CN114417569A - Carrier rocket cross-sliding section thrust descent online re-planning initial value estimation method - Google Patents

Carrier rocket cross-sliding section thrust descent online re-planning initial value estimation method Download PDF

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CN114417569A
CN114417569A CN202111595032.7A CN202111595032A CN114417569A CN 114417569 A CN114417569 A CN 114417569A CN 202111595032 A CN202111595032 A CN 202111595032A CN 114417569 A CN114417569 A CN 114417569A
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section
active segment
active
taxiing
segment
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宋征宇
王聪
巩庆海
施国兴
何勇
胡海峰
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Beijing Aerospace Automatic Control Research Institute
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Beijing Aerospace Automatic Control Research Institute
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    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces

Abstract

The application provides a method for estimating an initial value of thrust descent online re-planning of a carrier rocket in a cross-sliding section, which comprises the following steps: dividing a track of a carrier rocket under the condition of thrust descent failure into a first active section, a gliding section and a second active section; determining a flight state sequence of a first active section, a flight state sequence of a taxiing section and a flight state sequence of a second active section based on the flight state of the carrier rocket; and carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment. The method divides the track of the carrier rocket under the condition of thrust descent fault into a first active section, a gliding section and a second active section, and simultaneously considers the flight state sequence of the first active section, the flight state sequence of the gliding section and the flight state sequence of the second active section to carry out re-planning initial value estimation, so that the estimation process is more reasonable, and the convergence and the rapidity of numerical value re-planning are further improved.

Description

Carrier rocket cross-sliding section thrust descent online re-planning initial value estimation method
Technical Field
The application relates to the technical field of carrier rocket control, in particular to an initial value estimation method for thrust descent online re-planning of a carrier rocket in a cross-sliding section.
Background
The carrier rocket is divided into an active section and a gliding section in the process of taking off from the ignition of the launching platform until the spacecraft is delivered to a preset orbit.
The active section is an orbit section which flies during the working period of the rocket engine, and the gliding section is an orbit section which does not generate thrust after the engine is shut down, namely an inertia flying section.
Aiming at a carrier rocket flight mission section with a gliding section at an orbit entering stage, how to improve the fault adaptability becomes the key point of the current research.
Disclosure of Invention
In order to solve one of the technical defects, the application provides an initial value estimation method for thrust descent online re-planning of a carrier rocket in a cross-sliding section.
In a first aspect of the application, a method for estimating an initial value of thrust descent online re-planning of a carrier rocket in a cross-sliding section is provided, and the method comprises the following steps:
dividing a track of a carrier rocket under the condition of thrust descent failure into a first active section, a gliding section and a second active section;
determining a flight state sequence of the first active segment, a flight state sequence of the taxiing segment, and a flight state sequence of the second active segment based on a flight state of a launch vehicle;
and carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment.
Optionally, the pitch program angle and the yaw program angle of the first active segment are both constant values;
the height of the far place of the taxiing section is the height of the near place of the original target track;
in the second active section, the thrust direction of the carrier rocket is always vertical to the direction of the earth center distance in the orbital plane;
wherein the launch vehicle enters the second active segment when taxiing to the remote site.
Optionally, determining a flight state sequence of the first active segment based on the flight state of the launch vehicle comprises:
constructing a first active section re-planning problem; the first active segment re-planning problem comprises: motion equation, initial state, terminal state and controlled variable constraint;
solving the re-planning problem of the first active segment by adopting a nonlinear planning method to obtain a flight state sequence X of the first active segment1And a sequence of control quantities U1；
Wherein, U1The control quantity of each discrete point of the first active section is
Figure BDA0003430835140000021
Figure BDA0003430835140000022
For the pitch program angle, ψ, of the first active segment at the moment of failure0Is a yaw program angle of the first active segment at a failure time]TIs a transpose operation.
Optionally, the equation of motion is:
Figure BDA0003430835140000023
T＝Ispdm；
Figure BDA0003430835140000024
wherein, for the operator of first derivative, P is position vector, V is velocity vector, T is engine thrust after fault, m is mass, u is thrust direction, mu is constant coefficient of earth gravity, r is distance from center of mass to center of earth, r is component of center of earth to under launching inertia coordinate system, dm is second flow after fault, IspIs the specific impulse of the engine,
Figure BDA0003430835140000025
is the pitch program angle of the first active segment and psi is the yaw program angle of the first active segment.
Optionally, in the emission inertial coordinate system, the origin O is at an emission point, the OX axis points to the emission direction in a horizontal plane, the OY axis points to the sky in a vertical direction to the local horizontal plane of the emission point, and the OZ axis satisfies the right-hand rule.
Optionally, the initial state is:
[P,V,m]T(t0)＝[P0,V0,m0]T；
where P is the position vector, V is the velocity vector, m is the mass, t0To the moment of failure, P0Is a position vector of the moment of failure, V0Is the velocity vector at the moment of failure, m0Is the quality of the moment of failure.
Optionally, the terminal state is:
[af1，ef1,if1,Ωf1,wf1,ff1]T＝Funorbit(P(tf1),V(tf1))；
Figure BDA0003430835140000031
Figure BDA0003430835140000032
|if1-iref|≤εi,|Ωf1-Ωref|≤εΩ；
wherein, af1Is the terminal time orbit semi-major axis, e, of the first active segmentf1Is the terminal moment track eccentricity of the first active segment if1Is the terminal time orbit inclination angle, omega, of the first active segmentf1An ascending node longitude, w, of the terminal time orbit of the first active segmentf1Is the terminal time orbit near-place argument, f of the first active segmentf1Is the terminal time true approach angle, Fun, of the first active segmentorbit() Is a conversion relation function between the number of the tracks of the first active section and the position and the speed under a launching inertia coordinate system, tf1Is the terminal time of the first active segment, P (t)f1) Is the position vector of the terminal moment of the first active segment, V (t)f1) Is the velocity vector, ha, of the terminal instant of the first active segmentf1The track apogee height at the terminal time, rBIs the telecentric distance of the sliding track, ReIs the radius of the earth, nTFor the variable to be solved, m0M is the quality of the moment of failuresIs the structural mass, mloadFor payload mass, VBIs the near-to-earth velocity of the original target orbit, mu is the constant coefficient of the earth's gravity, IspFor engine specific impulse, irefIs the track inclination of the target track, epsiloniMaximum value of track inclination deviation, omegarefIs the elevation point longitude, epsilon, of the target trackΩIs the maximum value of the elevation point longitude deviation.
Alternatively,
Figure BDA0003430835140000033
wherein r isAThe close center distance of the sliding track.
Optionally, the control amount constraint is:
Figure BDA0003430835140000034
wherein t is any time of the first active segment,
Figure BDA0003430835140000041
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.
