CN116185058B - Carrier rocket attitude control method and device and flight control computer - Google Patents
Carrier rocket attitude control method and device and flight control computer Download PDFInfo
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- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
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Abstract
The invention provides a carrier rocket attitude control method, a carrier rocket attitude control device and a flight control computer, and relates to the technical field of rocket solid booster attitude control, wherein the carrier rocket attitude control method comprises the following steps: acquiring at least one interference moment of a booster of the carrier rocket in a thrust descending stage; determining the ignition time of the core engine when the flight attitude of the carrier rocket changes according to at least one interference moment and the control moment provided by the booster of the carrier rocket in the thrust descending stage; acquiring a flight attitude control scheme; and controlling the flight attitude of the carrier rocket according to the flight attitude control scheme according to the ignition time. The invention can make the attitude control of the carrier rocket in the boosting flight section more stable.
Description
Technical Field
The invention relates to the technical field of rocket solid booster attitude control, in particular to a carrier rocket attitude control method and device and a flight control computer.
Background
At present, the launch frequency of the carrier rocket is higher and higher, the thrust requirement on the rocket is also increased along with the launch frequency, and in order to meet the requirement, more and more rockets adopt a booster binding structure so as to improve the carrying capacity.
The rocket engines with smaller thrust are bundled around a single-stage rocket engine (core stage) with large thrust to form a bundled rocket, and the bundled rocket engines with smaller thrust are called thrusters, so that larger thrust can be generated.
Typically, the booster and core engine will fire at the same time, providing thrust to the rocket. At the moment, the posture of the rocket can be controlled by adjusting the swing angles of different thrusters and the swing angles of the spray pipes of the core-stage engine. At the end of the flight phase, insufficient control and increased control disturbances are caused by the unsynchronized or other deviations in the combustion time of the booster.
Disclosure of Invention
The technical problem to be solved by the invention is to provide a carrier rocket attitude control method, a carrier rocket attitude control device and a flight control computer, so that the stability of controlling flight can be achieved, the attitude control of the carrier rocket in a boosting flight section can be more stable, and the requirements are met.
In order to solve the technical problems, the technical scheme of the invention is as follows:
in a first aspect, a method for controlling attitude of a launch vehicle, the method comprising:
acquiring at least one interference moment of a booster of the carrier rocket in a thrust descending stage;
determining the ignition time of the core engine when the flight attitude of the carrier rocket changes according to at least one interference moment and the control moment provided by the booster of the carrier rocket in the thrust descending stage;
acquiring a flight attitude control scheme;
and controlling the flight attitude of the carrier rocket according to the flight attitude control scheme according to the ignition time.
Further, acquiring at least one disturbance moment of the booster of the carrier rocket in the thrust descending stage comprises the following steps:
at least one of the engine axis deflection aerodynamic disturbance moment, the disturbance moment generated by the mass center deviation axis, the disturbance moment generated by the thrust line sideslip, the disturbance moment generated by the thrust line deflection and the disturbance moment of the engine thrust descent section relative to the rated value of the booster of the carrier rocket in the thrust descent stage is obtained.
Further, acquiring the engine axis deflection aerodynamic disturbance moment of the booster of the carrier rocket in the thrust descending stage comprises the following steps:
according to
Acquiring the engine axis deflection pneumatic interference moment of a booster of the carrier rocket in a thrust descending stage;
wherein,,
for the skew aerodynamic disturbance moment of the engine axis, +.>
；
Further, acquiring a disturbance moment generated by a centroid of a booster of the carrier rocket in a thrust descent stage deviating from an axis comprises:
according to
Acquiring interference moment generated by the fact that the mass center of a booster of the carrier rocket deviates from an axis in a thrust descending stage;
wherein,,
disturbance moment generated for centroid off-axis, +.>
Is engine thrust; />
Is the distance of the thrust from the theoretical axis.
