CN116301008A - Carrier rocket control method, carrier rocket, electronic device and storage medium - Google Patents

Carrier rocket control method, carrier rocket, electronic device and storage medium Download PDF

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Publication number
CN116301008A
CN116301008A CN202310566867.2A CN202310566867A CN116301008A CN 116301008 A CN116301008 A CN 116301008A CN 202310566867 A CN202310566867 A CN 202310566867A CN 116301008 A CN116301008 A CN 116301008A
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time
real
coordinate system
angle
carrier rocket
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CN116301008B (en
Inventor
熊少锋
刘百奇
梅金平
何建华
王振华
孙国伟
刘建设
王博
雷克非
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Beijing Xinghe Power Aerospace Technology Co ltd
Beijing Xinghe Power Equipment Technology Co Ltd
Anhui Galaxy Power Equipment Technology Co Ltd
Galactic Energy Shandong Aerospace Technology Co Ltd
Jiangsu Galatic Aerospace Technology Co Ltd
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Beijing Xinghe Power Aerospace Technology Co ltd
Beijing Xinghe Power Equipment Technology Co Ltd
Anhui Galaxy Power Equipment Technology Co Ltd
Galactic Energy Shandong Aerospace Technology Co Ltd
Jiangsu Galatic Aerospace Technology Co Ltd
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

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  • Aviation & Aerospace Engineering (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)

Abstract

The application discloses a carrier rocket control method, a carrier rocket, electronic equipment and a storage medium, and relates to the technical field of aerospace, wherein the method comprises the following steps: acquiring a target attack angle, a real-time sideslip angle and a real-time velocity vector in a launching coordinate system of a carrier rocket at the current flight time; determining a real-time ballistic inclination angle and a real-time ballistic deflection angle of the carrier rocket at the current flight moment based on the real-time velocity vector; determining a real-time pitch angle and a real-time yaw angle of the carrier rocket at the current flight time based on the real-time trajectory inclination angle, the real-time trajectory deflection angle, the target attack angle, the real-time sideslip angle and the posture association relation of the carrier rocket at the current flight time; and controlling the flight attitude of the carrier rocket. The method and the device provided by the application improve the solving efficiency of the pitch angle and the yaw angle and improve the control response speed of the carrier rocket.

Description

Carrier rocket control method, carrier rocket, electronic device and storage medium
Technical Field
The application relates to the technical field of aerospace, in particular to a carrier rocket control method, a carrier rocket, electronic equipment and a storage medium.
Background
In the development process of the carrier rocket, the ballistic design is an overall design work which needs to be subjected to top-level planning and first research, and provides input for the guidance control system and the rocket body structural system design, so the ballistic design is a very important work. When the carrier rocket flies along the trajectory, a control mode of program angle turning is generally adopted, an instruction attack angle is designed in advance, and then the attitude angle of the rocket in a launching inertial coordinate system is obtained through conversion of the instruction attack angle. The attitude angle includes at least a pitch angle and a yaw angle.
Therefore, how to quickly and accurately determine the pitch angle and yaw angle of the carrier rocket and improve the control response speed of the carrier rocket are technical problems to be solved in the industry.
Disclosure of Invention
The application provides a carrier rocket control method, a carrier rocket, electronic equipment and a storage medium, which are used for solving the technical problem of how to quickly and accurately determine the pitch angle and yaw angle of the carrier rocket and improving the control response speed of the carrier rocket.
The application provides a carrier rocket control method, which comprises the following steps:
acquiring a target attack angle, a real-time sideslip angle and a real-time velocity vector in a launching coordinate system of a carrier rocket at the current flight time;
determining a real-time ballistic inclination angle and a real-time ballistic deflection angle of the carrier rocket at the current flight moment based on the real-time velocity vector;
determining a real-time pitch angle and a real-time yaw angle of the carrier rocket at the current flight time based on the real-time trajectory inclination angle, the real-time trajectory deflection angle, the target attack angle, the real-time sideslip angle and the posture association relation of the carrier rocket at the current flight time;
controlling the flight attitude of the carrier rocket based on the real-time pitch angle and the real-time yaw angle;
the attitude association relation is determined based on an attitude transfer matrix among an launching inertial coordinate system, an launching coordinate system, a speed coordinate system and an rocket body coordinate system of the carrier rocket.
In some embodiments, the gesture association is determined based on the steps of:
determining a first attitude transfer matrix of the launching inertial coordinate system transformed to the launching coordinate system based on the azimuth angle of the carrier rocket, the geographic latitude of the launching point, the current flight time and the rotation angular speed of the earth;
determining a second attitude transfer matrix for transforming the launching coordinate system to the speed coordinate system based on the real-time speed inclination angle, the real-time trajectory deflection angle and the real-time trajectory inclination angle of the carrier rocket at the current flight time;
determining a third attitude transfer matrix of the speed coordinate system transformed to the rocket body coordinate system based on a target attack angle and a real-time sideslip angle of the carrier rocket at the current flight time;
determining a fourth pose transfer matrix for transforming the launching inertial coordinate system to the arrow body coordinate system based on the first pose transfer matrix, the second pose transfer matrix, and the third pose transfer matrix;
determining a fifth attitude transfer matrix for transforming the launch inertial coordinate system to the rocket body coordinate system based on a real-time pitch angle, a real-time yaw angle and a real-time roll angle of the carrier rocket in the launch inertial coordinate system at the current flight time;
and determining the posture association relation based on the fourth posture transfer matrix and the fifth posture transfer matrix.
In some embodiments, the determining the first attitude transfer matrix for transforming the launch inertial coordinate system to the launch coordinate system based on the azimuth of the launch vehicle, the launch point geographical latitude, the current time of flight, and the earth rotation angular velocity includes:
Figure SMS_1
wherein ,
Figure SMS_3
for the emission inertial coordinate system +.>
Figure SMS_8
To the emission coordinate system->
Figure SMS_12
Is a first posture transfer matrix of->
Figure SMS_4
For the azimuth angle>
Figure SMS_9
For the geographical latitude of the emission point +.>
Figure SMS_10
For the current flight time>
Figure SMS_13
For the earth rotation angular velocity, +.>
Figure SMS_2
For representing a surrounding coordinate system +.>
Figure SMS_7
A rotation matrix for rotation of the shaft; />
Figure SMS_11
For representing a surrounding coordinate system +.>
Figure SMS_14
A rotation matrix for rotation of the shaft;
Figure SMS_5
for representing a surrounding coordinate system +.>
Figure SMS_6
A rotation matrix for rotating the shaft.