Optionally, determining a sequence of flight states of the taxiing sections based on the flight state of the launch vehicle comprises:
determining the number of orbits of the gliding section according to the terminal state of the carrier rocket at the moment of entering the gliding orbit;
determining a velocity vector of each discrete point of the taxiing section based on the number of the taxiing section tracks
Figure BDA0003430835140000042
And position vector
Figure BDA0003430835140000043
Wherein j is the taxiing section discrete point identifier, j is 1, …, N +1 is the total number of discrete points of the taxiing section, and N is the number of sections bisected by the real approximate point angle of the taxiing section;
determining the time of each discrete point of the taxiing section;
according to the velocity vector of each discrete point of the taxiing section
Figure BDA0003430835140000044
And position vector
Figure BDA0003430835140000045
Obtaining the flight state sequence X of the taxiing section according to the time of each discrete point of the taxiing sectioncAnd a sequence of control quantities Uc；
Wherein, UcThe control quantity of each discrete point of the taxiing section is 0.
Optionally, the number of the sliding section rails is
Figure BDA0003430835140000046
Wherein the content of the first and second substances,
Figure BDA0003430835140000047
is the semi-long axis of the track of the sliding section,
Figure BDA0003430835140000048
is the orbital eccentricity of the skid section,
Figure BDA0003430835140000049
is the angle of inclination of the track of the skid section,
Figure BDA00034308351400000410
the point longitude is incremented for the track of the taxiing section,
Figure BDA00034308351400000411
is the orbital perigee argument of the taxiing section,
Figure BDA00034308351400000412
is the true periorbital angle of the skid section.
Optionally, the velocity vector of each discrete point of the taxiing section
Figure BDA00034308351400000413
And position vector
Figure BDA00034308351400000414
The following relationship is satisfied:
Figure BDA00034308351400000415
wherein, FunPV() As a function of the conversion relation between the number of the taxiing section orbit and the position and the speed under the emission inertia coordinate system,
Figure BDA00034308351400000416
optionally, the time of each discrete point of the taxiing section satisfies the following relationship:
Figure BDA00034308351400000417
Figure BDA00034308351400000418
wherein, tf1Is the firstThe terminal time of the active segment, j ', is the non-first discrete point identification of the taxiing segment, j' is 2, …, N +1,
Figure BDA0003430835140000051
is the approximate point angle of the discrete point j' of the taxiing section, and mu is the gravity constant coefficient of the earth.
Optionally, determining a flight state sequence of the second active segment based on the flight state of the launch vehicle comprises:
determining the time interval and the initial state of the second active segment according to the remaining flight time of the carrier rocket;
calculating the flight state quantity sequence X of the second active segment by adopting numerical integration based on the time interval and the initial state of the second active segment2And a sequence of control quantities U2。
Optionally, a time interval of the second active segment
Figure BDA0003430835140000052
Wherein, tf2And M is the total number of the discrete points of the second active section, namely-1, wherein M is the residual flight time of the carrier rocket.
Optionally, the initial state of the second active segment is a terminal state of the taxiing segment.
In a second aspect of the present application, there is provided an electronic device comprising:
a memory;
a processor; and
a computer program;
wherein the computer program is stored in the memory and configured to be executed by the processor to implement the method according to the first aspect.
In a third aspect of the present application, there is provided a computer readable storage medium having a computer program stored thereon; the computer program is executed by a processor to implement the method according to the first aspect as described above.
The application provides a method for estimating an initial value of thrust descent online re-planning of a carrier rocket in a cross-sliding section, which comprises the following steps: dividing a track of a carrier rocket under the condition of thrust descent failure into a first active section, a gliding section and a second active section; determining a flight state sequence of a first active section, a flight state sequence of a taxiing section and a flight state sequence of a second active section based on the flight state of the carrier rocket; and carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment. The method divides the track of the carrier rocket under the condition of thrust descent fault into a first active section, a gliding section and a second active section, and simultaneously considers the flight state sequence of the first active section, the flight state sequence of the gliding section and the flight state sequence of the second active section to carry out re-planning initial value estimation, so that the estimation process is more reasonable, and the convergence and the rapidity of numerical value re-planning are further improved.
Drawings
The accompanying drawings, which are included to provide a further understanding of the application and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the application and together with the description serve to explain the application and not to limit the application. In the drawings:
fig. 1 is a flowchart of an initial value estimation method for thrust descent online re-planning of a carrier rocket in a cross-sliding section according to an embodiment of the present application;
fig. 2 is a flowchart of another method for estimating an initial value of thrust descent online re-planning of a carrier rocket in a cross-sliding section according to an embodiment of the present application.
Detailed Description
In order to make the technical solutions and advantages of the embodiments of the present application more apparent, the following further detailed description of the exemplary embodiments of the present application with reference to the accompanying drawings makes it clear that the described embodiments are only a part of the embodiments of the present application, and are not exhaustive of all embodiments. It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict.
In the process of implementing the application, the inventor finds that how to improve the fault adaptability becomes the focus of current research aiming at the flight mission section of the carrier rocket with the gliding section at the stage of entering the orbit.
In view of the above problem, an embodiment of the present application provides a method for estimating an initial value of thrust descent online re-planning of a carrier rocket during a cross-sliding section, where the method includes: dividing a track of a carrier rocket under the condition of thrust descent failure into a first active section, a gliding section and a second active section; determining a flight state sequence of a first active section, a flight state sequence of a taxiing section and a flight state sequence of a second active section based on the flight state of the carrier rocket; and carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment. The method divides the track of the carrier rocket under the condition of thrust descent fault into a first active section, a gliding section and a second active section, and simultaneously considers the flight state sequence of the first active section, the flight state sequence of the gliding section and the flight state sequence of the second active section to carry out re-planning initial value estimation, so that the estimation process is more reasonable, and the convergence and the rapidity of numerical value re-planning are further improved.
Referring to fig. 1, the implementation flow of the method for estimating the initial value of thrust descent online re-planning of the carrier rocket in the cross-sliding section according to the embodiment is as follows:
101, dividing the track of the carrier rocket under the condition of thrust descent fault into a first active section, a gliding section and a second active section.
And the pitching program angle and the yawing program angle of the first active section are constant values.
The height of the far place of the sliding section is the height of the near place of the original target track.
In the second active section, the thrust direction of the carrier rocket is always vertical to the direction of the earth center distance in the orbital plane.
And the carrier rocket enters the second active section when sliding to a far place.
That is to say, in the method of this embodiment, it is assumed that the pitch and yaw process angles of the first active segment are both constant values, the glide track surface to be solved is close to the original glide track surface, and meanwhile, the glide segment apogee height is considered to be the original target track apogee height
And 102, determining a flight state sequence of the first active section, a flight state sequence of the taxiing section and a flight state sequence of the second active section based on the flight state of the carrier rocket.
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 taxiing 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.
1. Implementation process for determining flight state sequence of first active segment based on flight state of carrier rocket
1) And constructing a first active section re-planning problem.
The re-planning of the first active segment after the fault needs to consider the motion equation constraint, the initial state constraint at the fault moment, the terminal constraint of the sliding track and the control quantity constraint, so that the problem of re-planning of the first active segment comprises the following steps: the motion equation, the initial state, the terminal state and the controlled variable constraint.
(1) Equation of motion
Namely constraint of motion equation, which is a state equation describing the motion of rocket centroid under the launching inertia coordinate system. Wherein, the origin O of the emission inertial coordinate system (called as the emission inertial system for short) is at the emission point, the OX axis points to the emission direction in the horizontal plane, the OY axis points to the sky perpendicular to the local horizontal plane of the emission point, and the OZ axis meets the right-hand rule.