Further, obtaining a disturbing moment generated by the thrust line traversing of the booster of the carrier rocket in the thrust descending stage, the method comprises the following steps:
according to
Acquiring interference moment generated by the thrust line traversing of a booster of the carrier rocket in a thrust descending stage; wherein (1)>
The interference moment generated by the thrust line transverse movement of the booster in the thrust descending stage is provided;
the method for acquiring the interference moment generated by the thrust line deflection of the booster of the carrier rocket in the thrust descending stage comprises the following steps:
according to
Acquiring interference moment generated by thrust line deflection of a booster of the carrier rocket in a thrust descending stage;
wherein,,
for the disturbance moment generated by the thrust line deflection of the booster in the thrust down phase, +.>
For engine thrust +.>
For thrust deflection angle->
Is the distance between the rocket centroid and the theoretical vertex, +.>
Is the distance from the engine nozzle swing point to the theoretical vertex.
Further, determining an ignition time of the core engine when the flight attitude of the carrier rocket changes according to at least one of the disturbance moment and a control moment provided by the booster of the carrier rocket in the thrust descending stage, including:
obtaining a difference value between a control moment and an interference moment provided by a booster of the carrier rocket in a thrust descending stage;
judging whether a control moment provided by a booster of the carrier rocket in a thrust descending stage counteracts an interference moment or not according to the difference value, if so, keeping the original flight attitude of the carrier rocket, and if not, acquiring a time point when the flight attitude of the carrier rocket changes;
and determining the ignition time of the core engine according to the time point and the time interval of the core engine capable of providing the preset control force.
Further, acquiring a flight attitude control scheme includes:
determining a static amplification coefficient of a boosting flight section of a core engine ignition time point, a format of a control network of a boosting stable flight section, and a swing angle amplitude limit of an engine spray pipe of the boosting stable flight section and a swing angle amplitude limit of the engine spray pipe of the core engine;
and obtaining the flight attitude control scheme according to the static amplification factor of the boosting flight section, the format of a control network of the boosting stable flight section, the swing angle amplitude limit of the engine spray pipe of the boosting stable flight section and the swing angle amplitude limit of the core engine spray pipe.
In a second aspect, a launch vehicle attitude control device includes:
the acquisition module is used for acquiring at least one interference moment of the booster of the carrier rocket in the thrust descending stage;
the processing module is used for determining the ignition time of the core engine when the flight attitude of the carrier rocket changes according to at least one of the interference moment and the control moment provided by the booster of the carrier rocket in the thrust descending stage; acquiring a flight attitude control scheme; and controlling the flight attitude of the carrier rocket according to the flight attitude control scheme according to the ignition time.
In a third aspect, a launch vehicle flight control computer comprises:
one or more processors;
and a storage means for storing one or more programs that, when executed by the one or more processors, cause the one or more processors to implement the method.
In a fourth aspect, a computer readable storage medium has a program stored therein, which when executed by a processor, implements the method.
The scheme of the invention at least comprises the following beneficial effects:
according to the scheme, the ignition time of the core engine when the flight attitude of the carrier rocket changes is determined according to at least one interference moment and the control moment provided by the booster of the carrier rocket in the thrust descending stage, and the flight attitude control scheme is obtained and the flight attitude of the carrier rocket is controlled according to the ignition time according to the flight attitude control scheme, so that the attitude control of the carrier rocket in the booster flight stage is more stable, and the requirements are met.
Drawings
Fig. 1 is a schematic flow chart of a method for controlling the attitude of a carrier rocket according to an embodiment of the present invention.
Fig. 2 is a schematic diagram of a carrier rocket attitude control device according to an embodiment of the present invention.
Detailed Description
Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be embodied in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art.