In some embodiments, the determining a second attitude transfer matrix for transforming the launch coordinate system to the velocity coordinate system based on the real-time velocity tilt, the real-time ballistic deflection, and the real-time ballistic tilt of the launch vehicle at the current time of flight comprises:
Figure SMS_15
wherein ,
Figure SMS_16
for the emission coordinate system->
Figure SMS_17
To the velocity coordinate system->
Figure SMS_18
Is a second posture transfer matrix of->
Figure SMS_19
For the real-time speed dip +.>
Figure SMS_20
For the real-time ballistic deflection, +.>
Figure SMS_21
For the real-time ballistic dip angle.
In some embodiments, the determining a third attitude transfer matrix for transforming the velocity coordinate system to the rocket body coordinate system based on a target angle of attack and a real-time sideslip angle of the launch vehicle at a current time of flight comprises:
Figure SMS_22
wherein ,
Figure SMS_23
for the speed coordinate system->
Figure SMS_24
To the arrow coordinate system->
Figure SMS_25
Third pose transition matrix of>
Figure SMS_26
For the target angle of attack, < >>
Figure SMS_27
And the real-time sideslip angle is the real-time sideslip angle.
In some embodiments, the determining a fifth attitude transfer matrix for transforming the launch inertial coordinate system to the rocket body coordinate system based on the real-time pitch angle, the real-time yaw angle, and the real-time roll angle of the launch inertial coordinate system at the current time of flight of the launch vehicle comprises:
Figure SMS_28
wherein ,
Figure SMS_29
for the emission inertial coordinate system +.>
Figure SMS_30
To the arrow coordinate system->
Figure SMS_31
Fifth pose transition matrix of>
Figure SMS_32
For the real-time pitch angle +.>
Figure SMS_33
For said real time yaw angle +.>
Figure SMS_34
And (5) the real-time roll angle is the real-time roll angle.
In some embodiments, the real-time yaw angle
Figure SMS_35
Solving based on the following formula:
Figure SMS_36
the real-time pitch angle
Figure SMS_37
Solving based on the following formula:
Figure SMS_38
in the formula ,
Figure SMS_40
is a three-dimensional matrix->
Figure SMS_43
Element of (a)>
Figure SMS_46
Is a three-dimensional matrix->
Figure SMS_41
Element of (a)>
Figure SMS_44
For line number, ->
Figure SMS_47
Is a column number;
Figure SMS_48
is a first intermediate variable; />
Figure SMS_39
Is a second intermediate variable; />
Figure SMS_42
Is a third intermediate variable; />
Figure SMS_45
Is a fourth intermediate variable;
Figure SMS_49
；
Figure SMS_50
；
Figure SMS_51
。
the application provides a carrier rocket, which comprises a rocket body and an rocket-borne computer arranged on the rocket body;
the rocket-borne computer is used for executing the carrier rocket control method.
The application provides an electronic device, which comprises a memory, a processor and a computer program stored in the memory and capable of running on the processor, wherein the processor realizes the carrier rocket control method when executing the program.
The present application provides a non-transitory computer readable storage medium having stored thereon a computer program which when executed by a processor implements the launch vehicle control method.
The application provides a carrier rocket, which comprises a rocket body and an rocket-borne computer arranged on the rocket body; the rocket-borne computer is used for executing the carrier rocket control method.
The application provides an electronic device, which comprises a memory, a processor and a computer program stored in the memory and capable of running on the processor, wherein the processor realizes the carrier rocket control method when executing the program.
The present application provides a non-transitory computer readable storage medium having stored thereon a computer program which when executed by a processor implements the launch vehicle control method.
According to the carrier rocket control method, the carrier rocket, the electronic equipment and the storage medium, the posture association relation is determined according to the posture transfer matrixes among the launch inertial coordinate system, the launch coordinate system, the speed coordinate system and the rocket body coordinate system of the carrier rocket, the real-time trajectory dip angle and the real-time trajectory deflection angle are determined according to the real-time velocity vector of the carrier rocket at the current flight time, the real-time pitch angle and the real-time yaw angle of the carrier rocket at the current flight time and the posture association relation of the carrier rocket at the current flight time are determined, the flight posture of the carrier rocket is controlled, and as the posture association relation is established according to the posture transfer matrixes among the coordinate systems, compared with the method for solving the problem in the prior art, the method is easy to solve quickly through a computer, the complex calculation process of expanding each posture transfer matrix one by one and multiplying the pitch angle and the yaw angle is avoided, the solving time of the pitch angle and the yaw angle is shortened, the control response speed of the carrier rocket at the current flight time is shortened, and the control response time of the carrier rocket is shortened, and the control performance of the carrier rocket is improved.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the application and together with the description, serve to explain the principles of the application.
In order to more clearly illustrate the technical solutions of the present application or the prior art, the following description will briefly introduce the drawings used in the embodiments or the description of the prior art, and it is obvious that, in the following description, the drawings are some embodiments of the present application, and other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a schematic flow chart of a method for controlling a launch vehicle according to an embodiment of the present application;
FIG. 2 is a schematic representation of a transformation of a carrier rocket coordinate system provided in one embodiment of the present application;
FIG. 3 is a schematic view of a launch vehicle according to one embodiment of the present application;
fig. 4 is a schematic structural diagram of an electronic device according to an embodiment of the present application.
Detailed Description
In order to make the present application solution better understood by those skilled in the art, the following description will be made in detail and with reference to the accompanying drawings in the embodiments of the present application, it is apparent that the described embodiments are only some embodiments of the present application, not all embodiments. All other embodiments, which can be made by one of ordinary skill in the art based on the embodiments herein without making any inventive effort, shall fall within the scope of the present application.
