CN112461060B - Rocket final-stage derailment control method and device - Google Patents
Rocket final-stage derailment control method and device Download PDFInfo
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- CN112461060B CN112461060B CN202011287038.3A CN202011287038A CN112461060B CN 112461060 B CN112461060 B CN 112461060B CN 202011287038 A CN202011287038 A CN 202011287038A CN 112461060 B CN112461060 B CN 112461060B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
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- F42B15/01—Arrangements thereon for guidance or control
Abstract
The invention relates to a rocket final-stage derailment control method and a rocket final-stage derailment control device, wherein the method comprises the following steps: acquiring a first position vector, a first speed vector and available fuel mass in the rocket navigation coordinate system at the current moment, wherein the current moment is any moment of a powerless section before the rocket starts to execute the derailment action; determining a maximum ignition duration for the rocket based on the available fuel mass; and determining the actual flight attack angle of the rocket at the departure starting moment by utilizing iterative calculation based on the first position vector, the first speed vector, the maximum ignition time and the preset flight attack angle, so that the perigee height corresponding to the close orbit element after the last stage of the rocket leaves the orbit active section is equal to the preset target height. The invention can enhance the stability of rocket final stage derailment and avoid incomplete derailment and incomplete propellant consumption caused by deviation in rocket flight.
Description
Technical Field
The invention relates to the technical field of guidance control, in particular to a solid carrier rocket final stage derailment control method and device.
Background
The traditional rocket derailing scheme is that before the solid carrier rocket is launched, the initial condition of rocket final-stage derailing is generally obtained by calculating a standard trajectory, and the thrust vector direction of the derailing active section (namely the pitch angle and the yaw angle when the rocket is designed to derail) and the engine startup time of the derailing active section are designed according to simulation analysis. And binding pre-designed thrust vector parameters (pitch angle and yaw angle) and the engine starting time of the off-orbit active section to an rocket-borne computer, and executing off-orbit action by the rocket-borne computer according to the pre-bound value during an actual flight test. However, in an actual flight test task, because actual conditions are not completely consistent with theoretical calculation, parameters such as an actual orbit-in point of a rocket final stage and residual fuel quality of a rocket final stage engine are not consistent with design conditions of a theoretical orbit-off scheme before shooting, and if the orbit-off action is completely executed by using a binding value before shooting, the orbit-off effect cannot be guaranteed to be consistent with theoretical analysis.
For a multi-stage solid launch vehicle, if the first stage solid engine has energy deviation from the theoretical value, the last stage engine needs to consume unequal fuel for correcting the first stage deviation, so that the residual fuel quantity entering the off-rail section has deviation. If the remaining fuel is lower than the theoretical value, the last-stage engine may not actually maintain the startup duration of the pre-injection binding, and at this time, the off-rail posture of the pre-injection binding is continued, which may result in the failure of the off-rail, and if the remaining fuel is higher than the theoretical value, the incomplete combustion of the engine fuel may be caused.
Disclosure of Invention
The invention aims to provide a rocket final stage derailment control method and equipment, which can control the flight attack angle of the derailment according to the real-time state of a rocket, enhance the stability of the rocket final stage derailment, and avoid incomplete derailment and incomplete propellant consumption caused by deviation in rocket flight.
The embodiment of the invention provides the following scheme:
in a first aspect, an embodiment of the present invention provides a rocket final stage derailment control method, where the method includes:
acquiring a first position vector, a first speed vector and available fuel mass in the rocket navigation coordinate system at the current moment, wherein the current moment is any moment of a powerless section before the rocket starts to execute the derailment action;
determining a maximum ignition duration for the rocket based on the available fuel mass;
and determining the actual flight attack angle of the rocket at the departure starting moment by utilizing iterative computation based on the first position vector, the first velocity vector, the maximum ignition time and the preset flight attack angle, so that the perigee height corresponding to the close orbit element after the last stage of the rocket leaves the active section is equal to the preset target height.