As shown in fig. 1, an embodiment of the present invention proposes a method for controlling the attitude of a carrier rocket, including the following steps:
step 11: acquiring at least one interference moment of a booster of the carrier rocket in a thrust descending stage;
step 12: determining the ignition time of the core engine when the flight attitude of the carrier rocket changes according to at least one interference moment and the control moment provided by the booster of the carrier rocket in the thrust descending stage;
step 13: acquiring a flight attitude control scheme;
step 14: and controlling the flight attitude of the carrier rocket according to the flight attitude control scheme according to the ignition time.
In the embodiment of the invention, the booster ignites first in the flight process of the carrier rocket, and when the booster is exhausted or the thrust is reduced, the core engine ignites again, so that the ignition time of the core engine when the flight attitude of the carrier rocket changes is determined according to at least one of the interference moment and the control moment provided by the booster of the carrier rocket in the thrust reduction stage, thereby the ignition time of the core engine can be controlled at proper time, the resource is saved, and the flight attitude of the carrier rocket is controlled according to the flight attitude control scheme by acquiring the flight attitude control scheme and the ignition time, so that the attitude in the boosting flight stage meets the control requirement, and the flight attitude of the carrier rocket is more stable.
In the invention, the carrier rocket is a solid booster carrier rocket, and particularly has a structure of binding four solid boosters around a core engine, wherein the carrier rocket can be a missile or a carrier rocket, and when the missile or the carrier rocket flies, the flying attitude of the missile or the carrier rocket is regulated and controlled by an attitude control system, so the attitude control system is an important component of a carrier rocket (missile or carrier rocket) flying control system, and the function of the carrier rocket is to stabilize and control the angular movement of the carrier rocket around the mass center. In the boosting flight stage, after the attitude angle deviation and the attitude angle speed are respectively processed by an output correction network of the flight control, the servo mechanism of the booster is controlled to swing so as to obtain a control moment, and the attitude stability control of the boosting flight stage is realized.
In another alternative embodiment of the present invention, the step 11 may include:
step 111: at least one of the engine axis deflection aerodynamic disturbance moment, the disturbance moment generated by the mass center deviation axis, the disturbance moment generated by the thrust line sideslip, the disturbance moment generated by the thrust line deflection and the disturbance moment of the engine thrust descent section relative to the rated value of the booster of the carrier rocket in the thrust descent stage is obtained.
In this embodiment, in the thrust lowering stage, the thrust of the booster is disturbed due to the installation deviation, the thrust line deviation, the working time deviation, and the like between the opposite installed boosters, so that the stable attitude of the carrier rocket is affected, and therefore, in order to realize the stable flying attitude of the carrier rocket, it is necessary to calculate disturbance moment caused by the installation deviation, the thrust line deviation, the working time deviation, and the like, and to analyze and calculate various deviation conditions, so that the stable control of the attitude of the booster flying stage can be realized.
In another optional embodiment of the present invention, in the step 111, acquiring the engine axis skew aerodynamic disturbance moment of the booster of the carrier rocket in the thrust down stage may include:
according to
Acquiring the engine axis deflection pneumatic interference moment of a booster of the carrier rocket in a thrust descending stage;
wherein,,
for the skew aerodynamic disturbance moment of the engine axis, +.>
；
In the embodiment, the engine axis deflection pneumatic interference moment of the booster of the carrier rocket in the thrust descending stage is calculated, and the engine axis deflection pneumatic interference moment can be analyzed, so that the ignition time of the core-level engine when the flight attitude of the carrier rocket changes can be calculated, and the stable control of the attitude of the booster flight section of the carrier rocket is realized.
In another optional embodiment of the present invention, in the step 111, acquiring the disturbing moment generated by the center of mass of the booster of the carrier rocket during the thrust lowering stage, may include:
according to
Acquiring interference moment generated by the fact that the mass center of a booster of the carrier rocket deviates from an axis in a thrust descending stage;
wherein,,
disturbance moment generated for centroid off-axis, +.>
Is engine thrust; />
Is the distance of the thrust from the theoretical axis.