It should be noted that the terms "first," "second," and the like herein are used for distinguishing between similar objects and not necessarily for describing a particular sequential or chronological order. It is to be understood that the data so used may be interchanged where appropriate such that embodiments of the present application described herein may be implemented in sequences other than those illustrated or otherwise described herein. Furthermore, the terms "comprises," "comprising," and "having," and any variations thereof, are intended to cover a non-exclusive inclusion, such that a process, method, system, article, or apparatus that comprises a list of steps or modules is not necessarily limited to those steps or modules that are expressly listed or inherent to such process, method, article, or apparatus.
Fig. 1 is a flow chart of a method for controlling a launch vehicle according to an embodiment of the present application, as shown in fig. 1, the method includes steps 110, 120, 130 and 140.
Step 110, obtaining a target attack angle, a real-time sideslip angle and a real-time speed vector in a launching coordinate system of the carrier rocket at the current flight time.
Specifically, the execution main body of the launch vehicle control method provided in the embodiment of the present application may be an rocket-borne computer of the launch vehicle.
The flight time refers to each time when the carrier rocket flies along the trajectory. The attack angle is the included angle between the speed direction of the carrier rocket and the longitudinal symmetry axis of the carrier rocket. The target attack angle is the attack angle which the carrier rocket needs to keep when turning at the current flight moment. The real-time velocity vector comprises the velocity magnitude and the velocity direction of the carrier rocket at the current flight time. The real-time sideslip angle refers to the angle between the real-time velocity vector of the launch vehicle and the longitudinal plane of symmetry of the launch vehicle.
In the launching process of the carrier rocket, the related coordinate system mainly comprises a launching inertial coordinate system, a launching coordinate system, a speed coordinate system and an rocket body coordinate system.
The emission coordinate system is a right-hand rectangular coordinate system which is formed by taking an emission point as an origin O, pointing an OX axis to an emission aiming direction in the horizontal plane of the emission point, enabling an OY axis to face upwards perpendicular to the horizontal plane of the emission point, and enabling an OZ axis to be perpendicular to the OX axis and the OY axis respectively. For example, the origin O in the launch coordinate system may be a projected point of the launch vehicle on the geodetic reference ellipsoidal surface. The OX axis may be directed in the emission direction in the tangential plane of the reference ellipsoid passing through the origin, and the OY may coincide with the local normal direction of the reference ellipsoid and be directed upward. Once the launch point of the launch vehicle is determined, the entire launch coordinate system is determined and remains unchanged throughout the launch of the launch vehicle.
The launching inertial coordinate system is overlapped with the launching coordinate system at the moment of launching the carrier rocket, and the directions of all coordinate axes in the launching coordinate system are kept unchanged in the inertial space after the carrier rocket is launched.
The velocity coordinate system uses the centroid of the launch vehicle as the origin O, OX axis (also denoted as
Figure SMS_52
Axis) is along the direction of the flight speed of the launch vehicle, the OY axis (also denoted +.>
Figure SMS_53
The axis) is positive in the main symmetry plane (longitudinal symmetry plane) of the launch vehicle and pointing perpendicular to the OX axis, the OZ axis (also denoted +.>
Figure SMS_54
Axis) perpendicular to the plane formed by the OX axis and the OY axis, pointing to the right as viewed in the direction of flight of the launch vehicle.
The rocket body coordinate system takes the mass center of the carrier rocket as an origin O, the OX axis points to the head of the carrier rocket along the symmetrical axis of the rocket body shell of the carrier rocket, the OY axis is positive in the main symmetrical plane of the carrier rocket and is perpendicular to the OX axis, and the OZ axis is perpendicular to the plane formed by the OX axis and the OY axis and points to the right when being seen along the launching direction of the carrier rocket.
Step 120, determining a real-time trajectory inclination angle and a real-time trajectory deflection angle of the carrier rocket at the current flight time based on the real-time velocity vector.
Specifically, the real-time velocity vector of the carrier rocket is analyzed, so that the real-time ballistic inclination angle and the real-time ballistic deflection angle of the carrier rocket at the current flight time can be obtained.
The real-time ballistic inclination angle refers to the angle between the real-time velocity vector of the launch vehicle and the horizontal plane. The real-time ballistic deflection refers to the angle between the projection of the real-time velocity vector of the launch vehicle on the horizontal plane and the OX axis of the launch coordinate system.
And 130, determining a real-time pitch angle and a real-time yaw angle of the carrier rocket at the current flight time based on the real-time trajectory inclination angle, the real-time trajectory deflection angle, the target attack angle and the real-time sideslip angle and the posture association relation of the carrier rocket at the current flight time.
The attitude association relation is determined based on an attitude transfer matrix among an launching inertial coordinate system, an launching coordinate system, a speed coordinate system and an rocket body coordinate system of the carrier rocket.
Specifically, there are transformation relations among the emission inertial coordinate system, the emission coordinate system, the velocity coordinate system and the arrow body coordinate system, and these transformation relations can be represented by the attitude transfer matrix. The posture transfer matrix is related to the posture angle in the corresponding coordinate system.
And performing matrix operation according to the gesture conversion matrix among the coordinate systems to obtain a gesture association relation. The posture association relationship is used to represent the association relationship between the posture angles in the respective coordinate systems.
These attitude angles include at least:
(1) Transmitting pitch angle, yaw angle, roll angle and the like in an inertial coordinate system;
(2) Velocity dip, trajectory dip, etc. in the launch coordinate system;
(3) Angle of attack and sideslip angle in a velocity coordinate system, and the like.
And substituting the real-time trajectory inclination angle and the real-time trajectory deflection angle into the posture association relation, and solving to obtain the real-time pitch angle and the real-time yaw angle of the carrier rocket at the current flight time. The real-time pitch angle is the pitch angle of the carrier rocket when turning at the current flight moment. The real-time yaw angle is the yaw angle of the carrier rocket when turning at the current flight time. The real-time roll angle is the roll angle when the carrier rocket turns at the current flight time.
And 140, controlling the flight attitude of the carrier rocket based on the real-time pitch angle and the real-time yaw angle.
Specifically, control parameters of the attitude and orbit control engine of the carrier rocket can be obtained according to the real-time pitch angle and the real-time yaw angle, so that the attitude and orbit control engine of the carrier rocket is controlled to change the output, the flight attitude of the carrier rocket is controlled, the pitch angle and the yaw angle of the carrier rocket are changed, and the turning action is completed.