Optionally, determining the actual flight angle of attack of the rocket off-orbit active segment by using iterative computation based on the first position vector, the first velocity vector, the maximum ignition duration and a preset flight angle of attack includes:
calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of an off-orbit active section based on the preset flight attack angle;
determining the osculating rail element at the end of off-track based on the pitch angle, the yaw angle, the first position vector, the first velocity vector and the maximum ignition duration;
determining the height of the perigee corresponding to the close orbit element;
determining that the difference between the height of the near place and the preset target height is greater than or equal to a preset value, and determining the adjustment quantity of the preset flight attack angle based on the difference;
adjusting the preset flight attack angle based on the adjustment amount;
calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active section based on the adjusted preset flight attack angle; determining the osculating orbit element at the off-track ending time based on the calculated pitch and yaw angles, the first position vector, the first velocity vector and the maximum ignition duration; determining the perigee height corresponding to the osculating orbit element; and determining that the difference value between the near-place altitude and the preset target altitude is greater than or equal to a preset value, determining the adjustment amount of the adjusted preset flight attack angle based on the difference value, repeating iterative calculation in this way until the difference value between the near-place altitude corresponding to the close orbit element and the preset target altitude is determined to be smaller than the preset value, ending the iterative calculation process, and determining the adjusted preset flight attack angle as the actual flight attack angle of the rocket at the departure starting moment.
Optionally, the determining the close track element at the off-track end time based on the pitch angle, the yaw angle, the first position vector, the first velocity vector, and the maximum ignition duration includes:
determining the rocket derailment starting time based on the current time and a standard flight time sequence;
extrapolating to obtain a second position vector and a second velocity vector of the rocket in the navigation coordinate system at the off-orbit starting moment based on the first position vector and the first velocity vector in the navigation coordinate system of the rocket at the current moment;
extrapolating the second position vector and the second velocity vector to obtain a third position vector and a third velocity vector of the rocket derailing active section ending time based on the pitch angle and the yaw angle, wherein the rocket derailing active section ending time t is 1 ＝t 0 +t f ，t 0 Is the rocket off-orbit starting time, t f Is the maximum ignition time period;
performing coordinate conversion on the third position vector and the third speed vector to obtain a position vector and a speed vector in a geocentric equatorial coordinate system of the rocket; and
and determining the close track element at the off-track ending time based on the position vector and the speed vector in the geocentric equatorial coordinate system.
Optionally, the method uses a kinetic equation to extrapolate a second position vector and a second velocity vector of the rocket in the navigation coordinate system at the departure starting time, and extrapolate a third position vector and a third velocity vector of the rocket at the departure active section ending time.
Optionally, the determining the height of the perigee corresponding to the close track element includes:
and calculating to obtain the height of the perigee by utilizing a orbit determination formula based on the close orbit elements.
Optionally, the calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active segment based on the preset flight attack angle includes:
determining a coordinate value of a rocket final thrust vector in a geocentric orbit coordinate system based on the trigonometric function values of the preset flight attack angle, the local trajectory inclination angle, the sideslip angle and the orbit true approach point angle;
converting the rocket final stage thrust vector from the geocentric orbit coordinate system to a geocentric equatorial coordinate system;
converting the rocket final stage thrust vector from the geocentric equatorial coordinate system to a launching coordinate system;
converting the rocket final thrust vector from the launching coordinate system to a navigation coordinate system; and
and calculating the pitch angle and the yaw angle of the rocket in a navigation coordinate system.
Optionally, the determining a maximum ignition duration of the rocket based on the available fuel mass comprises:
determining the maximum ignition time period based on the available fuel mass and a final propellant specific impulse of the rocket, a sea level gravitational acceleration constant, and a final engine thrust of the rocket.
In a second aspect, an embodiment of the present invention provides a rocket final-stage derailment control device, where the device includes: the system comprises an acquisition module, a control module and a control module, wherein the acquisition module is used for acquiring a first position vector, a first speed vector and available fuel quality in the rocket navigation coordinate system at the current moment, and the current moment is any moment of a powerless section before the rocket starts to execute the off-orbit action;
a first determination module to determine a maximum ignition duration of the rocket based on the available fuel mass;
and the second determining module is used for determining the actual flight attack angle of the rocket at the departure starting moment by utilizing iterative calculation based on the first position vector, the first speed vector, the maximum ignition duration and the preset flight attack angle, so that the near-point height corresponding to the close orbit element after the last stage of the rocket leaves the orbit active section is equal to the preset target height.