In the embodiment, the interference moment generated by the mass center of the booster of the carrier rocket in the thrust descending stage deviating from the axis is calculated, and the pneumatic interference moment of the deflection of the engine axis can be analyzed, so that the ignition time of the core-level engine when the flight attitude of the carrier rocket changes can be calculated, and the stable control of the attitude of the booster flight section of the carrier rocket is realized.
In another optional embodiment of the present invention, in the step 111, acquiring the disturbing moment generated by the thrust line traversing of the booster of the carrier rocket in the thrust down stage may include:
according to
Acquiring interference moment generated by the thrust line traversing of a booster of the carrier rocket in a thrust descending stage; wherein (1)>
Interference moment generated for thrust line traversing of booster in thrust lowering stage +.>
For engine thrust +.>
The distance is the thrust line traversing distance;
in step 111, obtaining a disturbing moment generated by the thrust line deflection of the booster of the carrier rocket in the thrust down stage includes:
according to
Acquiring interference moment generated by thrust line deflection of a booster of the carrier rocket in a thrust descending stage;
wherein,,
for the disturbance moment generated by the thrust line deflection of the booster in the thrust down phase, +.>
For engine thrust +.>
For thrust deflection angle->
Is the distance between the rocket centroid and the theoretical vertex, +.>
Is the distance from the engine nozzle swing point to the theoretical vertex.
In this embodiment, by calculating the disturbance moment generated by the thrust line deflection of the booster in the thrust lowering stage, and analyzing the disturbance moment generated by the thrust line deflection of the booster in the thrust lowering stage, the ignition time of the core engine when the flight attitude of the carrier rocket changes can be calculated, so as to realize stable control of the attitude of the carrier rocket booster flight stage, wherein when calculating, the maximum disturbance moment is obtained by superposing the disturbance moment generated by the deviation of the engine axis from the pneumatic disturbance moment, the disturbance moment generated by the deviation of the center of mass from the axis, and the disturbance moment generated by the thrust line sideslip of the booster in the thrust lowering stage, wherein the maximum disturbance moment is:
in another alternative embodiment of the present invention, the step 12 may include:
step 121: obtaining a difference value between a control moment and an interference moment provided by a booster of the carrier rocket in a thrust descending stage;
step 122: judging whether the control moment provided by the booster of the carrier rocket in the thrust descending stage can offset the interference moment according to the difference value, if so, keeping the original flight attitude of the carrier rocket, and if not, acquiring the time point when the flight attitude of the carrier rocket changes;
step 123: and determining the ignition time of the core engine according to the time point and the time interval of the core engine capable of providing the preset control force.
In this embodiment, in order to determine whether the control moment provided by the booster of the carrier rocket in the thrust descending stage can offset the disturbance moment, it is required to specify that the disturbance moment is the maximum disturbance moment, so that it is required to calculate the control moment provided by the booster of the carrier rocket in the thrust descending stage and the maximum disturbance moment, and then determine the time point when the flight attitude of the carrier rocket changes according to the difference between the control moment provided by the booster of the carrier rocket in the thrust descending stage and the maximum disturbance moment, for example, if the difference is greater than or equal to 0, that is, the control moment provided by the booster of the carrier rocket in the thrust descending stage can offset the maximum disturbance moment, the flight attitude of the carrier rocket is stable, and in order to obtain the time point when the flight attitude of the carrier rocket changes, it is also required to keep the original flight attitude until the flight attitude of the carrier rocket changes, and obtain the time point when the flight attitude of the carrier rocket changes; if the difference is less than 0, that is, the control moment provided by the booster of the carrier rocket in the thrust descending stage cannot offset the maximum interference moment, the flight attitude of the carrier rocket is changed, and meanwhile, the time point of the change of the flight attitude of the carrier rocket is required to be acquired.