According to the carrier rocket control method, the posture association relation is determined according to the launching inertial coordinate system, the launching coordinate system, the speed coordinate system and the posture transfer matrix between the rocket body coordinate systems of the carrier rocket, the real-time trajectory dip angle and the real-time trajectory dip angle are determined according to the real-time speed vector of the carrier rocket at the current flight time, the real-time pitch angle and the real-time yaw angle of the carrier rocket at the current flight time and the posture association relation of the carrier rocket at the current flight time are determined, the flight posture of the carrier rocket is controlled, and as the posture association relation is established according to the posture transfer matrix between the coordinate systems, compared with the trigonometric function in the related art, the inverse solution is easy to be quickly solved through a computer, the calculation speed of the pitch angle and the yaw angle of the carrier rocket is improved, the complex calculation process of expanding and multiplying the posture transfer matrix one by one is avoided, the solution efficiency of the pitch angle and the yaw angle is improved, the solution time of the pitch angle and the yaw angle is shortened, the control response speed of the carrier rocket is improved, and the control response time of the carrier rocket is shortened, and the control response performance of the carrier rocket is improved.
It should be noted that each embodiment of the present application may be freely combined, permuted, or executed separately, and does not need to rely on or rely on a fixed execution sequence.
In some embodiments, the gesture association is determined based on the steps of:
determining a first attitude transfer matrix for transforming a launch inertial coordinate system into a launch coordinate system based on an azimuth angle of the carrier rocket, a geographic latitude of a launch point, a current flight time and an earth rotation angular speed;
determining a second attitude transfer matrix for transforming the launching coordinate system into a speed coordinate system based on the real-time speed inclination angle, the real-time trajectory deflection angle and the real-time trajectory inclination angle of the carrier rocket at the current flight time;
determining a third attitude transfer matrix for transforming the velocity coordinate system to an rocket body coordinate system based on a target attack angle and a real-time sideslip angle of the carrier rocket at the current flight time;
determining a fourth attitude transfer matrix for transforming the launching inertial coordinate system to the arrow body coordinate system based on the first, second and third attitude transfer matrices;
determining a fifth gesture transfer matrix for transforming the launching inertial coordinate system to an rocket body coordinate system based on a real-time pitch angle, a real-time yaw angle and a real-time roll angle of the carrier rocket in the launching inertial coordinate system at the current flight time;
and determining the posture association relation based on the fourth posture transfer matrix and the fifth posture transfer matrix.
Specifically, fig. 2 is a schematic diagram of transformation of a carrier rocket coordinate system provided in an embodiment of the present application, and as shown in fig. 2, there are two transformation paths from an launching inertial coordinate system to an rocket body coordinate system. The first is to transform from the emission inertial coordinate system to the emission coordinate system, from the emission coordinate system to the velocity coordinate system, and from the velocity coordinate system to the arrow body coordinate system; the second is a direct transformation from the launching inertial coordinate system to the arrow coordinate system. The results obtained for both transformation paths are equal. Therefore, the posture association relation can be obtained according to matrix operation of the posture conversion matrix under the two transformation paths.
Starting from the first transformation path, the azimuth angle of the carrier rocket refers to the horizontal included angle between the north-pointing direction line of the launching point and the target direction line from the clockwise direction. According to the current flight time and the earth rotation angular velocity, the geographic longitude change between the current position of the carrier rocket and the launching point can be obtained, and then according to the azimuth angle and the geographic latitude of the launching point, the first attitude transfer matrix from the launching inertial coordinate system to the launching coordinate system can be obtained through solving. And according to the real-time velocity dip angle, the real-time trajectory deflection angle and the real-time trajectory dip angle of the carrier rocket at the current flight time, a second gesture transfer matrix from the launching coordinate system to the velocity coordinate system can be obtained through solving. And according to the target attack angle and the real-time sideslip angle of the carrier rocket at the current flight time, a third attitude transfer matrix from the velocity coordinate system to the rocket body coordinate system can be obtained by solving. And multiplying the three gesture transfer matrixes to obtain a fourth gesture transfer matrix from the emission inertial coordinate system to the arrow body coordinate system.
Starting from the second transformation path, according to the real-time pitch angle, the real-time yaw angle and the real-time roll angle of the carrier rocket in the launching inertial coordinate system at the current flight time, a fifth gesture transfer matrix for transforming the launching inertial coordinate system into the rocket body coordinate system can be obtained.
Because the fourth gesture transfer matrix and the fifth gesture transfer matrix are both transformed from the emission inertial coordinate system to the arrow body coordinate system and are equal, the gesture association relationship can be obtained.
According to the carrier rocket control method, the posture association relation is obtained according to the posture transfer matrix among the coordinate systems involved in the carrier rocket launching process, so that the posture angles under the coordinate systems can be quickly calculated, the control response speed of the carrier rocket is improved, the control response time of the carrier rocket is shortened, and the control performance of the carrier rocket is improved.
In some embodiments, determining a first attitude transfer matrix for transforming a launch inertial coordinate system to a launch coordinate system based on an azimuth of the launch vehicle, a launch point geographic latitude, a current time of flight, and an earth rotation angular velocity, comprises:
Figure SMS_55
wherein ,
Figure SMS_58
for transmitting inertial coordinate system->
Figure SMS_63
To the emission coordinate system->
Figure SMS_66
Is a first posture transfer matrix of->
Figure SMS_57
For azimuth angle->
Figure SMS_61
For the geographical latitude of the transmitting point>
Figure SMS_65
For the current flight time>
Figure SMS_67
For the rotation angular velocity of the earth>
Figure SMS_56
For representing a surrounding coordinate system +.>
Figure SMS_60
A rotation matrix for rotation of the shaft; />
Figure SMS_64
For representing a surrounding coordinate system +.>
Figure SMS_68
A rotation matrix for rotation of the shaft; />
Figure SMS_59
For representing a wound coordinate system
Figure SMS_62
A rotation matrix for rotating the shaft.