In a third aspect, an embodiment of the present invention provides a rocket final stage derailment control device, including:
a memory for storing a computer program;
a processor for executing the computer program to implement the steps of the method for measuring a light emission characteristic of an infrared thermal radiation light source according to any one of the first aspect.
In a fourth aspect, an embodiment of the present invention provides a computer-readable storage medium, on which a computer program is stored, where the computer program is executed by a processor to implement the steps of the rocket final stage derailment control method in any one of the first aspect.
Compared with the prior art, the invention has the following advantages and beneficial effects:
the invention controls the flight attack angle of the derailment according to the real-time state of the rocket, such as the first position vector, the first speed vector and the available fuel quality, enhances the stability of the final derailment of the rocket, and avoids incomplete derailment and incomplete propellant consumption caused by deviation in the rocket flight.
Drawings
In order to more clearly illustrate the embodiments of the present specification or the technical solutions in the prior art, the drawings needed to be used in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present specification, and it is obvious for those skilled in the art that other drawings can be obtained according to the drawings without creative efforts.
Fig. 1 is a flowchart of a rocket final-stage derailment control method according to an embodiment of the present invention;
FIG. 2 is a partial flow diagram of a rocket final-stage derailment control method shown in FIG. 1;
fig. 3 is a schematic structural diagram of a rocket final stage derailment control device according to an embodiment of the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be described clearly and completely with reference to the accompanying drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, rather than all embodiments, and all other embodiments obtained by those skilled in the art based on the embodiments of the present invention belong to the scope of protection of the embodiments of the present invention.
Referring to fig. 1, fig. 1 is a flowchart of a rocket final stage derailment control method according to an embodiment of the present invention. The method for controlling the final stage derailment of the rocket determines a derailment strategy according to the actual flight state of the final stage of the rocket before the final stage of the rocket derails.
The orbit determination formula of the semi-major axis of the orbit is as follows:
wherein r is a position vector, v is a velocity vector, a is a semimajor axis of the orbit, and μ is a constant of the earth gravity, the semimajor axis a of the orbit decreases with the decrease of the velocity of the spacecraft, and when the semimajor axis of the orbit decreases, the altitude of the corresponding near-site of the spacecraft decreases. Therefore, theoretically, if the rocket final stage is ignited completely along the opposite direction of the current speed direction of the rocket final stage in the off-orbit section, the height of the rocket final stage at the near site can be reduced to the maximum extent. Considering other factors such as safety of the off-orbit track, the height of the first circle of track near the ground after the last stage of the rocket is off the orbit is generally set to be 100km, and the last stage of the rocket can crash within a preset time under the action of atmospheric resistance. In addition, in order to avoid the influence of pollution and the like caused by residual fuel of the engine, the residual fuel of the engine needs to be exhausted as much as possible. Therefore, an attitude angle during derailment needs to be set, so that the rocket can be derailed with thrust force which is not completely opposite to the speed, the attitude angle corresponds to a pitch angle and a yaw angle in a navigation coordinate system, but the rocket needs to be kept in the current orbit plane during the final derailment process of the rocket, so that only one independent variable is actually needed, and the independent variable can be replaced by a flight attack angle, namely, the derailment strategy can be determined by determining the flight attack angle. The method comprises the following steps:
step S1, acquiring a first position vector, a first speed vector and available fuel mass in the rocket navigation coordinate system at the current moment, wherein the current moment is any moment of a no-power section before the rocket starts to execute the off-orbit action. After the effective load is sent into the target orbit to carry out satellite-rocket separation, in order to ensure the safety of a far field, the solid carrier rocket has a long-time unpowered gliding section, and the calculation of the off-orbit instruction can be carried out at the gliding section.
And S2, determining the maximum ignition time length of the rocket based on the available fuel quality.
