eccentricity at the terminal moment of the first active segment, i f1 is the orbital inclination at the terminal moment of the first active segment, Ω f1 is the longitude of the ascending node of the first active segment at the terminal time, w f1 is the orbital perigee argument at the terminal moment of the first active segment, f f1 is the true anomaly at the terminal moment of the first active segment, Fun orbit ( ) is the conversion function between the orbital elements of the first active segment and the position and velocity in the launch inertial coordinate system, t f1 is the terminal time of the first active segment, P(t f1 ) is the position vector of the first active segment at the terminal moment, V(t f1 ) is the velocity vector at the terminal moment of the first active segment, ha f1 is the orbital apogee height at the terminal moment, r B is the telecentricity of the sliding track, R e is the radius of the Earth, n T is the variable to be solved, m0 is the mass at the time of failure, m s is the structural mass, m load is the payload mass, V B is the perigee velocity of the original target orbit, μ is the constant coefficient of the earth's gravity, I sp is the engine specific impulse, i ref is the orbital inclination of the target orbit, ε i is the maximum value of the orbital inclination deviation (i.e. the allowable value of the orbital inclination deviation), Ω ref is the longitude of the ascending node of the target orbit, ε Ω It is the maximum value of the longitude deviation of the ascending node (i.e. the allowable value of the longitude deviation of the ascending node).

[0116] (4) Control quantity constraints

[0117] In order to simplify the replanning problem of the first active segment, the pitch program angle of the first active segment is defined as and the yaw program angle ψ are both constants, then the control quantity constraint can be expressed as:

[0118]

[0119] Where t is any moment in the first active segment, is the pitch program angle of the first active segment at time t, and ψ(t) is the yaw program angle of the first active segment at time t.

[0120] In summary, if the control variable is defined as Then the first active segment replanning problem can be expressed as:

[0121] Equations of motion:

[0122]

[0123] Initial state: [P, V, m] T (t0) = [P0, V0, m0] T ,

[0124] Terminal status: [a f1 ,e f1 ,i f1 ,Ω f1 ,w f1 ,f f1 ] T =Fun orbit (P(t f1 ),V(t f1 )),

[0125]

[0126]

[0127] |i f1 -i ref |≤ε i ,|Ω f1 -Ω ref |≤ε Ω .

[0128] Control quantity constraints:

[0129] 2) The nonlinear programming method is used to solve the re-planning problem of the first active segment, and the flight state sequence X1 and control quantity sequence U1 of the first active segment are obtained.

[0130] Among them, U1 is composed of the control quantities at each discrete point of the first active segment, and the control quantities at each discrete point of the first active segment are

[0131] is the pitch program angle of the first active segment at the time of failure, ψ0 is the yaw program angle of the first active segment at the time of failure, [ ] T is the transpose operation.

[0132] This step will describe the re-planning problem of the first active segment for step 1), and use nonlinear programming methods (such as interior point method, sequential quadratic programming, etc.) to solve it. Without considering the objective function, it can quickly converge to a feasible solution, thereby obtaining the flight state sequence X1 of the first active segment. The control quantity at each discrete point is The corresponding control quantity sequence can be expressed as U1.

[0133] 2. The implementation process of determining the flight state sequence of the glide segment based on the flight state of the launch vehicle

[0134] 1) According to the terminal state of the launch vehicle when it enters the glide orbit, determine the number of orbital elements in the glide segment.

[0135] Specifically, the number of rails in the sliding section is

[0136] in, is the semi-major axis of the glide segment, is the orbit eccentricity of the sliding segment, is the track inclination of the sliding segment, is the longitude of the orbit ascending node of the glide segment, is the orbit perigee argument of the glide segment, is the true anomaly of the orbit during the taxiing segment.

[0137] 2) Determine the velocity vector of each discrete point in the sliding segment based on the number of track elements in the sliding segment and position vector

[0138] Wherein, j is the identification of the discrete point of the glide segment, j=1,…,N+1, N+1 is the total number of discrete points of the glide segment, and N is the number of intervals bisected by the true anomaly angle of the glide segment.

[0139] In addition, the velocity vector of each discrete point in the sliding segment and position vector Satisfies the following relationship:

[0140]

[0141] Among them, Fun PV ( ) is the conversion function between the orbital elements of the glide segment and the position and velocity in the launch inertial coordinate system,

[0142] 3) Determine the time of each discrete point in the sliding segment.