It should be noted that, when determining the ignition time of the core engine, it is also necessary to determine whether the core engine can provide sufficient control force, because it is generally not possible to select to ignite the core engine at the right time point, and it is necessary to consider the time interval from ignition to providing sufficient control force of the core engine, so when determining the ignition time of the core engine, by calculating the time interval at which the core engine can provide the preset control force, the ignition time point of the core engine can be more accurate, thereby saving resources.
In another alternative embodiment of the present invention, the step 13 may include:
step 131: determining a static amplification coefficient of a boosting flight section of a core engine ignition time point, a format of a control network of a boosting stable flight section, and a swing angle amplitude limit of an engine spray pipe of the boosting stable flight section and a swing angle amplitude limit of the engine spray pipe of the core engine;
step 132: and obtaining the flight attitude control scheme according to the static amplification factor of the boosting flight section, the format of a control network of the boosting stable flight section, the swing angle amplitude limit of the engine spray pipe of the boosting stable flight section and the swing angle amplitude limit of the core engine spray pipe.
In the embodiment, the static amplification factor of the boosting flight section at the ignition time point of the core engine is determined through calculation, and the static amplification factor of the boosting flight section is kept unchanged in the whole descending stage; the format of the control network of the boosting stable flight section and the parameters of the control network are determined, and the format of the control network of the whole descending stage can be directly used for boosting the format of the control network of the stable flight section, and the design can be carried out for the special flight period; and the swing angle limiting of the engine spray pipe in the boosting stable flight section and the swing angle limiting of the engine spray pipe in the core stage ensure the safety of the swing angle of the engine in the control section (wherein the swing angle limiting of the engine spray pipe in the boosting stable flight section is the same as that in the stable section), after the ignition of the engine in the core stage is determined, the swing angle of a step signal can be prevented from appearing by controlling the swing angle in the parameter control process, and the interference of the generation of the abrupt force to the gesture is avoided.
In another embodiment of the present invention, after step 14, the carrier rocket attitude control method may further include:
step 15: and simulating the preset control force provided by the core-level engine, judging whether the preset control force meets the preset control force not less than the interference moment, if so, enabling the flight attitude control scheme to be correct, otherwise, enabling the flight attitude control scheme to be incorrect, and readjusting the new flight attitude control scheme.
In the above embodiment of the present invention, it should be noted that, the disturbing moment is the largest disturbing moment, and whether the control force provided by the rocket can counteract the largest disturbing moment and provide enough stable control force is determined by determining whether the control force provided by the rocket in the thrust descending section of the booster, so as to determine whether the core engine ignites; after the ignition of the core-level engine is combined, judging the interval duration capable of providing enough control force and judging whether the requirement can be met; and designing a corresponding gesture control strategy; and then judging whether the attitude control strategy meets the requirement of attitude stabilization, so that an attitude control system in the boosting flight section meets the control requirement.
As shown in fig. 2, an embodiment of the present invention further provides a carrier rocket attitude control device 20, including:
an acquisition module 21, configured to acquire at least one disturbance moment of the booster of the carrier rocket in a thrust descent phase;
a processing module 22, configured to determine an ignition time of the core engine when a flight attitude of the carrier rocket changes according to at least one of the disturbance torque and a control torque provided by a booster of the carrier rocket in a thrust down stage; acquiring a flight attitude control scheme; and controlling the flight attitude of the carrier rocket according to the flight attitude control scheme according to the ignition time.
Optionally, acquiring at least one disturbing moment of the booster of the carrier rocket in the thrust descending stage includes:
at least one of the engine axis deflection aerodynamic disturbance moment, the disturbance moment generated by the mass center deviation axis, the disturbance moment generated by the thrust line sideslip, the disturbance moment generated by the thrust line deflection and the disturbance moment of the engine thrust descent section relative to the rated value of the booster of the carrier rocket in the thrust descent stage is obtained.