Specifically, in each coordinate systemIn the course of the transformation the phase of the transformation,
Figure SMS_69
for representing a surrounding coordinate system +.>
Figure SMS_70
The rotation angle of the shaft is +.>
Figure SMS_71
Can be expressed as:
Figure SMS_72
Figure SMS_73
for representing a surrounding coordinate system +.>
Figure SMS_74
The rotation angle of the shaft is +.>
Figure SMS_75
Can be expressed as:
Figure SMS_76
Figure SMS_77
for representing a surrounding coordinate system +.>
Figure SMS_78
The rotation angle of the shaft is +.>
Figure SMS_79
Can be expressed as:
Figure SMS_80
relative to the emission inertial frame
Figure SMS_81
Is a transmission coordinate system +.>
Figure SMS_82
There is at least one rotation of each of the corresponding three shafts.
Emission coordinate system
Figure SMS_83
Is->
Figure SMS_84
Axes are +.>
Figure SMS_85
Is->
Figure SMS_86
The angle by which the axis passes is the angle of rotation of the earth (change in geographical longitude), and the rotation angular velocity of the earth can be used>
Figure SMS_87
Is +.>
Figure SMS_88
Is obtained by the product of (2).
Emission coordinate system
Figure SMS_89
Is->
Figure SMS_90
Axes are +.>
Figure SMS_91
Is->
Figure SMS_92
The angle at which the axis passes around is the azimuth angle. During the transformation of the coordinate system, two rotations occur.
Emission coordinate system
Figure SMS_93
Is->
Figure SMS_94
Axes are +.>
Figure SMS_95
Is->
Figure SMS_96
The angle of axis bypass is the launch point geographical latitude. During the transformation of the coordinate system, two rotations occur.
The sequence of the above rotations is:
1. emission coordinate system
Figure SMS_97
Is->
Figure SMS_98
Axes are +.>
Figure SMS_99
Is->
Figure SMS_100
The angle of the shaft bypass is +.>
Figure SMS_101
；
2. Emission coordinate system
Figure SMS_102
Is->
Figure SMS_103
Axes are +.>
Figure SMS_104
Is->
Figure SMS_105
The angle of the shaft bypass is +.>
Figure SMS_106
；
3. Emission coordinate system
Figure SMS_107
Is->
Figure SMS_108
Axes are +.>
Figure SMS_109
Is->
Figure SMS_110
The angle of the shaft bypass is +.>
Figure SMS_111
；
4. Emission coordinate system
Figure SMS_112
Is->
Figure SMS_113
Axes are +.>
Figure SMS_114
Is->
Figure SMS_115
The angle of the shaft bypass is +.>
Figure SMS_116
；
5. Emission coordinate system
Figure SMS_117
Is->
Figure SMS_118
Axes are +.>
Figure SMS_119
Is->
Figure SMS_120
The angle of the shaft bypass is +.>
Figure SMS_121
。
According to the above angle rotation relationship, the rotation matrix corresponding to each axis and the rotation conversion sequence of each axis, a first posture transfer matrix can be obtained
Figure SMS_122
。
According to the carrier rocket control method, the first attitude rotation matrix is obtained according to the transmission inertial coordinate system and rotation transformation of three coordinate axes in the transmission coordinate system, and the conversion relation between the transmission inertial coordinate system and the transmission coordinate system can be accurately represented.
In some embodiments, determining a second attitude transfer matrix for transforming the launch coordinate system to the velocity coordinate system based on the real-time velocity tilt, the real-time ballistic deflection, and the real-time ballistic tilt of the launch vehicle at the current time of flight comprises:
Figure SMS_123
wherein ,
Figure SMS_124
for transmitting the coordinate system->
Figure SMS_125
To the speed coordinate system->
Figure SMS_126
Is a second posture transfer matrix of->
Figure SMS_127
For real-time speed dip->
Figure SMS_128
For real-time ballistic deflection +.>
Figure SMS_129
Is the real-time ballistic dip angle.
Specifically, first, a velocity coordinate system
Figure SMS_131
Is->
Figure SMS_136
Axis is +.>
Figure SMS_139
Is->
Figure SMS_132
The angle of the shaft bypass is real-time ballistic inclination angle +.>
Figure SMS_134
The method comprises the steps of carrying out a first treatment on the surface of the Second, the velocity coordinate System +.>
Figure SMS_138
Is->
Figure SMS_142
Axis is +.>
Figure SMS_130
Is->
Figure SMS_137
The angle of the shaft bypass is real-time ballistic deflection angle +.>
Figure SMS_141
The method comprises the steps of carrying out a first treatment on the surface of the Finally, the velocity coordinate System>
Figure SMS_144
Is->
Figure SMS_133
Axis is +.>
Figure SMS_135
Is->
Figure SMS_140
The angle of the shaft bypass is real-time speed dip angle +.>
Figure SMS_143
。
According to the above angle rotation relation, the rotation matrix corresponding to each axis and the rotation conversion sequence of each axis, a second posture transfer matrix can be obtained
Figure SMS_145
。
According to the carrier rocket control method, the second attitude rotation matrix is obtained according to rotation transformation of three coordinate axes in the launching coordinate system and the speed coordinate system, and the conversion relation between the launching coordinate system and the speed coordinate system can be accurately represented.
In some embodiments, determining a third attitude transfer matrix for transforming the velocity coordinate system to the rocket body coordinate system based on the target angle of attack and the real-time sideslip angle of the launch vehicle at the current time of flight comprises:
Figure SMS_146
wherein ,
Figure SMS_147
for the speed coordinate system->
Figure SMS_148
To the arrow coordinate system->
Figure SMS_149
Third pose transition matrix of>
Figure SMS_150
For the target angle of attack->
Figure SMS_151
Is the real-time sideslip angle.
Specifically, first, a velocity coordinate system
Figure SMS_153
Is->
Figure SMS_155
The axes are +.>
Figure SMS_159
Is->
Figure SMS_154
The angle of the shaft bypass is real-time sideslip angle +.>
Figure SMS_156
The method comprises the steps of carrying out a first treatment on the surface of the Second, the velocity coordinate System +.>
Figure SMS_158
Is->
Figure SMS_161
The axes are +.>
Figure SMS_152
Is->
Figure SMS_157
The angle of axis bypass is the target angle of attack +.>
Figure SMS_160
。
According to the above angle rotation relation, the rotation matrix corresponding to each axis and the rotation transformation sequence of each axis, a third posture transfer matrix can be obtained
Figure SMS_162
。
According to the carrier rocket control method, the third attitude rotation matrix is obtained according to rotation transformation of three coordinate axes in the speed coordinate system and the rocket body coordinate system, and the conversion relation between the speed coordinate system and the rocket body coordinate system can be accurately represented.