[0143] Specifically, the time of each discrete point in the sliding segment satisfies the following relationship:

[0144]

[0145]

[0146] Among them, t f1 is the terminal time of the first active segment, j′ is the non-first discrete point identifier of the sliding segment, j′=2,…,N+1, is the eccentric anomaly angle of the discrete point j′ in the sliding segment, and μ is the constant coefficient of the earth’s gravity.

[0147] 4) According to the velocity vector of each discrete point in the fixed sliding segment and position vector The time of each discrete point in the taxiing segment is used to obtain the flight state sequence X of the taxiing segment. c and control quantity sequence U c .

[0148] Among them, U c It is composed of the control quantities at each discrete point of the sliding segment, and the control quantities at each discrete point of the sliding segment are all 0.

[0149] When determining the flight state sequence of the glide segment, the corresponding glide segment orbital elements can be obtained according to the terminal state of the rocket when it enters the glide orbit. As the rocket glides toward the apogee, its true anomaly angle gradually increases. When it reaches 180°, it indicates that the rocket has reached the apogee and the gliding segment ends. Define the total number of discrete points in the gliding segment as N+1, and divide the true anomaly angle into N intervals (that is, N is the number of intervals divided by the true anomaly angle of the gliding segment), and the interval between each interval is df c , then the true anomaly sequence can be expressed as:

[0150]

[0151]

[0152] According to the conversion relationship function Fun between the orbital elements of the glide segment and the position and velocity in the launch inertial coordinate system PV ( ), the velocity vector corresponding to each discrete point can be obtained and position vector As shown below,

[0153]

[0154] The time corresponding to each discrete point can be calculated using the following formula:

[0155]

[0156]

[0157] Define the sliding segment state sequence as X c Since there is no thrust effect, the control quantity sequence U c All elements in are set to zero.

[0158] 3. The implementation process of determining the flight state sequence of the second active segment based on the flight state of the launch vehicle

[0159] In the second active segment, it is necessary to keep the thrust direction u in the orbital plane always perpendicular to the direction of the distance from the center of the earth r, that is, u·r=1. r is the component of the distance from the center of the earth in the launching inertial coordinate system.

[0160] 1) Determine the time interval and initial state of the second active segment according to the remaining flight time of the launch vehicle.

[0161] Specifically, the time interval of the second active segment

[0162] Among them, t f2 is the remaining flight time of the launch vehicle, and M is the total number of discrete points in the second active segment - 1.

[0163] In addition, the initial state of the second active segment is the terminal state of the sliding segment.

[0164] 2) Based on the time interval and initial state of the second active segment, the flight state quantity sequence X2 and the control quantity sequence U2 of the second active segment are calculated by numerical integration.

[0165] In the process of determining the flight state sequence of the second active segment, according to the remaining flight time t f2 , define the second active segment to take (M+1) discrete points, then is the time interval, with the terminal state of the taxiing segment is the initial state of the second active segment. The flight state quantity sequence X2 and control quantity sequence U2 of the second active segment can be calculated by numerical integration methods (Euler integration, Runge-Kutta integration, etc.).

[0166] in, is the terminal position vector of the taxiing segment, is the terminal velocity vector of the glide segment, is the terminal mass of the taxiing segment.

[0167] 103. Estimating the initial value of replanning according to the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment.

[0168] After executing step 102, the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment will be obtained. This step will combine the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment (such as combining the state quantity sequence and the control quantity sequence of each segment), and the initial guesses of the three flight segments corresponding to the online replanning problem across the glide segment can be obtained, thereby achieving a rapid solution to the problem.

[0169] This embodiment is aimed at the mission profile of a launch vehicle with a glide stage in the orbital stage. In order to improve the fault adaptability, the launch vehicle needs to jointly plan the first active segment, the glide segment, and the second active segment online according to the flight status in the case of a thrust drop failure, so as to send the payload into the original target orbit. A method for estimating the initial value of online replanning of thrust drop across the glide stage of a launch vehicle is proposed. By reasonably estimating the initial conjecture of the optimization problem, the iterative solution process of the numerical optimization method can converge quickly.

[0170] In the specific implementation, a glide trajectory optimization problem is constructed, and a numerical iteration method is used to quickly search for feasible solutions as the initial guess of the first active segment. Then, when the rocket glides to the apogee, it enters the second active segment, and the initial guess of the glide segment is obtained based on the conversion relationship between the orbital elements and the speed and position. After entering the second active segment, the thrust direction is kept perpendicular to the direction of the distance from the center of the earth in the orbital plane, and the initial guess of the second active segment is obtained by numerical integration. Finally, the initial values ​​estimated in the three flight segments are combined as the initial guess of the online trajectory planning problem of the launch vehicle across the glide segment to achieve a fast solution.