Specifically, the motion equation is:
Figure BDA0003430835140000081
T＝Isp dm。
Figure BDA0003430835140000082
wherein, for the operator of first derivative, P is position vector, V is velocity vector, T is thrust of engine after fault, m is mass, u is thrust direction, μ is constant coefficient of earth gravity,r is the distance from the rocket center of mass to the geocentric, r is the component of the geocentric distance under the launching inertia coordinate system, dm is the second flow after the fault, IspIs the specific impulse of the engine,
Figure BDA0003430835140000083
is the pitch program angle of the first active segment and psi is the yaw program angle of the first active segment.
Thus, the state quantity is [ P, V, m]TThe control quantity is
Figure BDA0003430835140000084
(2) Initial state
I.e. the initial state constraint at the moment of failure.
If the fault time is defined as t0Corresponding to an initial state of [ P0,V0,m0]TThen the initial state is:
[P,V,m]T(t0)＝[P0,V0,m0]T。
where P is the position vector, V is the velocity vector, m is the mass, P0Is a position vector of the moment of failure, V0Is the velocity vector at the moment of failure, m0Is the quality of the moment of failure.
(3) Terminal state
I.e. the end constraint of the glide track.
If the terminal time of the first active segment is defined as tf1The number of the sliding tracks corresponding to the speed and the position of the terminal needs to satisfy the following constraints,
[af1,ef1,if1,Ωf1,wf1,ff1]T＝Funorbit(P(tf1),V(tf1))。
Figure BDA0003430835140000085
wherein, af1Is the terminal time orbit semi-major axis, e, of the first active segmentf1Is the terminal moment track eccentricity of the first active segment if1Is the terminal time orbit inclination angle, omega, of the first active segmentf1An ascending node longitude, w, of the terminal time orbit of the first active segmentf1Is the terminal time orbit near-place argument, f of the first active segmentf1Is the terminal time true approach angle, Fun, of the first active segmentorbit() Is a conversion relation function between the number of the tracks of the first active section and the position and the speed under a launching inertia coordinate system, tf1Is the terminal time of the first active segment, P (t)f1) Is the position vector of the terminal moment of the first active segment, V (t)f1) Is the velocity vector, ha, of the terminal instant of the first active segmentf1The track apogee height at the terminal time, rBIs the telecentric distance of the sliding track, ReIs the radius of the earth, nTAre variables to be solved.
If the close center distance of the sliding track is defined as rAThen, then
Figure BDA0003430835140000091
In addition, if the speed of the original target track at the near point is defined as VBThe speed of the carrier rocket when sliding to the remote place of the sliding track is VTBThen, then
Figure BDA0003430835140000092
Then, the velocity increment DeltaV of the second active section orbit change is implemented by the carrier rocket at the long-distance site of the gliding sectionB＝VB-VTB，
Figure BDA0003430835140000093
Remaining time of flight of the first active segment
Figure BDA0003430835140000094
Wherein m is0M is the quality of the moment of failuresIs the structural mass, mloadIs a payloadMass, VBIs the near-to-earth velocity of the original target orbit, mu is the constant coefficient of the earth's gravity, tf2Is the flight time of the second active segment.
Suppose the velocity increment Δ V required for the second active section to change trackBProvided by engine thrust, the flight duration of the second active segment
Figure BDA0003430835140000095
Thus, the terminal constraint may be denoted as nTThe function of (2), namely the terminal state, is:
[af1,ef1,if1,Ωf1,wf1,ff1]T＝Funorbit(P(tf1),V(tf1))。
Figure BDA0003430835140000096
Figure BDA0003430835140000097
|if1-iref|≤εi,|Ωf1-Ωref|≤εΩ。
wherein, af1Is the terminal time orbit semi-major axis, e, of the first active segmentf1Is the terminal moment track eccentricity of the first active segment if1Is the terminal time orbit inclination angle, omega, of the first active segmentf1An ascending node longitude, w, of the terminal time orbit of the first active segmentf1Is the terminal time orbit near-place argument, f of the first active segmentf1Is the terminal time true approach angle, Fun, of the first active segmentorbit() Is a conversion relation function between the number of the tracks of the first active section and the position and the speed under a launching inertia coordinate system, tf1Is the terminal time of the first active segment, P (t)f1) Is the position vector of the terminal moment of the first active segment, V (t)f1) Is the first active sectionVelocity vector of terminal time of (ha)f1The track apogee height at the terminal time, rBIs the telecentric distance of the sliding track, ReIs the radius of the earth, nTFor the variable to be solved, m0M is the quality of the moment of failuresIs the structural mass, mloadFor payload mass, VBIs the near-to-earth velocity of the original target orbit, mu is the constant coefficient of the earth's gravity, IspFor engine specific impulse, irefIs the track inclination of the target track, epsiloniIs the maximum value of the track inclination deviation (i.e. the allowable value of the track inclination deviation), omegarefIs the elevation point longitude, epsilon, of the target trackΩIs the maximum value of the elevation point longitude deviation (i.e., the allowable value of the elevation point longitude deviation).
(4) Control quantity constraint
To simplify the re-planning problem of the first active segment, a first active segment pitch program angle is defined
Figure BDA0003430835140000101
And the yaw program angle ψ are both constant values, the control amount constraint can be expressed as:
Figure BDA0003430835140000102
wherein t is any time of the first active segment,
Figure BDA0003430835140000103
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.
In summary, if the control variable is defined as
Figure BDA0003430835140000104
The first active segment re-planning problem may be represented as:
equation of motion:
Figure BDA0003430835140000105
Figure BDA0003430835140000106
initial state: [ P, V, m ]]T(t0)＝[P0,V0,m0]T，
The terminal state: [ a ] Af1,ef1,if1,Ωf1,wf1,ff1]T＝Funorbit(P(tf1),V(tf1))，
Figure BDA0003430835140000107
Figure BDA0003430835140000111
|if1-iref|≤εi，|Ωf1-Ωref|≤εΩ。
And (3) controlling quantity constraint:
Figure BDA0003430835140000112
2) solving the re-planning problem of the first active segment by adopting a nonlinear planning method to obtain a flight state sequence X of the first active segment1And a sequence of control quantities U1。
Wherein, U1The control quantity of each discrete point of the first active section is
Figure BDA0003430835140000113
Figure BDA0003430835140000114
For the pitch program angle of the first active segment at the moment of failure, psi0Is the yaw program angle of the first active segment at the time of failure]TIs a transpose operation.
Aiming at the step 1), describing the re-planning problem of the first active segment, solving by adopting a non-linear planning method (such as an interior point method, a sequence quadratic planning and the like), and rapidly converging to a feasible solution under the condition of not considering an objective function so as to obtain a flight state quantity sequence X of the first active segment1The control quantity at each discrete point is
Figure BDA0003430835140000115
The corresponding sequence of control quantities may be denoted as U1。
2. Implementation process for determining flight state sequence of taxiing section based on flight state of carrier rocket
1) And determining the number of the orbits in the gliding section according to the terminal state of the carrier rocket at the moment of entering the gliding orbit.