In some embodiments, a fourth pose transfer matrix is determined based on the first, second, and third pose transfer matrices, with the launching inertial coordinate system transformed to the arrow body coordinate system.
In particular, the emission inertial coordinate system
Figure SMS_163
Transformation to arrow coordinate System->
Figure SMS_164
Fourth pose transition matrix of->
Figure SMS_165
Can be expressed as:
Figure SMS_166
in some embodiments, determining a fifth attitude transfer matrix for transforming the launch inertial coordinate system to the rocket body coordinate system based on the real-time pitch angle, the real-time yaw angle, and the real-time roll angle of the launch inertial coordinate system of the launch vehicle at the current time of flight comprises:
Figure SMS_167
wherein ,
Figure SMS_168
for transmitting inertial coordinate system->
Figure SMS_169
To the arrow coordinate system->
Figure SMS_170
Fifth pose transition matrix of>
Figure SMS_171
For real time pitch angle>
Figure SMS_172
For real-time yaw angle>
Figure SMS_173
Is the real-time roll angle.
Specifically, first, an arrow body coordinate system
Figure SMS_175
Is->
Figure SMS_181
Axes are +.>
Figure SMS_185
Is->
Figure SMS_177
The angle of shaft bypass is real-time pitch angle +.>
Figure SMS_180
The method comprises the steps of carrying out a first treatment on the surface of the Secondly, arrow coordinate system->
Figure SMS_184
Is->
Figure SMS_188
Axes are +.>
Figure SMS_174
Is->
Figure SMS_178
The angle of the shaft bypass is real-time yaw angle +.>
Figure SMS_182
The method comprises the steps of carrying out a first treatment on the surface of the Finally, arrow coordinate System->
Figure SMS_186
Is->
Figure SMS_176
Axes are +.>
Figure SMS_179
Is->
Figure SMS_183
The angle of the shaft bypass is real-time roll angle +.>
Figure SMS_187
。
According to the above angle rotation relation, the rotation matrix corresponding to each axis and the rotation conversion sequence of each axis, a fifth posture transfer matrix can be obtained
Figure SMS_189
。
According to the carrier rocket control method, the fifth gesture rotation matrix is obtained according to rotation transformation of three coordinate axes in the rocket body coordinate system and the launching inertial coordinate system, and the conversion relation between the rocket body coordinate system and the launching inertial coordinate system can be accurately represented.
In some embodiments, the pose correlation is determined based on the fourth pose transfer matrix and the fifth pose transfer matrix.
Specifically, since the fourth posture transfer matrix and the fifth posture transfer matrix are both indicative of the conversion relationship from the emission inertial coordinate system to the arrow body coordinate system, which are equal, the first relational expression can be obtained:
Figure SMS_190
the first relationship may be used to represent a gesture association.
Substitution into
Figure SMS_191
and />
Figure SMS_192
The second relation can be further obtained:
Figure SMS_193
in the second relation:
Figure SMS_195
middle azimuth +.>
Figure SMS_197
Geographical latitude of transmitting point->
Figure SMS_202
Rotational speed of the earth->
Figure SMS_196
Is known, the current flight moment +.>
Figure SMS_199
Are also known; real-time ballistic deflection->
Figure SMS_201
And real-time ballistic dip>
Figure SMS_203
The real-time velocity vector can be calculated according to a transmitting coordinate system; target attack angle->
Figure SMS_194
Is designed in advance and is input through a program; real-time sideslip angle->
Figure SMS_198
The method can be calculated according to the included angle between the real-time velocity vector and the longitudinal symmetry plane of the carrier rocket; for launch vehicle ballistic design, real-time roll angle under launch inertial frame +.>
Figure SMS_200
May be set to zero.
Therefore, of the 8 attitude angles involved in the second relation, the real-time roll angle
Figure SMS_205
Target attack angle->
Figure SMS_208
Real-time sideslip angle->
Figure SMS_210
Real-time ballistic deflection->
Figure SMS_206
And real-time ballistic dip>
Figure SMS_207
Are known; only real-time speed dip +.>
Figure SMS_209
Real-time pitch angle->
Figure SMS_211
Real-time yaw angle->
Figure SMS_204
Is unknown.
Taking into account that
Figure SMS_212
Then->
Figure SMS_213
，/>
Figure SMS_214
For a 3-dimensional identity matrix, the second relation can be reduced to a third relation:
Figure SMS_215
the transformation of the rotation matrix is known as:
Figure SMS_216
thus, the third relationship may be expressed as a fourth relationship:
Figure SMS_217
respectively using matrix
Figure SMS_218
Sum matrix->
Figure SMS_219
Representing a rotation transformation:
Figure SMS_220
wherein ,
Figure SMS_221
for matrix->
Figure SMS_222
Element of (a)>
Figure SMS_223
For matrix->
Figure SMS_224
Element of (a)>
Figure SMS_225
For line number, ->
Figure SMS_226
Are column numbers.
The fourth relationship may be expressed as:
Figure SMS_227
expanding the fourth relation can obtain a fifth relation:
Figure SMS_228
further, a size of
Figure SMS_229
Matrix->
Figure SMS_230
Representation->
Figure SMS_231
The following steps are:
Figure SMS_232
the method can obtain the following steps:
Figure SMS_233
wherein ,
Figure SMS_234
for matrix->
Figure SMS_235
Element of (a)>
Figure SMS_236
For line number, ->
Figure SMS_237
Are column numbers.
Since the elements corresponding to the matrix are equal, there is
Figure SMS_238
I.e. +.>
Figure SMS_239
. Consider->
Figure SMS_240
Substitution can yield the equation:
Figure SMS_241
recording device
Figure SMS_242
，/>
Figure SMS_243
Is a first intermediate variable; />
Figure SMS_244
Is a second intermediate variable; />
Figure SMS_245
As a third intermediate variable, the equation may represent:
Figure SMS_246
solving the equation can obtain:
Figure SMS_247
taking into account that
Figure SMS_250
Are small values that vary around 0, so +.>
Figure SMS_253
Thus get +.>
Figure SMS_255
By->
Figure SMS_249
Calculate->
Figure SMS_252
According to the value of->
Figure SMS_254
and />
Figure SMS_256
Can obtain real-time yaw angle
Figure SMS_248
And judge->
Figure SMS_251
Is the sign of (c).