[0171] by Figure 2 As an example, the method provided in this embodiment constructs the first active segment replanning problem, and then iteratively solves the first active segment flight state sequence based on the problem. The glide segment flight state sequence is calculated, and the second active segment flight state sequence is calculated. Then, according to the first active segment flight state sequence, the glide segment flight state sequence, and the second active segment flight state sequence, an initial guess of the cross-glide segment online replanning problem is constructed, and then the replanning initial value estimation is performed.

[0172] The initial value estimation method for online re-planning of thrust reduction across the glide phase of a launch vehicle provided in this embodiment combines the motion characteristics of the rocket's orbit entry process and decomposes the online planning problem across the glide phase into three flight phases for separate calculation, so that the initial guess satisfies the constraints of the motion equation and reduces the difficulty of numerical solution.

[0173] In addition, the initial value estimation method for online replanning of thrust descent across the glide phase of a launch vehicle provided in this embodiment simplifies the form of the control variables of the first active phase into constant values, reduces the dimension of the variables to be solved, and constructs a first active phase replanning problem that is easy to iteratively solve.

[0174] In addition, the online re-planning initial value estimation method for the thrust reduction of the launch vehicle across the glide phase provided in this embodiment is based on the orbit transfer theory, and a motion state estimation method for the glide phase and the second active phase is designed that does not require numerical iterative solution. While satisfying the rationality of the initial guess, it improves the speed of generating the initial guess.

[0175] The present application provides a method for estimating initial values ​​of online replanning of thrust reduction across a glide segment of a launch vehicle, the method comprising: dividing the trajectory of a launch vehicle under a thrust reduction failure into a first active segment, a glide segment, and a second active segment; determining the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment based on the flight state of the launch vehicle; and estimating initial values ​​of replanning according to the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment. The present application divides the trajectory of a launch vehicle under a thrust reduction failure into a first active segment, a glide segment, and a second active segment, and simultaneously considers the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment to estimate initial values ​​of replanning, making the estimation process more reasonable, thereby improving the convergence and rapidity of numerical replanning.

[0176] Based on the same inventive concept of the initial value estimation method for online re-planning of thrust drop across the glide phase of a launch vehicle, this embodiment provides an electronic device, which includes: a memory, a processor, and a computer program.

[0177] The computer program is stored in the memory and is configured to be executed by the processor to implement the above Figure 1 The initial value estimation method for online re-planning of thrust descent of the launch vehicle across the glide phase is shown.

[0178] Specifically,

[0179] The trajectory of the launch vehicle in the case of thrust reduction failure is divided into the first active segment, the glide segment, and the second active segment.

[0180] The flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment are determined based on the flight state of the launch vehicle.

[0181] The replanning initial value estimation is performed according to the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment.

[0182] Optionally, the pitch program angle and the yaw program angle of the first active segment are both constant values.

[0183] The apogee altitude of the glide segment is the perigee altitude of the original target orbit.

[0184] In the second active segment, the thrust direction of the launch vehicle is always perpendicular to the direction of the Earth's center in the orbital plane.

[0185] Among them, the carrier rocket enters the second active stage when it glides to the apogee.

[0186] Optionally, determining a flight state sequence of the first active segment based on the flight state of the launch vehicle includes:

[0187] The first active segment replanning problem is constructed. The first active segment replanning problem includes: motion equations, initial state, terminal state and control quantity constraints.

[0188] The nonlinear programming method is used to solve the re-planning problem of the first active segment, and the flight state sequence X1 and control quantity sequence U1 of the first active segment are obtained.

[0189] Among them, U1 is composed of the control quantities at each discrete point of the first active segment, and the control quantities at each discrete point of the first active segment are

[0190] is the pitch program angle of the first active segment at the time of failure, ψ0 is the yaw program angle of the first active segment at the time of failure, [] T is the transpose operation.

[0191] Alternatively, the equation of motion is:

[0192]

[0193] T=I sp dm.

[0194]

[0195] Where · is the first-order derivative operator, P is the position vector, V is the velocity vector, T is the engine thrust after the failure, m is the mass, u is the thrust direction, μ is the constant coefficient of the earth's gravity, r is the distance from the center of mass of the rocket to the center of the earth, r is the component of the distance from the center of the earth in the launch inertial coordinate system, dm is the flow rate per second after the failure, I sp is the engine specific impulse, is the pitch program angle of the first active segment, and ψ is the yaw program angle of the first active segment.