Specifically, the number of the sliding section rails is
Figure BDA0003430835140000116
Wherein the content of the first and second substances,
Figure BDA0003430835140000117
is a semi-long axis of the track of the sliding section,
Figure BDA0003430835140000118
is the track eccentricity of the sliding section,
Figure BDA0003430835140000119
is the track inclination angle of the sliding section,
Figure BDA00034308351400001110
the point longitude is incremented for the trajectory of the skid section,
Figure BDA00034308351400001111
is the orbital perigee argument of the taxiing section,
Figure BDA00034308351400001112
is the true periorbital angle of the skid section.
2) Determining speed vector of each discrete point of taxiing section based on number of tracks of taxiing section
Figure BDA00034308351400001113
And position vector
Figure BDA00034308351400001114
Wherein j is a taxiing section discrete point identifier, j is 1, …, N +1 is the total number of discrete points of the taxiing section, and N is the number of sections equally divided by the true approximate point angle of the taxiing section.
In addition, the velocity vector of each discrete point of the taxiing section
Figure BDA00034308351400001115
And position vector
Figure BDA00034308351400001116
The following relationship is satisfied:
Figure BDA00034308351400001117
wherein, FunPV() Is a conversion relation function between the number of the tracks of the sliding section and the position and the speed under a launching inertia coordinate system,
Figure BDA0003430835140000121
3) the time of each discrete point of the taxiing section is determined.
Specifically, the time of each discrete point of the taxiing section satisfies the following relation:
Figure BDA0003430835140000122
Figure BDA0003430835140000123
wherein, tf1Is the first initiativeThe terminal time of the segment, j 'is the non-first discrete point identification of the taxiing segment, j' is 2, …, N +1,
Figure BDA0003430835140000124
is the approximate point angle of the discrete point j' of the taxiing section, mu is the earth gravity constant coefficient.
4) According to the velocity vector of each discrete point of the fixed sliding section
Figure BDA0003430835140000125
And position vector
Figure BDA0003430835140000126
Obtaining flight state sequence X of the taxiing section according to time of each discrete point of the taxiing sectioncAnd a sequence of control quantities Uc。
Wherein, UcThe control quantity of each discrete point of the sliding section is 0.
When determining the flight state sequence of the gliding section, the corresponding gliding section orbit number can be obtained according to the terminal state of the rocket entering the gliding orbit
Figure BDA0003430835140000127
As the rocket taxis to the far point, the true near point angle gradually becomes larger, and when 180 degrees is reached, the rocket reaches the far point, and the taxiing section is ended. Defining the total quantity of discrete points of the taxiing section as N +1, and averagely dividing the true proximal angle into N intervals (namely N is the quantity of the intervals divided by the true proximal angle of the taxiing section), wherein the interval of each interval is dfcThen the true paraxial angular sequence can be expressed as:
Figure BDA0003430835140000128
Figure BDA0003430835140000129
based on the number of tracks in the sliding section and the position and speed of the launching inertia coordinate systemFun function of the relationship between the twoPV() The velocity vector corresponding to each discrete point can be obtained
Figure BDA00034308351400001210
And position vector
Figure BDA00034308351400001211
As will be shown below, in the following,
Figure BDA00034308351400001212
the time corresponding to each discrete point can be calculated using the following equation,
Figure BDA0003430835140000131
Figure BDA0003430835140000132
defining a sequence of state quantities of taxis as Xc. Control quantity sequence U due to no thrust actioncAll elements in (1) take zero.
3. Implementation process for determining flight state sequence of second active segment based on flight state of carrier rocket
In the second active stage, it is necessary to keep the thrust direction u always perpendicular to the ground center distance direction r in the track plane, that is, u · r is 1. Wherein the content of the first and second substances,
Figure BDA0003430835140000133
r is the component of the earth's center distance in the transmit inertial frame.
1) And determining the time interval and the initial state of the second active segment according to the remaining flight time of the carrier rocket.
In particular, the time interval of the second active segment
Figure BDA0003430835140000134
Wherein, tf2M is the total number of discrete points of the second active segment-1, which is the remaining flight time of the launch vehicle.
In addition, the initial state of the second active section is the terminal state of the sliding section.
2) Calculating the flight state quantity sequence X of the second active segment by numerical integration based on the time interval and the initial state of the second active segment2And a sequence of control quantities U2。
In the implementation process of determining the flight state sequence of the second active segment, the residual flight time t is determinedf2Defining the second active segment to take (M +1) discrete points, and
Figure BDA0003430835140000135
at intervals of time, in the terminal state of the sliding section
Figure BDA0003430835140000136
For the initial state of the second active segment, the flight state quantity sequence X of the second active segment can be calculated by adopting a numerical integration mode (Euler integration, Longge Kutta integration and the like)2And a sequence of control quantities U2。
Wherein the content of the first and second substances,
Figure BDA0003430835140000137
as a result of the taxiing section terminal position vector,
Figure BDA0003430835140000138
for the terminal velocity vector of the taxiing section,
Figure BDA0003430835140000139
is the terminal mass of the skid section.
103, carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment.
After step 102 is executed, a flight state sequence of the first active segment, a flight state sequence of the taxiing segment, and a flight state sequence of the second active segment are obtained, and in this step, the flight state sequence of the first active segment, the flight state sequence of the taxiing segment, and the flight state sequence of the second active segment are linked (for example, the state quantity sequence and the control quantity sequence of each segment are linked), so that initial guesses of three flight segments corresponding to the online re-planning problem across the taxiing segment can be obtained, and the problem can be solved quickly.
In the embodiment, aiming at the flight mission section of the carrier rocket with the glide section at the orbit entering stage, in order to improve the fault adaptability, the first active section, the glide section and the second active section need to be planned in an online simultaneous manner according to the flight state of the carrier rocket under the condition of thrust descent fault, and the effective load is sent to the original target orbit. The method for estimating the initial value of thrust descent on-line re-planning of the carrier rocket in the cross-sliding section enables the iterative solution process of the numerical optimization method to be fast converged through reasonable initial guess of estimation optimization problems.
In concrete implementation, a sliding track optimization problem is constructed, and a numerical iteration method is utilized to quickly search a feasible solution as an initial guess of a first active segment. And then the rocket enters a second active section when sliding to a far place, so that the initial guess of the sliding section is obtained according to the conversion relation between the number of the tracks and the speed and the position. After entering the second active section, the thrust direction is kept to be always vertical to the earth center distance direction in the track surface, and the initial guess of the second active section is obtained by utilizing numerical integration. And finally, combining the initial values estimated by the three flight sections to serve as an initial guess of the online trajectory planning problem of the carrier rocket in the cross-sliding section so as to realize rapid solution.
Taking fig. 2 as an example, the method provided in this embodiment constructs a first active segment re-planning problem, and then iteratively solves the first active segment flight state sequence based on the problem. Calculating a taxiing section flight state sequence, calculating a second active section flight state sequence, and further constructing an initial guess of the online re-planning problem of the cross-taxiing section according to the first active section flight state sequence, the taxiing section flight state sequence and the second active section flight state sequence so as to estimate the re-planning initial value.