Similarly, there is
Figure SMS_257
，/>
Figure SMS_258
。
Can be expressed in a matrix as:
Figure SMS_259
recording device
Figure SMS_260
，/>
Figure SMS_261
Substituting the matrix for the fourth intermediate variable can result in: />
Figure SMS_262
Solving the matrix to obtain
Figure SMS_263
and />
Figure SMS_264
Furthermore, the real-time pitch angle can be obtained>
Figure SMS_265
And judge->
Figure SMS_266
Is the sign of (c).
According to the carrier rocket control method, the matrix capable of representing the posture association relation is obtained through operation of the posture conversion matrix under the two conversion paths, the matrix can be obtained through matrix multiplication realized through computer code programming without manual expansion and multiplication, the solving efficiency of the pitch angle and the yaw angle can be improved, the solving time of the pitch angle and the yaw angle is shortened, and the control response speed of the carrier rocket is improved.
FIG. 3 is a schematic view of a carrier rocket according to one embodiment of the present application, and as shown in FIG. 3, the carrier rocket 300 includes a rocket body 310, and an rocket-borne computer 320 disposed on the rocket body 310; the rocket-borne computer 320 is used to perform the launch vehicle control method in the above-described embodiments.
The control method in the embodiment of the carrier rocket is executed in the launching process, so that the calculation speed of the pitch angle and the yaw angle of the carrier rocket is improved, the complex operation process of expanding and multiplying each attitude transfer matrix one by one is avoided, the solving efficiency of the pitch angle and the yaw angle is improved, the solving time of the pitch angle and the yaw angle is shortened, the control response speed of the carrier rocket is improved, the control response time of the carrier rocket is shortened, and the control performance of the carrier rocket is improved.
Fig. 4 is a schematic structural diagram of an electronic device according to an embodiment of the present application, as shown in fig. 4, the electronic device may include: processor (Processor) 410, communication interface (Communications Interface) 420, memory (Memory) 430, and communication bus (Communications Bus) 440, wherein Processor 410, communication interface 420, memory 430 complete communication with each other via communication bus 440. The processor 410 may invoke logic commands in the memory 430 to perform the method described above, including:
acquiring a target attack angle, a real-time sideslip angle and a real-time velocity vector in a launching coordinate system of a carrier rocket at the current flight time; based on the real-time velocity vector, determining a real-time trajectory inclination angle and a real-time trajectory deflection angle of the carrier rocket at the current flight time; determining a real-time pitch angle and a real-time yaw angle of the carrier rocket at the current flight time based on the real-time trajectory inclination angle, the real-time trajectory deflection angle, the target attack angle, the real-time sideslip angle and the posture association relation of the carrier rocket at the current flight time; controlling the flight attitude of the carrier rocket based on the real-time pitch angle and the real-time yaw angle; the attitude association relation is determined based on an attitude transfer matrix among an launching inertial coordinate system, an launching coordinate system, a speed coordinate system and an rocket body coordinate system of the carrier rocket.
In addition, the logic commands in the memory described above may be implemented in the form of software functional units and may be stored in a computer readable storage medium when sold or used as a stand alone product. Based on such understanding, the technical solution of the present application may be embodied essentially or in a part contributing to the prior art or in a part of the technical solution, in the form of a software product stored in a storage medium, comprising several commands for causing a computer device (which may be a personal computer, a server, or a network device, etc.) to perform all or part of the steps of the methods described in the embodiments of the present application. And the aforementioned storage medium includes: a U-disk, a removable hard disk, a Read-Only Memory (ROM), a random access Memory (RAM, random Access Memory), a magnetic disk, or an optical disk, or other various media capable of storing program codes.
The processor in the electronic device provided by the embodiment of the present application may call the logic instruction in the memory to implement the above method, and the specific implementation manner of the processor is consistent with the implementation manner of the foregoing method, and may achieve the same beneficial effects, which are not described herein again.
The present application also provides a computer-readable storage medium having stored thereon a computer program which, when executed by a processor, is implemented to perform the methods provided by the above embodiments.
The specific embodiment is consistent with the foregoing method embodiment, and the same beneficial effects can be achieved, and will not be described herein.
Embodiments of the present application provide a computer program product comprising a computer program which, when executed by a processor, implements a method as described above.
The apparatus embodiments described above are merely illustrative, wherein the elements illustrated as separate elements may or may not be physically separate, and the elements shown as elements may or may not be physical elements, may be located in one place, or may be distributed over a plurality of network elements. Some or all of the modules may be selected according to actual needs to achieve the purpose of the solution of this embodiment. Those of ordinary skill in the art will understand and implement the present invention without undue burden.
From the above description of the embodiments, it will be apparent to those skilled in the art that the embodiments may be implemented by means of software plus necessary general hardware platforms, or of course may be implemented by means of hardware. Based on this understanding, the foregoing technical solution may be embodied essentially or in a part contributing to the prior art in the form of a software product, which may be stored in a computer readable storage medium, such as ROM/RAM, a magnetic disk, an optical disk, etc., including several instructions for causing a computer device (which may be a personal computer, a server, or a network device, etc.) to execute the method described in the respective embodiments or some parts of the embodiments.
Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present application, and are not limiting thereof; although the present application has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit and scope of the corresponding technical solutions.

Claims (10)

1. A method of controlling a launch vehicle, comprising:
acquiring a target attack angle, a real-time sideslip angle and a real-time velocity vector in a launching coordinate system of a carrier rocket at the current flight time;
determining a real-time ballistic inclination angle and a real-time ballistic deflection angle of the carrier rocket at the current flight moment based on the real-time velocity vector;
determining a real-time pitch angle and a real-time yaw angle of the carrier rocket at the current flight time based on the real-time trajectory inclination angle, the real-time trajectory deflection angle, the target attack angle, the real-time sideslip angle and the posture association relation of the carrier rocket at the current flight time;
controlling the flight attitude of the carrier rocket based on the real-time pitch angle and the real-time yaw angle;
the attitude association relation is determined based on an attitude transfer matrix among an launching inertial coordinate system, an launching coordinate system, a speed coordinate system and an rocket body coordinate system of the carrier rocket.