[0196] Optionally, in the launch inertial coordinate system, the origin O is at the launch point, the OX axis points to the launch direction in the horizontal plane, the OY axis is perpendicular to the local horizontal plane of the launch point and points to the sky, and the OZ axis satisfies the right-hand rule.

[0197] Optionally, the initial state is:

[0198] [P,V,m] T (t0) = [P0, V0, m0]T .

[0199] Among them, P is the position vector, V is the velocity vector, m is the mass, t0 is the fault time, P0 is the position vector at the fault time, V0 is the velocity vector at the fault time, and m0 is the mass at the fault time.

[0200] Optionally, the terminal status is:

[0201] [a f1 ,e f1 ,i f1 ,Ω f1 ,w f1 ,f f1 ] T =Fun orbit (P(t f1 ),V(t f1 )).

[0202]

[0203]

[0204] |i f1 -i ref |≤ε i ,|Ω f1 -Ω ref |≤ε Ω .

[0205] Among them, a f1 is the semi-major axis of the orbit at the terminal moment of the first active segment, e f1 is the orbital eccentricity at the terminal moment of the first active segment, i f1 is the orbital inclination at the terminal moment of the first active segment, Ω f1 is the longitude of the ascending node of the first active segment at the terminal time, w f1 is the orbital perigee argument at the terminal moment of the first active segment, f f1 is the true anomaly at the terminal moment of the first active segment, Fun orbit ( ) is the conversion function between the orbital elements of the first active segment and the position and velocity in the launch inertial coordinate system, t f1 is the terminal time of the first active segment, P(t f1 ) is the position vector of the first active segment at the terminal moment, V(t f1 ) is the velocity vector at the terminal moment of the first active segment, ha f1 is the orbital apogee height at the terminal moment, r B is the telecentricity of the sliding track, R e is the radius of the Earth, n Tis the variable to be solved, m0 is the mass at the time of failure, m s is the structural mass, m load is the payload mass, V B is the perigee velocity of the original target orbit, μ is the constant coefficient of the earth's gravity, I sp is the engine specific impulse, i ref is the orbital inclination of the target orbit, ε i is the maximum value of the orbital inclination deviation, Ω ref is the longitude of the ascending node of the target orbit, ε Ω is the maximum value of the longitude deviation of the ascending node.

[0206] Optionally,

[0207] Among them, r A is the pericentric distance of the sliding track.

[0208] Optionally, the control volume constraint is:

[0209]

[0210] Where t is any moment in the first active segment, is the pitch program angle of the first active segment at time t, and ψ(t) is the yaw program angle of the first active segment at time t.

[0211] Optionally, determining a flight state sequence of the glide segment based on the flight state of the launch vehicle includes:

[0212] The number of orbital roots of the glide segment is determined based on the terminal state of the launch vehicle when it enters the glide orbit.

[0213] Determine the velocity vector of each discrete point of the sliding segment based on the number of track elements of the sliding segment and position vector Wherein, j is the identification of the discrete point of the glide segment, j=1,…,N+1, N+1 is the total number of discrete points of the glide segment, and N is the number of intervals bisected by the true anomaly angle of the glide segment.

[0214] Determine the time of each discrete point in the taxiing segment.

[0215] According to the velocity vector of each discrete point in the fixed sliding segment and position vector The time of each discrete point in the taxiing segment is used to obtain the flight state sequence X of the taxiing segment. c and control quantity sequence U c .

[0216] Among them, U c It is composed of the control quantities at each discrete point of the sliding segment, and the control quantities at each discrete point of the sliding segment are all 0.

[0217] Optionally, the number of sliding segment tracks is

[0218] in, is the semi-major axis of the glide segment, is the orbit eccentricity of the sliding segment, is the track inclination of the sliding segment, is the longitude of the orbit ascending node of the glide segment, is the orbit perigee argument of the glide segment, is the true anomaly of the orbit during the taxiing segment.

[0219] Optionally, the velocity vector of each discrete point in the sliding segment and position vector Satisfies the following relationship:

[0220]

[0221] Among them, Fun PV ( ) is the conversion function between the orbital elements of the glide segment and the position and velocity in the launch inertial coordinate system,

[0222] Optionally, the time of each discrete point in the sliding segment satisfies the following relationship:

[0223]

[0224]

[0225] Among them, t f1 is the terminal time of the first active segment, j′ is the non-first discrete point identifier of the sliding segment, j′=2,…,N+1, is the eccentric anomaly angle of the discrete point j′ in the sliding segment, and μ is the constant coefficient of the earth’s gravity.