The method for estimating the initial value of the carrier rocket cross-sliding section thrust descent online re-planning provided by the embodiment combines the motion characteristics of the rocket in-orbit process, decomposes the cross-sliding section online planning problem into three flight sections for independent calculation, so that the initial guess meets the motion equation constraint, and the numerical solving difficulty is reduced.
In addition, the method for estimating the initial value of thrust descent online re-planning of the carrier rocket in the cross-sliding section simplifies the form of the control variable of the first active section into a constant value, reduces the dimension of the variable to be solved, and constructs a re-planning problem of the first active section which is easy to solve iteratively.
In addition, the method for estimating the initial value of thrust descent online re-planning of the carrier rocket in the cross-sliding section is designed according to the orbit transfer theory, and a sliding section and a second active section motion state estimation method which do not need numerical value iterative solution are designed, so that the initial guess rationality is met, and the generation speed of the initial guess is increased.
The application provides a method for estimating an initial value of thrust descent online re-planning of a carrier rocket in a cross-sliding section, which comprises the following steps: dividing a track of a carrier rocket under the condition of thrust descent failure into a first active section, a gliding section and a second active section; determining a flight state sequence of a first active section, a flight state sequence of a taxiing section and a flight state sequence of a second active section based on the flight state of the carrier rocket; and carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment. The method divides the track of the carrier rocket under the condition of thrust descent fault into a first active section, a gliding section and a second active section, and simultaneously considers the flight state sequence of the first active section, the flight state sequence of the gliding section and the flight state sequence of the second active section to carry out re-planning initial value estimation, so that the estimation process is more reasonable, and the convergence and the rapidity of numerical value re-planning are further improved.
Based on the same inventive concept of the method for estimating the initial value of the thrust descent online re-planning of the cross-sliding section of the carrier rocket, the embodiment provides electronic equipment, which comprises: memory, processor, and computer programs.
Wherein a computer program is stored in the memory and configured to be executed by the processor to implement the method for estimating an initial value of thrust descent online re-planning of a carrier rocket across taxis as described above with reference to figure 1.
In particular, the method comprises the following steps of,
the trajectory under the condition of the thrust descent fault of the carrier rocket is divided into a first active section, a gliding section and a second active section.
And determining the flight state sequence of the first active section, the flight state sequence of the taxiing section and the flight state sequence of the second active section based on the flight state of the carrier rocket.
And carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment.
Optionally, the pitch program angle and the yaw program angle of the first active segment are both constant.
The height of the far place of the sliding section is the height of the near place of the original target track.
In the second active section, the thrust direction of the carrier rocket is always vertical to the direction of the earth center distance in the orbital plane.
Wherein the carrier rocket enters the second active section when sliding to a remote site.
Optionally, determining a flight state sequence of the first active segment based on the flight state of the launch vehicle comprises:
and constructing a first active section re-planning problem. The first active segment re-planning problem includes: the motion equation, the initial state, the terminal state and the controlled variable constraint.
Solving the re-planning problem of the first active segment by adopting a nonlinear planning method to obtain a flight state sequence X of the first active segment1And a sequence of control quantities U1。
Wherein, U1The control quantity of each discrete point of the first active section is
Figure BDA0003430835140000161
Figure BDA0003430835140000162
For the pitch program angle of the first active segment at the moment of failure, psi0Is the yaw program angle of the first active segment at the time of failure]TIs a transpose operation.
Optionally, the equation of motion is:
Figure BDA0003430835140000163
T＝Ispdm。
Figure BDA0003430835140000164
wherein, for an operator for solving a first derivative, P is a position vector, V is a velocity vector, T is engine thrust after a fault, m is mass, u is a thrust direction, mu is an earth gravity constant coefficient, r is a distance from a rocket centroid to an earth center, r is a component of the earth center distance under a launching inertia coordinate system, dm is a second flow after the fault, IspIs the specific impulse of the engine,
Figure BDA0003430835140000165
is the pitch program angle of the first active segment and psi is the yaw program angle of the first active segment.
Optionally, in the emission inertial coordinate system, the origin O is at the emission point, the OX axis points to the emission direction in the horizontal plane, the OY axis points to the sky in the horizontal plane, and the OZ axis satisfies the right-hand rule.
Optionally, the initial state is:
[P,V,m]T(t0)＝[P0,V0,m0]T。
where P is the position vector, V is the velocity vector, m is the mass, t0To the moment of failure, P0Is a position vector of the moment of failure, V0Is the velocity vector at the moment of failure, m0Is the quality of the moment of failure.
Optionally, the terminal state is:
[af1,ef1,if1,Ωf1,wf1,ff1]T＝Funorbit(P(tf1),V(tf1))。
Figure BDA0003430835140000171
Figure BDA0003430835140000172
|if1-iref|≤εi,|Ωf1-Ωref|≤εΩ。
wherein, af1Is the terminal time orbit semi-major axis, e, of the first active segmentf1Is the terminal moment track eccentricity of the first active segment if1Is the terminal time orbit inclination angle, omega, of the first active segmentf1An ascending node longitude, w, of the terminal time orbit of the first active segmentf1Is the terminal time orbit near-place argument, f of the first active segmentf1Is the terminal time true approach angle, Fun, of the first active segmentorbit() Is a conversion relation function between the number of the tracks of the first active section and the position and the speed under a launching inertia coordinate system, tf1Is the terminal time of the first active segment, P (t)f1) Is the position vector of the terminal moment of the first active segment, V (t)f1) Is the velocity vector, ha, of the terminal instant of the first active segmentf1The track apogee height at the terminal time, rBIs the telecentric distance of the sliding track, ReIs the radius of the earth, nTFor the variable to be solved, m0M is the quality of the moment of failuresIs the structural mass, mloadFor payload mass, VBIs the near-to-earth velocity of the original target orbit, mu is the constant coefficient of the earth's gravity, IspFor engine specific impulse, irefIs the track inclination of the target track, epsiloniMaximum value of track inclination deviation, omegarefIs the elevation point longitude, epsilon, of the target trackΩIs a Chinese traditional patent medicine ofMaximum value of point longitude deviation.
Alternatively,
Figure BDA0003430835140000173
wherein r isAThe close center distance of the sliding track.
Optionally, the control quantity is constrained to:
Figure BDA0003430835140000174
wherein t is any time of the first active segment,
Figure BDA0003430835140000175
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.
Optionally, determining a sequence of flight states of the taxiing section based on the flight state of the launch vehicle comprises:
and determining the number of the orbits in the gliding section according to the terminal state of the carrier rocket at the moment of entering the gliding orbit.
Determining speed vector of each discrete point of taxiing section based on number of tracks of taxiing section
Figure BDA0003430835140000181
And position vector
Figure BDA0003430835140000182
Wherein j is a taxiing section discrete point identifier, j is 1, …, N +1 is the total number of discrete points of the taxiing section, and N is the number of sections equally divided by the true approximate point angle of the taxiing section.
The time of each discrete point of the taxiing section is determined.
According to the velocity vector of each discrete point of the fixed sliding section
Figure BDA0003430835140000183
And position vector
Figure BDA0003430835140000184
Obtaining flight state sequence X of the taxiing section according to time of each discrete point of the taxiing sectioncAnd a sequence of control quantities Uc。
Wherein, UcThe control quantity of each discrete point of the sliding section is 0.