Optionally, acquiring the engine axis skew aerodynamic disturbance moment of the booster of the carrier rocket in the thrust descending stage includes:
according to
Acquiring the engine axis deflection pneumatic interference moment of a booster of the carrier rocket in a thrust descending stage;
wherein,,
for the skew aerodynamic disturbance moment of the engine axis, +.>
；
Optionally, acquiring the disturbing moment generated by the center of mass of the booster of the carrier rocket in the thrust descending stage deviating from the axis comprises:
according to
Acquiring interference moment generated by the fact that the mass center of a booster of the carrier rocket deviates from an axis in a thrust descending stage;
wherein,,
disturbance moment generated for centroid off-axis, +.>
Is engine thrust; />
Is the distance of the thrust from the theoretical axis.
Optionally, acquiring a disturbing moment generated by the thrust line traversing of the booster of the carrier rocket in the thrust descending stage includes:
according to
Acquiring interference moment generated by the thrust line traversing of a booster of the carrier rocket in a thrust descending stage; wherein (1)>
Interference moment generated for thrust line traversing of booster in thrust lowering stage +.>
For engine thrust +.>
The distance is the thrust line traversing distance;
the method for acquiring the interference moment generated by the thrust line deflection of the booster of the carrier rocket in the thrust descending stage comprises the following steps:
according to
Acquiring interference moment generated by thrust line deflection of a booster of the carrier rocket in a thrust descending stage;
wherein,,
for the disturbance moment generated by the thrust line deflection of the booster in the thrust down phase, +.>
For engine thrust +.>
For thrust deflection angle->
Is the distance between the rocket centroid and the theoretical vertex, +.>
Is the distance from the engine nozzle swing point to the theoretical vertex.
Optionally, determining the ignition time of the core engine when the flight attitude of the carrier rocket changes according to at least one of the disturbance moment and the control moment provided by the booster of the carrier rocket in the thrust descending stage includes:
obtaining a difference value between a control moment and an interference moment provided by a booster of the carrier rocket in a thrust descending stage;
judging whether a control moment provided by a booster of the carrier rocket in a thrust descending stage counteracts an interference moment or not according to the difference value, if so, keeping the original flight attitude of the carrier rocket, and if not, acquiring a time point when the flight attitude of the carrier rocket changes;
and determining the ignition time of the core engine according to the time point and the time interval of the core engine capable of providing the preset control force.
Optionally, acquiring a flight attitude control scheme includes:
determining a static amplification coefficient of a boosting flight section of a core engine ignition time point, a format of a control network of a boosting stable flight section, and a swing angle amplitude limit of an engine spray pipe of the boosting stable flight section and a swing angle amplitude limit of the engine spray pipe of the core engine;
and obtaining the flight attitude control scheme according to the static amplification factor of the boosting flight section, the format of a control network of the boosting stable flight section, the swing angle amplitude limit of the engine spray pipe of the boosting stable flight section and the swing angle amplitude limit of the core engine spray pipe.
It should be noted that the apparatus is an apparatus corresponding to the above method, and all implementation manners in the above method embodiment are applicable to this embodiment, so that the same technical effects can be achieved.
The embodiment of the invention also provides a carrier rocket flight control computer, which comprises: a processor, a memory storing a computer program which, when executed by the processor, performs the method as described above. All the implementation manners in the method embodiment are applicable to the embodiment, and the same technical effect can be achieved.
Embodiments of the present invention also provide a computer-readable storage medium storing instructions that, when executed on a computer, cause the computer to perform a method as described above. All the implementation manners in the method embodiment are applicable to the embodiment, and the same technical effect can be achieved.
Those of ordinary skill in the art will appreciate that the various illustrative elements and algorithm steps described in connection with the embodiments disclosed herein may be implemented as electronic hardware, or combinations of computer software and electronic hardware. Whether such functionality is implemented as hardware or software depends upon the particular application and design constraints imposed on the solution. Skilled artisans may implement the described functionality in varying ways for each particular application, but such implementation decisions should not be interpreted as causing a departure from the scope of the present invention.