Specifically, the maximum ignition time period is determined based on the available fuel mass and the rocket's final propellant specific impulse, sea level gravitational acceleration constant, and the rocket's final engine thrust, and in one embodiment, the maximum ignition time period
Wherein, I sp Specific impulse of final stage propellant of rocket g 0 The sea level gravity acceleration constant is, | u | | | is the thrust mode length, i.e. the magnitude of the thrust numerical value.
And S3, determining the actual flight attack angle of the rocket at the departure starting moment by utilizing iterative calculation based on the first position vector, the first speed vector, the maximum ignition time and a preset flight attack angle, so that the height of the near point corresponding to the close orbit element after the last stage of the rocket leaves the orbit active section is equal to the preset target height. In one embodiment, the predetermined flight angle of attack is 180 °.
Determining the actual flight angle of attack of the rocket off-orbit active segment by utilizing iterative computation based on the first position vector, the first velocity vector, the maximum ignition duration and a preset flight angle of attack, wherein the iterative computation comprises the following steps:
referring to fig. 2, in step S31, based on the preset flight attack angle, a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active segment are calculated.
Specifically, the calculating of the pitch angle and the yaw angle of the rocket in the navigation coordinate system of the off-orbit active segment based on the preset flight attack angle includes:
and step S311, determining the coordinate value of the rocket final thrust vector in the geocentric orbit coordinate system based on the trigonometric function values of the preset flight attack angle, the local trajectory inclination angle, the sideslip angle and the orbit true approach point angle.
Specifically, the coordinate value of the rocket final thrust vector in the geocentric orbit coordinate system is determined through an equation set (1), wherein the equation set (1) is as follows:
wherein the content of the first and second substances,
for local ballistic dip, β is the sideslip angle, f is the true approach angle of the orbit, and in one embodiment, β is 0 °.
And step S312, converting the rocket final stage thrust vector from the geocentric orbit coordinate system to a geocentric equatorial coordinate system.
And step S313, converting the rocket final stage thrust vector from the geocentric equator coordinate system to a launching coordinate system.
And step S314, converting the rocket final thrust vector from the launching coordinate system to a navigation coordinate system.
And step S315, calculating the pitch angle and the yaw angle of the rocket in a navigation coordinate system.
Specifically, the pitch angle and yaw angle of the rocket in a navigation coordinate system are calculated by an equation set (2), wherein the equation set (2) is as follows:
wherein the content of the first and second substances,
for the pitch angle in the navigation coordinate system,. Psi. dh 、y dh 、z dh Are coordinates within the navigational coordinate system.
Step S32 of determining the osculating rail element at the off-track end time based on the pitch angle, the yaw angle, the first position vector, the first velocity vector and the maximum ignition time.
Said determining said osculating orbit element based on said pitch angle, said yaw angle, said first position vector, said first velocity vector and said maximum firing duration at the end of off-track time comprises:
and S321, determining the departure starting time of the rocket based on the current time and the standard flight time sequence.
Step S322, based on the first position vector and the first speed vector in the navigation coordinate system of the rocket at the current moment, extrapolating to obtain a second position vector and a second speed vector in the navigation coordinate system of the rocket at the off-orbit starting moment.
Step S323, based on the pitch angle and the yaw angle, extrapolating the second position vector and the second velocity vector to obtain a third position vector and a third velocity vector of the rocket off-orbit active section ending time, wherein the rocket off-orbit active section ending time t is 1 ＝t 0 +t f ，t 0 Is the rocket off-orbit starting time, t f Is the maximum ignition time period.
In one embodiment, the method extrapolates a second position vector and a second velocity vector of the rocket in a navigation coordinate system at the off-orbit starting time by using a kinetic equation, and extrapolates a third position vector and a third velocity vector at the rocket off-orbit active segment ending time.
Specifically, the kinetic equation is:
where r is a position vector (e.g., a first position vector and a second position vector), v is a velocity vector (e.g., a first velocity vector and a second velocity vector), u is a thrust vector, | u | | | is a thrust mode length, m is a rocket final stage mass, I is sp The specific impulse of the rocket final-stage propellant is obtained, g is the gravitational acceleration borne by the rocket final stage, and the solution can be carried out according to a standard earth gravity model containing J2 terms.