[0226] Optionally, determining a flight state sequence of the second active segment based on the flight state of the launch vehicle includes:

[0227] The time interval and initial state of the second active segment are determined according to the remaining flight time of the launch vehicle.

[0228] Based on the time interval and initial state of the second active segment, the flight state quantity sequence X2 and the control quantity sequence U2 of the second active segment are calculated by numerical integration.

[0229] Optionally, the time interval of the second active segment

[0230] Among them, t f2 is the remaining flight time of the launch vehicle, and M is the total number of discrete points in the second active segment - 1.

[0231] Optionally, the initial state of the second active segment is the terminal state of the gliding segment.

[0232] The electronic device provided in this embodiment has a computer program executed by a processor to divide the trajectory of a carrier rocket in the event of a thrust drop failure into a first active segment, a glide segment, and a second active segment; determine the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment based on the flight state of the carrier rocket; and perform replanning initial value estimation based on the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment. This embodiment divides the trajectory of a carrier rocket in the event of a thrust drop failure into a first active segment, a glide segment, and a second active segment, and simultaneously considers the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment to perform replanning initial value estimation, making the estimation process more reasonable, thereby improving the convergence and rapidity of numerical replanning.

[0233] Based on the same inventive concept of the method for estimating the initial value of the thrust reduction of the launch vehicle across the glide phase, this embodiment provides a computer having a computer program stored thereon. The computer program is executed by a processor to implement the above Figure 1 The initial value estimation method for online re-planning of thrust descent of the launch vehicle across the glide phase is shown.

[0234] Specifically,

[0235] The trajectory of the launch vehicle in the case of thrust reduction failure is divided into the first active segment, the glide segment, and the second active segment.

[0236] The flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment are determined based on the flight state of the launch vehicle.

[0237] The replanning initial value estimation is performed according to the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment.

[0238] Optionally, the pitch program angle and the yaw program angle of the first active segment are both constant values.

[0239] The apogee altitude of the glide segment is the perigee altitude of the original target orbit.

[0240] In the second active segment, the thrust direction of the launch vehicle is always perpendicular to the direction of the Earth's center in the orbital plane.

[0241] Among them, the carrier rocket enters the second active stage when it glides to the apogee.

[0242] Optionally, determining a flight state sequence of the first active segment based on the flight state of the launch vehicle includes:

[0243] The first active segment replanning problem is constructed. The first active segment replanning problem includes: motion equations, initial state, terminal state and control quantity constraints.

[0244] The nonlinear programming method is used to solve the re-planning problem of the first active segment, and the flight state sequence X1 and control quantity sequence U1 of the first active segment are obtained.

[0245] Among them, U1 is composed of the control quantities at each discrete point of the first active segment, and the control quantities at each discrete point of the first active segment are

[0246] is the pitch program angle of the first active segment at the time of failure, ψ0 is the yaw program angle of the first active segment at the time of failure, [] T is the transpose operation.

[0247] Alternatively, the equation of motion is:

[0248]

[0249] T=I sp dm.

[0250]

[0251] Where · is the first-order derivative operator, P is the position vector, V is the velocity vector, T is the engine thrust after the failure, m is the mass, u is the thrust direction, μ is the constant coefficient of the earth's gravity, r is the distance from the center of mass of the rocket to the center of the earth, r is the component of the distance from the center of the earth in the launch inertial coordinate system, dm is the flow rate per second after the failure, I sp is the engine specific impulse, is the pitch program angle of the first active segment, and ψ is the yaw program angle of the first active segment.

[0252] Optionally, in the launch inertial coordinate system, the origin O is at the launch point, the OX axis points to the launch direction in the horizontal plane, the OY axis is perpendicular to the local horizontal plane of the launch point and points to the sky, and the OZ axis satisfies the right-hand rule.

[0253] Optionally, the initial state is:

[0254] [P,V,m] T (t0) = [P0, V0, m0] T .

[0255] Among them, P is the position vector, V is the velocity vector, m is the mass, t0 is the fault time, P0 is the position vector at the fault time, V0 is the velocity vector at the fault time, and m0 is the mass at the fault time.

[0256] Optionally, the terminal status is:

[0257] [a f1 , e f1 ,i f1 ,Ω f1 ,w f1 ,f f1 ] T =Fun orbit (P(t f1 ),V(t f1 )).