It will be clear to those skilled in the art that, for convenience and brevity of description, specific working procedures of the above-described systems, apparatuses and units may refer to corresponding procedures in the foregoing method embodiments, and are not repeated herein.
In the embodiments provided in the present invention, it should be understood that the disclosed apparatus and method may be implemented in other manners. For example, the apparatus embodiments described above are merely illustrative, e.g., the division of the units is merely a logical function division, and there may be additional divisions when actually implemented, e.g., multiple units or components may be combined or integrated into another system, or some features may be omitted or not performed. Alternatively, the coupling or direct coupling or communication connection shown or discussed with each other may be an indirect coupling or communication connection via some interfaces, devices or units, which may be in electrical, mechanical or other form.
The units described as separate units may or may not be physically separate, and units shown as units may or may not be physical units, may be located in one place, or may be distributed on a plurality of network units. Some or all of the units may be selected according to actual needs to achieve the purpose of the solution of this embodiment.
In addition, each functional unit in the embodiments of the present invention may be integrated in one processing unit, or each unit may exist alone physically, or two or more units may be integrated in one unit.
The functions, if implemented in the form of software functional units and sold or used as a stand-alone product, may be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present invention 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 instructions for causing a computer device (which may be a personal computer, a server, a network device, etc.) to perform all or part of the steps of the method according to the embodiments of the present invention. And the aforementioned storage medium includes: a usb disk, a removable hard disk, a ROM, a RAM, a magnetic disk, or an optical disk, etc.
Furthermore, it should be noted that in the apparatus and method of the present invention, it is apparent that the components or steps may be disassembled and/or assembled. Such decomposition and/or recombination should be considered as equivalent aspects of the present invention. Also, the steps of performing the series of processes described above may naturally be performed in chronological order in the order of description, but are not necessarily performed in chronological order, and some steps may be performed in parallel or independently of each other. It will be appreciated by those of ordinary skill in the art that all or any of the steps or components of the methods and apparatus of the present invention may be implemented in hardware, firmware, software, or a combination thereof in any computing device (including processors, storage media, etc.) or network of computing devices, as would be apparent to one of ordinary skill in the art after reading this description of the invention.
The object of the invention can thus also be achieved by running a program or a set of programs on any computing device. The computing device may be a well-known general purpose device. The object of the invention can thus also be achieved by merely providing a program product containing program code for implementing said method or apparatus. That is, such a program product also constitutes the present invention, and a storage medium storing such a program product also constitutes the present invention. It is apparent that the storage medium may be any known storage medium or any storage medium developed in the future. It should also be noted that in the apparatus and method of the present invention, it is apparent that the components or steps may be disassembled and/or assembled. Such decomposition and/or recombination should be considered as equivalent aspects of the present invention. The steps of executing the series of processes may naturally be executed in chronological order in the order described, but are not necessarily executed in chronological order. Some steps may be performed in parallel or independently of each other.
While the foregoing is directed to the preferred embodiments of the present invention, it will be appreciated by those skilled in the art that various modifications and adaptations can be made without departing from the principles of the present invention, and such modifications and adaptations are intended to be comprehended within the scope of the present invention.