Optionally, the number of sliding section rails is
Figure BDA0003430835140000185
Wherein the content of the first and second substances,
Figure BDA0003430835140000186
is a semi-long axis of the track of the sliding section,
Figure BDA0003430835140000187
is the track eccentricity of the sliding section,
Figure BDA0003430835140000188
is the track inclination angle of the sliding section,
Figure BDA0003430835140000189
the point longitude is incremented for the trajectory of the skid section,
Figure BDA00034308351400001810
is the orbital perigee argument of the taxiing section,
Figure BDA00034308351400001811
is the true periorbital angle of the skid section.
Optionally, the velocity vector of each discrete point of the taxiing section
Figure BDA00034308351400001812
And position vector
Figure BDA00034308351400001813
The following relationship is satisfied:
Figure BDA00034308351400001814
wherein, FunPV() Is a conversion relation function between the number of the tracks of the sliding section and the position and the speed under a launching inertia coordinate system,
Figure BDA00034308351400001815
optionally, the time of each discrete point of the taxiing section satisfies the following relationship:
Figure BDA00034308351400001816
Figure BDA00034308351400001817
wherein, tf1The terminal time of the first active segment, j 'is the identification of the non-first discrete point of the skid segment, j' is 2, …, N +1,
Figure BDA00034308351400001818
is the approximate point angle of the discrete point j' of the taxiing section, mu is the earth gravity constant coefficient.
Optionally, determining a flight state sequence of the second active segment based on the flight state of the launch vehicle comprises:
and determining the time interval and the initial state of the second active segment according to the remaining flight time of the carrier rocket.
Calculating the flight state quantity sequence X of the second active segment by numerical integration based on the time interval and the initial state of the second active segment2And a sequence of control quantities U2。
Optionally, the time interval of the second active segment
Figure BDA0003430835140000191
Wherein, tf2M is the total number of discrete points of the second active segment-1, which is the remaining flight time of the launch vehicle.
Optionally, the initial state of the second active segment is a terminal state of the skid segment.
The electronic device provided by the embodiment is provided with a computer program executed by a processor to divide the track under the condition of the thrust descent fault of the carrier rocket into a first active section, a gliding section and a second active section; determining a flight state sequence of a first active section, a flight state sequence of a taxiing section and a flight state sequence of a second active section based on the flight state of the carrier rocket; and carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment. In the embodiment, the trajectory of the carrier rocket under the condition of thrust descent fault is divided into the first active section, the gliding section and the second active section, and the flight state sequence of the first active section, the flight state sequence of the gliding section and the flight state sequence of the second active section are considered for re-planning initial value estimation, so that the estimation process is more reasonable, and the convergence and the rapidity of numerical value re-planning are further improved.
Based on the same inventive concept of the method for estimating the initial value of thrust descent online re-planning of the cross-sliding section of the carrier rocket, the embodiment provides a computer on which a computer program is stored. The computer program is executed by a processor to implement the method for estimating the initial value of thrust descent online re-planning of the carrier rocket across the taxiing section shown in fig. 1.
In particular, the method comprises the following steps of,
the trajectory under the condition of the thrust descent fault of the carrier rocket is divided into a first active section, a gliding section and a second active section.
And determining the flight state sequence of the first active section, the flight state sequence of the taxiing section and the flight state sequence of the second active section based on the flight state of the carrier rocket.
And carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment.
Optionally, the pitch program angle and the yaw program angle of the first active segment are both constant.
The height of the far place of the sliding section is the height of the near place of the original target track.
In the second active section, the thrust direction of the carrier rocket is always vertical to the direction of the earth center distance in the orbital plane.
Wherein the carrier rocket enters the second active section when sliding to a remote site.
Optionally, determining a flight state sequence of the first active segment based on the flight state of the launch vehicle comprises:
and constructing a first active section re-planning problem. The first active segment re-planning problem includes: the motion equation, the initial state, the terminal state and the controlled variable constraint.
Solving the re-planning problem of the first active segment by adopting a nonlinear planning method to obtain a flight state sequence X of the first active segment1And a sequence of control quantities U1。
Wherein, U1The control quantity of each discrete point of the first active section is
Figure BDA0003430835140000201
Figure BDA0003430835140000202
For the pitch program angle of the first active segment at the moment of failure, psi0Is the yaw program angle of the first active segment at the time of failure]TIs a transpose operation.
Optionally, the equation of motion is:
Figure BDA0003430835140000203
T＝Ispdm。
Figure BDA0003430835140000204
where, for the operator to find the first derivative, P is the position vector, V is the velocity vector, T is the post-fault engine thrust, and m is the primeThe quantity u is the thrust direction, mu is the constant coefficient of the earth gravity, r is the distance from the center of mass of the rocket to the center of earth, r is the component of the center of earth distance under the launching inertia coordinate system, dm is the second flow after the fault, IspIs the specific impulse of the engine,
Figure BDA0003430835140000205
is the pitch program angle of the first active segment and psi is the yaw program angle of the first active segment.
Optionally, in the emission inertial coordinate system, the origin O is at the emission point, the OX axis points to the emission direction in the horizontal plane, the OY axis points to the sky in the horizontal plane, and the OZ axis satisfies the right-hand rule.
Optionally, the initial state is:
[P,V,m]T(t0)＝[P0,V0,m0]T。
where P is the position vector, V is the velocity vector, m is the mass, t0To the moment of failure, P0Is a position vector of the moment of failure, V0Is the velocity vector at the moment of failure, m0Is the quality of the moment of failure.
Optionally, the terminal state is:
[af1，ef1,if1,Ωf1,wf1,ff1]T＝Funorbit(P(tf1),V(tf1))。
Figure BDA0003430835140000211
Figure BDA0003430835140000212
|if1-iref|≤εi,|Ωf1-Ωref|≤εΩ。
wherein, af1Is the terminal time orbit semi-major axis, e, of the first active segmentf1Is a terminal time track of the first active segmentEccentricity, if1Is the terminal time orbit inclination angle, omega, of the first active segmentf1An ascending node longitude, w, of the terminal time orbit of the first active segmentf1Is the terminal time orbit near-place argument, f of the first active segmentf1Is the terminal time true approach angle, Fun, of the first active segmentorbit() Is a conversion relation function between the number of the tracks of the first active section and the position and the speed under a launching inertia coordinate system, tf1Is the terminal time of the first active segment, P (t)f1) Is the position vector of the terminal moment of the first active segment, V (t)f1) Is the velocity vector, ha, of the terminal instant of the first active segmentf1The track apogee height at the terminal time, rBIs the telecentric distance of the sliding track, ReIs the radius of the earth, nTFor the variable to be solved, m0M is the quality of the moment of failuresIs the structural mass, mloadFor payload mass, VBIs the near-to-earth velocity of the original target orbit, mu is the constant coefficient of the earth's gravity, IspFor engine specific impulse, irefIs the track inclination of the target track, epsiloniMaximum value of track inclination deviation, omegarefIs the elevation point longitude, epsilon, of the target trackΩIs the maximum value of the elevation point longitude deviation.