And step S324, performing coordinate conversion on the third position vector and the third speed vector to obtain a position vector and a speed vector in a geocentric equatorial coordinate system of the rocket.
Step S325, determining the close track element at the off-track end time based on the position vector and the velocity vector in the geocentric equatorial coordinate system.
And step S33, determining the height of the near point corresponding to the close track element.
Specifically, the height of the near place is calculated by using a orbit determination formula based on the close orbit element. The orbit determination formula is as follows:
wherein a is orbit semimajor axis, mu is earth gravity constant, p is orbit semidiameter, e is orbit eccentricity, R E The radius of the earth.
And S34, determining that the difference between the altitude of the near place and the preset target altitude is greater than or equal to a preset value, and determining the adjustment quantity of the preset flight attack angle based on the difference.
In one embodiment, the adjustment amount of the preset flight angle of attack Δ α = (H) per1 -H target ) /10000, wherein H per1 Height of the perigee corresponding to the elements of the close track, H target And presetting a target height. In one embodiment, H target Is 100km, corresponding to a preset value of 0.1km.
And S35, adjusting the preset flight attack angle based on the adjustment amount. Specifically, the adjusted preset flight angle of attack is equal to the sum of the preset flight angle of attack calculated in the previous iteration and the adjustment amount.
Step S36, calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active section based on the adjusted preset flight attack angle, namely, entering the step S31 to the step S34 again, and determining the close orbit element at the off-orbit ending moment based on the calculated pitch angle and yaw angle, the first position vector, the first velocity vector and the maximum ignition duration; determining the height of the perigee corresponding to the close orbit element; and if the difference between the perigee height and the preset target height is determined to be larger than or equal to the preset value, determining the adjustment amount of the adjusted preset flight attack angle based on the difference, repeating the iterative calculation in the way until the difference between the perigee height corresponding to the osculating orbit element and the preset target height is determined to be smaller than the preset value, ending the iterative calculation process, and entering the step S37.
And step S37, determining the adjusted preset flight attack angle as the actual flight attack angle of the rocket at the departure starting moment.
The flight procedure angle used for navigation and guidance of the solid carrier rocket is generally defined in a navigation coordinate system, the orientation of the coordinate axis of the flight procedure angle is consistent with that of a launching inertia coordinate system, and the origin point is coincident with that of the launching inertia coordinate system at the ignition moment. The origin of the emission inertial coordinate system is an emission point, OY points to the opposite direction of the gravity of the origin, OX is perpendicular to the OY axis and points to the emission direction, and OZ is determined by a right-hand rule. And the navigation coordinate system and the emission inertial coordinate system are overlapped at the emission moment, and then the origin of the navigation coordinate system makes uniform linear motion along the earth rotation linear velocity at the origin of the ignition moment. Since the definition of the attitude angle is not affected by the position of the origin, the attitude of all navigation coordinate systems in the present invention can be equivalent to the attitude of the emission inertial coordinate system.
The definition of other coordinate systems involved in the present invention has a clear specification in the field, and the transformation matrix between coordinate systems is easily derived by those skilled in the art and will not be described in detail.
The dynamic equations involved in the invention and the standard earth gravity model containing the J2 term have clear specification definitions in the field, and the specific calculation mode is easily derived by those skilled in the art and is not described again.
Based on the same inventive concept as the method, an embodiment of the present invention further provides a rocket final-stage derailment control device, as shown in fig. 3, which is a schematic structural diagram of the embodiment of the device, and the device includes:
and the obtaining module 10 is configured to obtain a first position vector, a first velocity vector, and an available fuel mass in the rocket navigation coordinate system at a current time, where the current time is any time of a no-power section before the rocket starts to execute an off-orbit action.
A first determination module 20 for determining a maximum ignition time period for the rocket based on the available fuel mass.
Specifically, the maximum ignition time period is determined based on the available fuel mass and the rocket's final propellant specific impulse, sea level gravitational acceleration constant, and the rocket's final engine thrust, and in one embodiment, the maximum ignition time period
Wherein, I sp Specific impulse of final stage propellant of rocket g 0 The sea level gravity acceleration constant is, | u | | | is the thrust mode length, i.e. the magnitude of the thrust numerical value.