2. A method of controlling a launch vehicle according to claim 1, wherein the attitude association is determined based on the steps of:
determining a first attitude transfer matrix of the launching inertial coordinate system transformed to the launching coordinate system based on the azimuth angle of the carrier rocket, the geographic latitude of the launching point, the current flight time and the rotation angular speed of the earth;
determining a second attitude transfer matrix for transforming the launching coordinate system to the speed coordinate system based on the real-time speed inclination angle, the real-time trajectory deflection angle and the real-time trajectory inclination angle of the carrier rocket at the current flight time;
determining a third attitude transfer matrix of the speed coordinate system transformed to the rocket body coordinate system based on a target attack angle and a real-time sideslip angle of the carrier rocket at the current flight time;
determining a fourth pose transfer matrix for transforming the launching inertial coordinate system to the arrow body coordinate system based on the first pose transfer matrix, the second pose transfer matrix, and the third pose transfer matrix;
determining a fifth attitude transfer matrix for transforming the launch inertial coordinate system to the rocket body coordinate system based on a real-time pitch angle, a real-time yaw angle and a real-time roll angle of the carrier rocket in the launch inertial coordinate system at the current flight time;
and determining the posture association relation based on the fourth posture transfer matrix and the fifth posture transfer matrix.
3. The method of claim 2, wherein determining the first attitude transfer matrix for transforming the launch inertial coordinate system to the launch coordinate system based on the azimuth angle of the launch vehicle, the launch point geographical latitude, the current time of flight, and the earth rotation angular velocity comprises:
Figure QLYQS_1
；
wherein ,
Figure QLYQS_4
for the emission inertial coordinate system +.>
Figure QLYQS_8
To the emission coordinate system->
Figure QLYQS_12
Is a first posture transfer matrix of->
Figure QLYQS_3
For the azimuth angle>
Figure QLYQS_7
For the geographical latitude of the emission point +.>
Figure QLYQS_11
For the current flight time>
Figure QLYQS_14
For the earth rotation angular velocity, +.>
Figure QLYQS_2
For representing a surrounding coordinate system +.>
Figure QLYQS_6
A rotation matrix for rotation of the shaft; />
Figure QLYQS_10
For the purpose of representationAround coordinate system->
Figure QLYQS_13
A rotation matrix for rotation of the shaft; />
Figure QLYQS_5
For representing a surrounding coordinate system +.>
Figure QLYQS_9
A rotation matrix for rotating the shaft.
4. A method of controlling a launch vehicle according to claim 3, wherein said determining a second attitude transfer matrix for transforming the launch coordinate system to the velocity coordinate system based on the real-time velocity tilt, the real-time ballistic deflection and the real-time ballistic tilt of the launch vehicle at the current time of flight comprises:
Figure QLYQS_15
；
wherein ,
Figure QLYQS_16
for the emission coordinate system->
Figure QLYQS_17
To the velocity coordinate system->
Figure QLYQS_18
Is a second posture transfer matrix of->
Figure QLYQS_19
For the real-time speed dip +.>
Figure QLYQS_20
For the real-time ballistic deflection, +.>
Figure QLYQS_21
Is saidReal-time ballistic dip angle.
5. The method of claim 4, wherein determining a third attitude transfer matrix for transforming the velocity coordinate system to the rocket body coordinate system based on a target angle of attack and a real-time sideslip angle of the rocket at a current time of flight comprises:
Figure QLYQS_22
；
wherein ,
Figure QLYQS_23
for the speed coordinate system->
Figure QLYQS_24
To the arrow coordinate system->
Figure QLYQS_25
Third pose transition matrix of>
Figure QLYQS_26
For the target angle of attack, < >>
Figure QLYQS_27
And the real-time sideslip angle is the real-time sideslip angle.
6. The method of claim 5, wherein determining a fifth attitude transfer matrix for transforming the launch inertial coordinate system to the rocket body coordinate system based on the real-time pitch angle, the real-time yaw angle, and the real-time roll angle of the launch inertial coordinate system at the current time of flight of the launch rocket, comprises:
Figure QLYQS_28
；
wherein ,
Figure QLYQS_29
for the emission inertial coordinate system +.>
Figure QLYQS_30
To the arrow coordinate system->
Figure QLYQS_31
Fifth pose transition matrix of>
Figure QLYQS_32
For the real-time pitch angle +.>
Figure QLYQS_33
For said real time yaw angle +.>
Figure QLYQS_34
And (5) the real-time roll angle is the real-time roll angle.
7. A method of controlling a launch vehicle according to claim 6 wherein the real time yaw angle
Figure QLYQS_35
Solving based on the following formula:
Figure QLYQS_36
；
the real-time pitch angle
Figure QLYQS_37
Solving based on the following formula:
Figure QLYQS_38
；
in the formula ,
Figure QLYQS_41
is a three-dimensional matrix->
Figure QLYQS_43
Element of (a)>
Figure QLYQS_46
Is a three-dimensional matrix->
Figure QLYQS_40
Element of (a)>
Figure QLYQS_42
For line number, ->
Figure QLYQS_45
Is a column number; />
Figure QLYQS_48
Is a first intermediate variable; />
Figure QLYQS_39
Is a second intermediate variable; />
Figure QLYQS_44
Is a third intermediate variable; />
Figure QLYQS_47
Is a fourth intermediate variable;
Figure QLYQS_49
；
Figure QLYQS_50
；
Figure QLYQS_51
。
8. the carrier rocket is characterized by comprising a rocket body and an rocket-borne computer arranged on the rocket body;
the rocket-borne computer being configured to perform the launch vehicle control method of any one of claims 1 to 7.
9. An electronic device comprising a memory, a processor and a computer program stored on the memory and executable on the processor, wherein the processor implements the launch vehicle control method of any one of claims 1 to 7 when the computer program is executed.
10. A non-transitory computer readable storage medium having stored thereon a computer program, which when executed by a processor implements a launch vehicle control method according to any one of claims 1 to 7.
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