Alternatively,
Figure BDA0003430835140000213
wherein r isAThe close center distance of the sliding track.
Optionally, the control quantity is constrained to:
Figure BDA0003430835140000214
wherein t is any time of the first active segment,
Figure BDA0003430835140000221
is the pitch program angle of the first active segment at time t, psi (t) is the pitch program angle of the first active segment at time tYaw program angle of moment.
Optionally, determining a sequence of flight states of the taxiing section based on the flight state of the launch vehicle comprises:
and determining the number of the orbits in the gliding section according to the terminal state of the carrier rocket at the moment of entering the gliding orbit.
Determining speed vector of each discrete point of taxiing section based on number of tracks of taxiing section
Figure BDA0003430835140000222
And position vector
Figure BDA0003430835140000223
Wherein j is a taxiing section discrete point identifier, j is 1, …, N +1 is the total number of discrete points of the taxiing section, and N is the number of sections equally divided by the true approximate point angle of the taxiing section.
The time of each discrete point of the taxiing section is determined.
According to the velocity vector of each discrete point of the fixed sliding section
Figure BDA0003430835140000224
And position vector
Figure BDA0003430835140000225
Obtaining flight state sequence X of the taxiing section according to time of each discrete point of the taxiing sectioncAnd a sequence of control quantities Uc。
Wherein, UcThe control quantity of each discrete point of the sliding section is 0.
Optionally, the number of sliding section rails is
Figure BDA0003430835140000226
Wherein the content of the first and second substances,
Figure BDA0003430835140000227
is a semi-long axis of the track of the sliding section,
Figure BDA0003430835140000228
is the track eccentricity of the sliding section,
Figure BDA0003430835140000229
is the track inclination angle of the sliding section,
Figure BDA00034308351400002210
the point longitude is incremented for the trajectory of the skid section,
Figure BDA00034308351400002211
is the orbital perigee argument of the taxiing section,
Figure BDA00034308351400002212
is the true periorbital angle of the skid section.
Optionally, the velocity vector of each discrete point of the taxiing section
Figure BDA00034308351400002213
And position vector
Figure BDA00034308351400002214
The following relationship is satisfied:
Figure BDA00034308351400002215
wherein, FunPV() Is a conversion relation function between the number of the tracks of the sliding section and the position and the speed under a launching inertia coordinate system,
Figure BDA00034308351400002216
optionally, the time of each discrete point of the taxiing section satisfies the following relationship:
Figure BDA00034308351400002217
Figure BDA00034308351400002218
wherein, tf1The terminal time of the first active segment, j 'is the identification of the non-first discrete point of the skid segment, j' is 2, …, N +1,
Figure BDA0003430835140000231
is the approximate point angle of the discrete point j' of the taxiing section, mu is the earth gravity constant coefficient.
Optionally, determining a flight state sequence of the second active segment based on the flight state of the launch vehicle comprises:
and determining the time interval and the initial state of the second active segment according to the remaining flight time of the carrier rocket.
Calculating the flight state quantity sequence X of the second active segment by numerical integration based on the time interval and the initial state of the second active segment2And a sequence of control quantities U2。
Optionally, the time interval of the second active segment
Figure BDA0003430835140000232
Wherein, tf2M is the total number of discrete points of the second active segment-1, which is the remaining flight time of the launch vehicle.
Optionally, the initial state of the second active segment is a terminal state of the skid segment.
The present embodiment provides a computer-readable storage medium having a computer program thereon executed by a processor to divide a trajectory in the event of a thrust descent fault of a launch vehicle into a first active segment, a gliding segment, and a second active segment; determining a flight state sequence of a first active section, a flight state sequence of a taxiing section and a flight state sequence of a second active section based on the flight state of the carrier rocket; and carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment. In the embodiment, the trajectory of the carrier rocket under the condition of thrust descent fault is divided into the first active section, the gliding section and the second active section, and the flight state sequence of the first active section, the flight state sequence of the gliding section and the flight state sequence of the second active section are considered for re-planning initial value estimation, so that the estimation process is more reasonable, and the convergence and the rapidity of numerical value re-planning are further improved.
As will be appreciated by one skilled in the art, embodiments of the present application may be provided as a method, system, or computer program product. Accordingly, the present application may take the form of an entirely hardware embodiment, an entirely software embodiment or an embodiment combining software and hardware aspects. Furthermore, the present application may take the form of a computer program product embodied on one or more computer-usable storage media (including, but not limited to, disk storage, CD-ROM, optical storage, and the like) having computer-usable program code embodied therein. The scheme in the embodiment of the application can be implemented by adopting various computer languages, such as object-oriented programming language Java and transliterated scripting language JavaScript.
The present application is described with reference to flowchart illustrations and/or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the application. It will be understood that each flow and/or block of the flow diagrams and/or block diagrams, and combinations of flows and/or blocks in the flow diagrams and/or block diagrams, can be implemented by computer program instructions. These computer program instructions may be provided to a processor of a general purpose computer, special purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, create means for implementing the functions specified in the flowchart flow or flows and/or block diagram block or blocks.
These computer program instructions may also be stored in a computer-readable memory that can direct a computer or other programmable data processing apparatus to function in a particular manner, such that the instructions stored in the computer-readable memory produce an article of manufacture including instruction means which implement the function specified in the flowchart flow or flows and/or block diagram block or blocks.
These computer program instructions may also be loaded onto a computer or other programmable data processing apparatus to cause a series of operational steps to be performed on the computer or other programmable apparatus to produce a computer implemented process such that the instructions which execute on the computer or other programmable apparatus provide steps for implementing the functions specified in the flowchart flow or flows and/or block diagram block or blocks.
While the preferred embodiments of the present application have been described, additional variations and modifications in those embodiments may occur to those skilled in the art once they learn of the basic inventive concepts. Therefore, it is intended that the appended claims be interpreted as including preferred embodiments and all alterations and modifications as fall within the scope of the application.
It will be apparent to those skilled in the art that various changes and modifications may be made in the present application without departing from the spirit and scope of the application. Thus, if such 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 intended to include such modifications and variations as well.

Claims (18)

1. A method for estimating an initial value of thrust descent online re-planning of a carrier rocket in a cross-sliding section is characterized by comprising the following steps:
dividing a track of a carrier rocket under the condition of thrust descent failure into a first active section, a gliding section and a second active section;
determining a flight state sequence of the first active segment, a flight state sequence of the taxiing segment, and a flight state sequence of the second active segment based on a flight state of a launch vehicle;
and carrying out re-planning initial value estimation according to the flight state sequence of the first active segment, the flight state sequence of the taxiing segment and the flight state sequence of the second active segment.
2. The method of claim 1, wherein the pitch program angle and yaw program angle of the first active segment are both constant;
the height of the far place of the taxiing section is the height of the near place of the original target track;
in the second active section, the thrust direction of the carrier rocket is always vertical to the direction of the earth center distance in the orbital plane;
wherein the launch vehicle enters the second active segment when taxiing to the remote site.