A second determining module 30, configured to determine, based on the first position vector, the first velocity vector, the maximum ignition duration, and a preset flight attack angle, an actual flight attack angle at the starting moment of the rocket derailing by using iterative computation, so that a near-location height corresponding to an osculating element after a last-stage derailing active segment of the rocket is equal to a preset target height.
Specifically, the second determining module 30 calculates a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active segment based on the preset flight attack angle; determining the osculating trajectory element at an off-track end time based on the pitch angle, the yaw angle, the first position vector, the first velocity vector and the maximum ignition duration; determining the height of the perigee corresponding to the close orbit element; determining that the difference between the height of the near place and the preset target height is greater than or equal to a preset value, and determining the adjustment quantity of the preset flight attack angle based on the difference; adjusting the preset flight attack angle based on the adjustment amount; calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active section based on the adjusted preset flight attack angle; determining the osculating orbit element at the off-track ending time based on the calculated pitch and yaw angles, the first position vector, the first velocity vector and the maximum ignition duration; determining the perigee height corresponding to the osculating orbit element; and determining that the difference value between the near-place altitude and the preset target altitude is greater than or equal to a preset value, determining the adjustment amount of the adjusted preset flight attack angle based on the difference value, repeating iterative calculation in this way until the difference value between the near-place altitude corresponding to the close orbit element and the preset target altitude is determined to be smaller than the preset value, ending the iterative calculation process, and determining the adjusted preset flight attack angle as the actual flight attack angle of the rocket at the departure starting moment.
Specifically, the second determining module 30 determines the rocket derailing starting time based on the current time and the standard flight time sequence; extrapolating to obtain a second position vector and a second velocity vector of the rocket in the navigation coordinate system at the off-orbit starting moment based on the first position vector and the first velocity vector in the navigation coordinate system of the rocket at the current moment; extrapolating the second position vector and the second velocity vector based on the pitch angle and the yaw angle to obtain a third position vector and a third velocity vector at the end time of the off-orbit active section of the rocket, wherein the end time t of the off-orbit active section of the rocket is 1 ＝t 0 +t f ，t 0 Is the rocket off-orbit starting time, t f Is the maximum ignition time period; performing coordinate conversion on the third position vector and the third speed vector to obtain a position vector and a speed vector in a geocentric equatorial coordinate system of the rocket; and determining the close track element at the off-track ending time based on the position vector and the speed vector in the geocentric equatorial coordinate system.
Based on the same inventive concept as in the previous embodiments, an embodiment of the present invention further provides a rocket final-stage derailment control device, which includes a memory, a processor, and a computer program stored in the memory and executable on the processor, and the processor implements the steps of any one of the methods described above when executing the program.
Based on the same inventive concept as in the previous embodiments, an embodiment of the present invention further provides a computer-readable storage medium, on which a computer program is stored, which, when being executed by a processor, implements the steps of any of the above-mentioned rocket final-stage derailment control methods.
The technical scheme provided by the embodiment of the invention at least has the following technical effects or advantages:
for the traditional method of binding before shooting, because various deviations in the rocket flying process can cause the difference between the actual flying state before off-orbit and the theoretical value, the binding value is used to use a fixed attitude to fix a final engine; compared with the prior art, the method has the advantages that the off-orbit ignition time length and the off-orbit effect are always different from the theoretical value, particularly when the energy of a solid engine of the solid carrier rocket is deflected downwards, the residual fuel after the final stage enters the orbit is low, and the final stage engine cannot actually work and bind the off-orbit time length during the off-orbit.
As will be appreciated by one skilled in the art, embodiments of the present invention may be provided as a method, system, or computer program product. Accordingly, the present invention may take the form of an entirely hardware embodiment, an entirely software embodiment or an embodiment combining software and hardware aspects. Furthermore, the present invention 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 present invention is described with reference to flowchart illustrations and/or block diagrams of methods, apparatus (modules, systems), and computer program products according to embodiments of the invention. It will be understood that each flow and/or block of the flowchart illustrations and/or block diagrams, and combinations of flows and/or blocks in the flowchart illustrations 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 computer, 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 preferred embodiments of the present invention 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 the preferred embodiment and all changes and modifications that fall within the scope of the invention.