[0258]

[0259]

[0260] |i f1 -i ref |≤ε i ,|Ω f1 -Ω ref |≤ε Ω .

[0261] Among them, a f1 is the semi-major axis of the orbit at the terminal moment of the first active segment, e f1 is the orbital eccentricity at the terminal moment of the first active segment, i f1 is the orbital inclination at the terminal moment of the first active segment, Ω f1 is the longitude of the ascending node of the first active segment at the terminal time, w f1 is the orbital perigee argument at the terminal moment of the first active segment, f f1 is the true anomaly at the terminal moment of the first active segment, Fun orbit ( ) is the conversion function between the orbital elements of the first active segment and the position and velocity in the launch inertial coordinate system, t f1 is the terminal time of the first active segment, P(t f1 ) is the position vector of the first active segment at the terminal moment, V(t f1 ) is the velocity vector at the terminal moment of the first active segment, ha f1 is the orbital apogee height at the terminal moment, r B is the telecentricity of the sliding track, R e is the radius of the Earth, n T is the variable to be solved, m0 is the mass at the time of failure, m s is the structural mass, m load is the payload mass, V B is the perigee velocity of the original target orbit, μ is the constant coefficient of the earth's gravity, I sp is the engine specific impulse, i ref is the orbital inclination of the target orbit, ε iis the maximum value of the orbital inclination deviation, Ω ref is the longitude of the ascending node of the target orbit, ε Ω is the maximum value of the longitude deviation of the ascending node.

[0262] Optionally,

[0263] Among them, r A is the pericentric distance of the sliding track.

[0264] Optionally, the control volume constraint is:

[0265]

[0266] Where t is any moment in the first active segment, is the pitch program angle of the first active segment at time t, and ψ(t) is the yaw program angle of the first active segment at time t.

[0267] Optionally, determining a flight state sequence of the glide segment based on the flight state of the launch vehicle includes:

[0268] The number of orbital roots of the glide segment is determined based on the terminal state of the launch vehicle when it enters the glide orbit.

[0269] Determine the velocity vector of each discrete point of the sliding segment based on the number of track elements of the sliding segment and position vector Wherein, j is the identification of the discrete point of the glide segment, j=1,…,N+1, N+1 is the total number of discrete points of the glide segment, and N is the number of intervals bisected by the true anomaly angle of the glide segment.

[0270] Determine the time of each discrete point in the taxiing segment.

[0271] According to the velocity vector of each discrete point in the fixed sliding segment and position vector The time of each discrete point in the taxiing segment is used to obtain the flight state sequence X of the taxiing segment. c and control quantity sequence U c .

[0272] Among them, U c It is composed of the control quantities at each discrete point of the sliding segment, and the control quantities at each discrete point of the sliding segment are all 0.

[0273] Optionally, the number of sliding segment tracks is

[0274] in, is the semi-major axis of the glide segment, is the orbit eccentricity of the sliding segment, is the track inclination of the sliding segment, is the longitude of the orbit ascending node of the glide segment, is the orbit perigee argument of the glide segment, is the true anomaly of the orbit during the taxiing segment.

[0275] Optionally, the velocity vector of each discrete point in the sliding segment and position vector Satisfies the following relationship:

[0276]

[0277] Among them, Fun PV ( ) is the conversion function between the orbital elements of the glide segment and the position and velocity in the launch inertial coordinate system,

[0278] Optionally, the time of each discrete point in the sliding segment satisfies the following relationship:

[0279]

[0280]

[0281] Among them, t f1 is the terminal time of the first active segment, j′ is the non-first discrete point identifier of the sliding segment, j′=2,…,N+1, is the eccentric anomaly angle of the discrete point j′ in the sliding segment, and μ is the constant coefficient of the earth’s gravity.

[0282] Optionally, determining a flight state sequence of the second active segment based on the flight state of the launch vehicle includes:

[0283] The time interval and initial state of the second active segment are determined according to the remaining flight time of the launch vehicle.

[0284] Based on the time interval and initial state of the second active segment, the flight state quantity sequence X2 and the control quantity sequence U2 of the second active segment are calculated by numerical integration.

[0285] Optionally, the time interval of the second active segment

[0286] Among them, t f2 is the remaining flight time of the launch vehicle, and M is the total number of discrete points in the second active segment - 1.

[0287] Optionally, the initial state of the second active segment is the terminal state of the gliding segment.