3. The method of claim 2, wherein determining the sequence of flight states for the first active segment based on the flight state of the launch vehicle comprises:
constructing a first active section re-planning problem; the first active segment re-planning problem comprises: motion equation, initial state, terminal state and controlled variable constraint;
solving the re-planning problem of the first active segment by adopting a nonlinear planning method to obtain a flight state sequence X of the first active segment1And a sequence of control quantities U1；
Wherein, U1The control quantity of each discrete point of the first active section is
Figure FDA0003430835130000011
Figure FDA0003430835130000012
For the pitch program angle, ψ, of the first active segment at the moment of failure0Is a yaw program angle of the first active segment at a failure time]TIs a transpose operation.
4. The method of claim 3, wherein the equation of motion is:
Figure FDA0003430835130000013
T＝Ispdm；
Figure FDA0003430835130000021
wherein, for an operator for solving a first derivative, P is a position vector, V is a velocity vector, T is engine thrust after a fault, m is mass, u is a thrust direction, mu is an earth gravity constant coefficient, r is a distance from a rocket centroid to an earth center, r is a component of the earth center distance under a launching inertia coordinate system, dm is a second flow after the fault, IspIs the specific impulse of the engine,
Figure FDA0003430835130000022
is the pitch program angle of the first active segment and psi is the yaw program angle of the first active segment.
5. The method of claim 4, wherein in the emission inertial coordinate system, an origin O is at an emission point, an OX axis points in a horizontal plane in the emission direction, an OY axis points perpendicular to the local horizontal plane of the emission point in the sky, and an OZ axis satisfies a right-hand rule.
6. The method of claim 3, wherein the initial state is:
[P，V，m]T(t0)＝[P0，V0，m0]T；
where P is the position vector, V is the velocity vector, m is the mass, t0To the moment of failure, P0Is a position vector of the moment of failure, V0Is the velocity vector at the moment of failure, m0Is the quality of the moment of failure.
7. The method of claim 3, wherein the terminal state is:
[af1，ef1，if1，Ωf1，wf1，ff1]T＝Funorbit(P(tf1)，V(tf1))；
haf1＝rB-Re，
Figure FDA0003430835130000024
Figure FDA0003430835130000023
|if1-iref|≤εi，|Ωf1-Ωref|≤εΩ；
wherein, af1Is the terminal time orbit semi-major axis, e, of the first active segmentf1Is the terminal moment track eccentricity of the first active segment if1Is the terminal time orbit inclination angle, omega, of the first active segmentf1An ascending node longitude, w, of the terminal time orbit of the first active segmentf1Is the terminal time orbit near-place argument, f of the first active segmentf1Is the terminal time true approach angle, Fun, of the first active segmentorbit() Is a conversion relation function between the number of the tracks of the first active section and the position and the speed under a launching inertia coordinate system, tf1Is the terminal time of the first active segment, P (t)f1) Is the position vector of the terminal moment of the first active segment, V (t)f1) Is the velocity vector, ha, of the terminal instant of the first active segmentf1The track apogee height at the terminal time, rBIs the telecentric distance of the sliding track, ReIs the radius of the earth, nTFor the variable to be solved, m0M is the quality of the moment of failuresIs the structural mass, mloadFor payload mass, VBIs the near-to-earth velocity of the original target orbit, mu is the constant coefficient of the earth's gravity, IspFor engine specific impulse, irefIs the track inclination of the target track, epsiloniMaximum value of track inclination deviation, omegarefIs the elevation point longitude, epsilon, of the target trackΩIs the maximum value of the elevation point longitude deviation.
8. The method of claim 7,
Figure FDA0003430835130000031
wherein r isAThe close center distance of the sliding track.
9. The method of claim 3, wherein the control quantity constraint is:
Figure FDA0003430835130000032
ψ(t)＝ψ0；
wherein t is any time of the first active segment,
Figure FDA0003430835130000033
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.
10. The method of claim 2, wherein determining the sequence of flight states for the taxiing section based on the flight state of the launch vehicle comprises:
determining the number of orbits of the gliding section according to the terminal state of the carrier rocket at the moment of entering the gliding orbit;
determining a velocity vector of each discrete point of the taxiing section based on the number of the taxiing section tracks
Figure FDA0003430835130000034
And position vector
Figure FDA0003430835130000035
Wherein j is the taxiing section discrete point identifier, j is 1, …, N +1 is the total number of discrete points of the taxiing section, and N is the number of sections bisected by the real approximate point angle of the taxiing section;
determining the time of each discrete point of the taxiing section;
according to the velocity vector of each discrete point of the taxiing section
Figure FDA0003430835130000036
And position vector
Figure FDA0003430835130000037
Obtaining the flight state sequence X of the taxiing section according to the time of each discrete point of the taxiing sectioncAnd a sequence of control quantities Uc；
Wherein, UcThe control quantity of each discrete point of the taxiing section is 0.
11. The method of claim 10 wherein the number of skid rails is
Figure FDA0003430835130000041
Wherein the content of the first and second substances,
Figure FDA0003430835130000042
is the semi-long axis of the track of the sliding section,
Figure FDA0003430835130000043
is the orbital eccentricity of the skid section,
Figure FDA0003430835130000044
is the angle of inclination of the track of the skid section,
Figure FDA0003430835130000045
the point longitude is incremented for the track of the taxiing section,
Figure FDA0003430835130000046
is the orbital perigee argument of the taxiing section,
Figure FDA0003430835130000047
is a track of the gliding sectionAngle of approach.
12. The method of claim 11 wherein the velocity vector of each discrete point of the taxiing segment
Figure FDA0003430835130000048
And position vector
Figure FDA0003430835130000049
The following relationship is satisfied:
Figure FDA00034308351300000410
wherein, FunPV() As a function of the conversion relation between the number of the taxiing section orbit and the position and the speed under the emission inertia coordinate system,
Figure FDA00034308351300000411
13. the method of claim 11 wherein the time at each discrete point of the skid satisfies the relationship:
Figure FDA00034308351300000412
Figure FDA00034308351300000413
wherein, tf1J ' is the terminal time of the first active segment, j ' is the non-first discrete point identification of the skid segment, j ' is 2, …, N +1,
Figure FDA00034308351300000414
is the approximate point angle of the discrete point j' of the taxiing section, mu is the gravitational constantAnd (4) counting.
14. The method of claim 2, wherein determining the sequence of flight states for the second active segment based on the flight state of the launch vehicle comprises:
determining the time interval and the initial state of the second active segment according to the remaining flight time of the carrier rocket;
calculating the flight state quantity sequence X of the second active segment by adopting numerical integration based on the time interval and the initial state of the second active segment2And a sequence of control quantities U2。
15. The method of claim 14, wherein the time interval of the second active segment is
Figure FDA00034308351300000415
Wherein, tf2And M is the total number of the discrete points of the second active section, namely-1, wherein M is the residual flight time of the carrier rocket.
16. The method of claim 14 wherein the initial state of the second active segment is a terminal state of the skid segment.
17. An electronic device, comprising:
a memory;
a processor; and
a computer program;
wherein the computer program is stored in the memory and configured to be executed by the processor to implement the method of any one of claims 1-16.
18. A computer-readable storage medium, having stored thereon a computer program; the computer program is executed by a processor to implement the method of any one of claims 1-16.
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