It will be apparent to those skilled in the art that various changes and modifications may be made in the present invention without departing from the spirit and scope of the invention. Thus, if such modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalents, the present invention is also intended to include such modifications and variations.
Claims (9)
1. A rocket final stage derailment control method, which is characterized by comprising the following steps:
acquiring a first position vector, a first speed vector and available fuel mass in the rocket navigation coordinate system at the current moment, wherein the current moment is any moment of a powerless section before the rocket starts to execute the derailment action;
determining a maximum ignition duration for the rocket based on the available fuel mass;
determining an actual flight attack angle of the rocket at the departure starting moment by utilizing iterative calculation based on the first position vector, the first speed vector, the maximum ignition duration and a preset flight attack angle, so that the perigee height corresponding to the close orbit element after the last stage of the rocket leaves the orbit active section is equal to a preset target height;
determining the actual flight angle of attack of the rocket off-orbit active segment by utilizing iterative computation based on the first position vector, the first velocity vector, the maximum ignition duration and a preset flight angle of attack, wherein the iterative computation comprises the following steps:
calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active section based on the preset flight attack angle;
determining the osculating trajectory element at an off-track end time based on the pitch angle, the yaw angle, the first position vector, the first velocity vector and the maximum ignition duration;
determining the height of the perigee corresponding to the close orbit element;
determining that the difference between the height of the near place and the preset target height is greater than or equal to a preset value, and determining the adjustment amount of the preset flight attack angle based on the difference;
adjusting the preset flight attack angle based on the adjustment amount;
calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active section based on the adjusted preset flight attack angle; determining the osculating orbit element at the off-track ending time based on the calculated pitch and yaw angles, the first position vector, the first velocity vector and the maximum ignition duration; determining the height of the perigee corresponding to the close orbit element; and determining that the difference value between the near-place altitude and the preset target altitude is greater than or equal to a preset value, determining the adjustment amount of the adjusted preset flight attack angle based on the difference value, repeating iterative calculation in this way until the difference value between the near-place altitude corresponding to the close orbit element and the preset target altitude is determined to be smaller than the preset value, ending the iterative calculation process, and determining the adjusted preset flight attack angle as the actual flight attack angle of the rocket at the departure starting moment.
2. A rocket final stage derailment control method according to claim 1, wherein the determining the osculating trajectory element at the derailment ending time based on the pitch angle, the yaw angle, the first position vector, the first velocity vector, and the maximum ignition time period comprises:
determining the rocket derailment starting time based on the current time and a standard flight time sequence;
extrapolating to obtain a second position vector and a second velocity vector of the rocket in the navigation coordinate system at the off-orbit starting moment based on the first position vector and the first velocity vector in the navigation coordinate system of the rocket at the current moment;
extrapolating the second position vector and the second velocity vector based on the pitch angle and the yaw angle to obtain a third position vector and a third velocity vector of the rocket off-orbit active section ending time, wherein the rocket off-orbit active section ending time t 1 ＝t 0 +t f ，t 0 Is the rocket off-orbit starting time, t f Is the maximum ignition time period;
performing coordinate conversion on the third position vector and the third speed vector to obtain a position vector and a speed vector in a geocentric equatorial coordinate system of the rocket; and
and determining the close track element at the off-track ending time based on the position vector and the speed vector in the geocentric equatorial coordinate system.
3. A rocket final stage derailment control method according to claim 2, wherein the method extrapolates a second position vector and a second velocity vector of the rocket in a navigation coordinate system at the start time of the derailment using a kinetic equation, and extrapolates the third position vector and the third velocity vector at the end time of the active stage of the rocket derailment.
4. A rocket final stage derailment control method according to claim 1, wherein the determining the perigee altitude corresponding to the osculating element comprises:
and calculating to obtain the height of the perigee by utilizing a orbit determination formula based on the close orbit elements.