[0288] The computer-readable storage medium provided in this embodiment has a computer program executed by a processor to divide the trajectory of a launch vehicle in the case of a thrust drop failure into a first active segment, a glide segment, and a second active segment; determine the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment based on the flight state of the launch vehicle; and perform replanning initial value estimation based on the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment. This embodiment divides the trajectory of a launch vehicle in the case of a thrust drop failure into a first active segment, a glide segment, and a second active segment, and simultaneously considers the flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment to perform replanning initial value estimation, so that the estimation process is more reasonable, thereby improving the convergence and rapidity of numerical replanning.

[0289] Those skilled in the art will appreciate that the embodiments of the present application can be provided as methods, systems, or computer program products. Therefore, the present application can adopt the form of complete hardware embodiments, complete software embodiments, or embodiments in combination with software and hardware. Moreover, the present application can adopt the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) that contain computer-usable program code. The scheme in the embodiments of the present application can be implemented in various computer languages, for example, object-oriented programming language Java and literal scripting language JavaScript, etc.

[0290] The present application is described with reference to the flowcharts and / or block diagrams of the methods, devices (systems), and computer program products according to the embodiments of the present application. It should be understood that each process and / or box in the flowchart and / or block diagram, as well as the combination of the processes and / or boxes in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to generate a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0291] These computer program instructions may also be stored in a computer-readable memory capable of directing a computer or other programmable data processing device to operate in a specific manner, so that the instructions stored in the computer-readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0292] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operating steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing instructions for implementing the process. Figure 1 A process or multiple processes and / or boxes Figure 1 The steps for the functions specified in one or more boxes.

[0293] Although the preferred embodiments of the present application have been described, those skilled in the art may make other changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications falling within the scope of the present application.

[0294] Obviously, those skilled in the art can make various changes and modifications to the present application without departing from the spirit and scope of the present application. Thus, if these modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is also intended to include these modifications and variations.

Claims

1. A method for estimating initial values ​​of online replanning of thrust reduction across the glide phase of a launch vehicle, characterized in that: The method comprises: The trajectory of the launch vehicle under thrust reduction failure condition is divided into the first active segment, the glide segment, and the second active segment; The flight state sequence of the first active segment, the flight state sequence of the glide segment, and the flight state sequence of the second active segment are determined based on the flight state of the carrier rocket; wherein: Determining a flight state sequence of the first active segment based on the flight state of the launch vehicle includes: Constructing a first active segment replanning problem; the first active segment replanning problem includes: a motion equation, an initial state, a terminal state and a control quantity constraint; A nonlinear programming method is used to solve the re-planning problem of the first active segment to obtain a flight state sequence X1 and a control quantity sequence U1 of the first active segment; Among them, U1 is composed of the control quantities at each discrete point of the first active segment, and the control quantities at each discrete point of the first active segment are is the pitch program angle of the first active segment at the time of failure, ψ0 is the yaw program angle of the first active segment at the time of failure, [] T is the transpose operation; Determining the flight state sequence of the glide segment based on the flight state of the carrier rocket includes: Determining the number of orbital elements of the glide segment according to the terminal state of the launch vehicle at the moment of entering the glide orbit; Determine the velocity vector of each discrete point of the sliding segment based on the number of track elements of the sliding segment and position vector Wherein, j is the identifier of the discrete point of the glide segment, j=1,…,N+1, N+1 is the total number of discrete points of the glide segment, and N is the number of intervals bisected by the true anomaly angle of the glide segment; Determining the time of each discrete point of the sliding segment; According to the velocity vector of each discrete point of the sliding segment and position vector The time of each discrete point of the glide segment is used to obtain the flight state sequence X of the glide segment. c and control quantity sequence U c ; Among them, U c It is composed of control quantities at each discrete point of the sliding segment, and the control quantities at each discrete point of the sliding segment are all 0; Determining a flight state sequence of the second active segment based on the flight state of the launch vehicle includes: Determining the time interval and initial state of the second active segment according to the remaining flight time of the carrier rocket; Based on the time interval and initial state of the second active segment, the flight state quantity sequence X2 and the control quantity sequence U2 of the second active segment are calculated by numerical integration; The replanning initial value estimation is performed according to the flight state sequence of the first active segment, the flight state sequence of the gliding segment, and the flight state sequence of the second active segment.

2. The method according to claim 1, characterized in that: The pitch program angle and the yaw program angle of the first active segment are both constant values; The apogee height of the glide segment is the perigee height of the original target orbit; In the second active segment, the thrust direction of the launch vehicle is always perpendicular to the direction of the distance from the center of the earth in the orbital plane; Wherein, the carrier rocket enters the second active section when gliding to the apogee.

3. The method according to claim 2, characterized in that The equation of motion is: T=I sp dm; Where · is the first-order derivative operator, P is the position vector, V is the velocity vector, T is the engine thrust after the failure, m is the mass, u is the thrust direction, μ is the constant coefficient of the earth's gravity, r is the distance from the center of mass of the rocket to the center of the earth, r is the component of the distance from the center of the earth in the launch inertial coordinate system, dm is the flow rate per second after the failure, Isp is the engine specific impulse, is the pitch program angle of the first active segment, and ψ is the yaw program angle of the first active segment.

4. The method according to claim 3, characterized in that In the launch inertial coordinate system, the origin O is at the launch point, the OX axis points to the launch direction in the horizontal plane, the OY axis is perpendicular to the local horizontal plane of the launch point and points to the sky, and the OZ axis satisfies the right-hand rule.

5. The method according to claim 2, characterized in that: The initial state is: [P,V,m] T (t0)=[P0,V0,m0] T ; Among them, P is the position vector, V is the velocity vector, m is the mass, t0 is the fault time, P0 is the position vector at the fault time, V0 is the velocity vector at the fault time, and m0 is the mass at the fault time.

6. The method according to claim 2, characterized in that The terminal status is: [a f1 ,e f1 ,i f1 ,Ω f1 ,w f1 ,f f1 ] T =Fun orbit (P(t f1 ),V(t f1 )); |i f1 -i ref |≤e i ,|Oh f1 -Oh ref |≤e Ω ; Among them, a f1 is the semi-major axis of the orbit at the terminal moment of the first active segment, e f1 is the orbital eccentricity at the terminal moment of the first active segment, i f1 is the orbital inclination at the terminal moment of the first active segment, Ω f1 is the longitude of the ascending node of the first active segment at the terminal time, w f1 is the orbital perigee argument at the terminal moment of the first active segment, f f1 is the true anomaly at the terminal moment of the first active segment, Fun orbit () is the conversion function between the orbital elements of the first active segment and the position and velocity in the launch inertial coordinate system, t f1 is the terminal time of the first active segment, P(t f1 ) is the position vector of the first active segment at the terminal moment, V(t f1 ) is the velocity vector at the terminal moment of the first active segment, ha f1 is the orbital apogee height at the terminal moment, r B is the telecentric distance of the sliding track, R e is the radius of the Earth, n T is the variable to be solved, m0 is the mass at the time of failure, m s is the structural mass, m load is the payload mass, V B is the perigee velocity of the original target orbit, μ is the constant coefficient of the earth's gravity, I sp is the engine specific impulse, i ref is the orbital inclination of the target orbit, ε i is the maximum value of the orbital inclination deviation, Ω ref is the longitude of the ascending node of the target orbit, ε Ω is the maximum value of the longitude deviation of the ascending node.

7. The method according to claim 6, characterized in that Among them, r A is the pericentric distance of the sliding track.

8. The method according to claim 2, characterized in that: The control quantity constraint is: Wherein, t is any moment of the first active segment, is the pitch program angle of the first active segment at time t, and ψ(t) is the yaw program angle of the first active segment at time t.

9. The method according to claim 2, characterized in that: The number of rails in the sliding section is in, is the semi-major axis of the orbit of the glide segment, is the track eccentricity of the sliding segment, is the track inclination of the sliding segment, is the longitude of the orbit ascending node of the glide segment, is the orbital perigee argument of the glide segment, is the true anomaly of the orbit of the glide segment.

10. The method according to claim 9, characterized in that The velocity vector of each discrete point in the sliding section and position vector Satisfies the following relationship: Among them, Fun PV () is the conversion function between the orbital elements of the sliding segment and the position and velocity in the launch inertial coordinate system, 11. The method according to claim 2, characterized in that The time interval of the second active segment Among them, t f2 is the remaining flight time of the launch vehicle, and M is the total number of discrete points in the second active segment - 1.

12. The method according to claim 2, characterized in that: The initial state of the second active segment is the terminal state of the gliding segment.

13. An electronic device, characterized in that: include: Memory; processor; as well as Computer programs; The computer program is stored in the memory and is configured to be executed by the processor to implement the method according to any one of claims 1 to 12.

14. A computer-readable storage medium, characterized in that: A computer program is stored thereon; the computer program is executed by a processor to implement the method according to any one of claims 1 to 12.

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