5. A rocket final-stage derailment control method according to claim 1, wherein the calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the derailment initiative section based on the preset flight attack angle comprises:
determining a coordinate value of a rocket final thrust vector in a geocentric orbit coordinate system based on trigonometric values of the preset flight attack angle, the local trajectory inclination angle, the sideslip angle and the orbit true approach point angle;
converting the rocket final stage thrust vector from the geocentric orbit coordinate system to a geocentric equatorial coordinate system;
converting the rocket final thrust vector from the geocentric equatorial coordinate system to a launch coordinate system;
converting the rocket final thrust vector from the launching coordinate system to a navigation coordinate system; and
calculating the pitch angle and the yaw angle of the rocket in a navigation coordinate system.
6. A rocket final derailment control method according to claim 1, wherein the determining a maximum ignition duration for the rocket based on the available fuel mass comprises:
determining the maximum ignition time period based on the available fuel mass and a final propellant thrust of the rocket, a sea level gravitational acceleration constant, and a final engine thrust of the rocket.
7. A rocket final off-track control device, characterized in that said device comprises:
the system comprises an acquisition module, a control module and a control module, wherein the acquisition module is used for acquiring a first position vector, a first speed vector and available fuel quality in a rocket navigation coordinate system at the current moment, and the current moment is any moment of a no-power section before the rocket starts to execute an off-orbit action;
a first determination module to determine a maximum ignition duration for the rocket based on the available fuel mass;
a second determining module, configured to determine, based on the first position vector, the first velocity vector, the maximum ignition duration, and a preset flight attack angle, an actual flight attack angle at the starting point of the rocket derailing by using iterative computation, so that a near-point altitude corresponding to an osculating element after a last-stage derailing active segment of the rocket is equal to a preset target altitude;
the second determining module calculates a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active section based on the preset flight attack angle;
determining the osculating trajectory element at an off-track end time based on the pitch angle, the yaw angle, the first position vector, the first velocity vector and the maximum ignition duration;
determining the height of the perigee corresponding to the close orbit element;
determining that the difference between the height of the near place and the preset target height is greater than or equal to a preset value, and determining the adjustment amount of the preset flight attack angle based on the difference;
adjusting the preset flight attack angle based on the adjustment amount;
calculating a pitch angle and a yaw angle of the rocket in a navigation coordinate system of the off-orbit active section based on the adjusted preset flight attack angle; determining the osculating rail element at the end of off-track based on the calculated pitch and yaw angles, the first position vector, the first velocity vector and the maximum ignition duration; determining the height of the perigee corresponding to the close orbit element; and determining that the difference between the perigee height and the preset target height is greater than or equal to a preset value, determining the adjustment amount of the adjusted preset flight attack angle based on the difference, repeating iterative calculation in the way until the difference between the perigee height corresponding to the osculating orbit element and the preset target height is smaller than the preset value, ending the iterative calculation process, and determining the adjusted preset flight attack angle as the actual flight attack angle of the rocket at the departure starting moment.
8. A rocket final derailment control apparatus, comprising:
a memory for storing a computer program;
a processor for executing the computer program to carry out the steps of the method of any one of claims 1 to 6.
9. A computer-readable storage medium, on which a computer program is stored, which, when being executed by a processor, is adapted to carry out the steps of the method of any one of claims 1 to 6.
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| CN114036780A (en) * | 2021-12-06 | 2022-02-11 | 航天科工火箭技术有限公司 | Rocket attitude angle design method constrained by space-based measurement and control |
| CN114281090B (en) * | 2021-12-17 | 2023-09-22 | 上海宇航系统工程研究所 | Rocket tail sub-level precise passivation emission control method and system |
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| CN114802819B (en) * | 2022-07-01 | 2022-10-14 | 星河动力(北京)空间科技有限公司 | Rocket terminal-level off-orbit control method and device |
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| CN116861811B (en) * | 2023-07-04 | 2024-02-02 | 东方空间技术(山东)有限公司 | Rocket final stage off-orbit thrust determination method, device and equipment |
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