Claims (8)
1. A method for controlling the attitude of a launch vehicle, the method comprising:
acquiring at least one interference moment of a booster of the carrier rocket in a thrust descending stage;
according to at least one interference moment and a control moment provided by a booster of the carrier rocket in a thrust descending stage, determining the ignition time of a core engine when the flight attitude of the carrier rocket changes, wherein the method comprises the steps of obtaining a difference value between the control moment provided by the booster of the carrier rocket in the thrust descending stage and the interference moment; judging whether a control moment provided by a booster of the carrier rocket in a thrust descending stage counteracts an interference moment or not according to the difference value, if so, keeping the original flight attitude of the carrier rocket, and if not, acquiring a time point when the flight attitude of the carrier rocket changes; determining the ignition time of the core engine according to the time point and the time interval at which the core engine can provide preset control force;
the method comprises the steps of obtaining a flight attitude control scheme, wherein the flight attitude control scheme comprises the steps of determining a static amplification coefficient of a boosting flight section of a core engine ignition time point, a format of a control network of a boosting stable flight section, and an engine nozzle swing angle amplitude limit and a core engine nozzle swing angle amplitude limit of the boosting stable flight section; obtaining the flight attitude control scheme according to the static amplification factor of the boosting flight section, the format of a control network of the boosting stable flight section, the swing angle amplitude limit of the engine spray pipe of the boosting stable flight section and the swing angle amplitude limit of the core engine spray pipe;
and controlling the flight attitude of the carrier rocket according to the flight attitude control scheme according to the ignition time.
2. A method of controlling the attitude of a launch vehicle according to claim 1, wherein obtaining at least one disturbance moment of a booster of the launch vehicle during a thrust-down phase comprises:
at least one of the engine axis deflection aerodynamic disturbance moment, the disturbance moment generated by the mass center deviation axis, the disturbance moment generated by the thrust line sideslip, the disturbance moment generated by the thrust line deflection and the disturbance moment of the engine thrust descent section relative to the rated value of the booster of the carrier rocket in the thrust descent stage is obtained.
3. A method of controlling the attitude of a launch vehicle according to claim 2, wherein obtaining the engine axis skew aerodynamic disturbance torque of the booster of the launch vehicle during the thrust-down phase comprises:
according to
Acquiring the engine axis deflection pneumatic interference moment of a booster of the carrier rocket in a thrust descending stage;
wherein,,
for the skew aerodynamic disturbance moment of the engine axis, +.>
；
4. A method of controlling the attitude of a launch vehicle according to claim 2, wherein obtaining the disturbance moment generated by the thrust-assist device of the launch vehicle during the thrust-down phase, which is off-axis, comprises:
according to
Acquiring interference moment generated by the fact that the mass center of a booster of the carrier rocket deviates from an axis in a thrust descending stage;
wherein,,
disturbance moment generated for centroid off-axis, +.>
Is engine thrust; />
Is the distance of the thrust from the theoretical axis.
5. The method of claim 2, wherein obtaining the disturbance moment generated by the thrust line traversing of the booster of the launch vehicle in the thrust down stage comprises:
according to
Acquiring interference moment generated by the thrust line traversing of a booster of the carrier rocket in a thrust descending stage; wherein (1)>
The interference moment generated by the thrust line transverse movement of the booster in the thrust descending stage is provided;
the method for acquiring the interference moment generated by the thrust line deflection of the booster of the carrier rocket in the thrust descending stage comprises the following steps:
according to
Acquiring interference moment generated by thrust line deflection of a booster of the carrier rocket in a thrust descending stage;
wherein,,
for the disturbance moment generated by the thrust line deflection of the booster in the thrust down phase, +.>
For the thrust of the engine,
for thrust deflection angle->
Is the distance between the rocket centroid and the theoretical vertex, +.>
Is the distance from the engine nozzle swing point to the theoretical vertex.
6. A launch vehicle attitude control device, comprising:
the acquisition module is used for acquiring at least one interference moment of the booster of the carrier rocket in the thrust descending stage;
the processing module is used for determining the ignition time of the core-stage engine when the flight attitude of the carrier rocket changes according to at least one interference moment and the control moment provided by the booster of the carrier rocket in the thrust descending stage, and comprises the steps of obtaining the difference value between the control moment provided by the booster of the carrier rocket in the thrust descending stage and the interference moment; judging whether a control moment provided by a booster of the carrier rocket in a thrust descending stage counteracts an interference moment or not according to the difference value, if so, keeping the original flight attitude of the carrier rocket, and if not, acquiring a time point when the flight attitude of the carrier rocket changes; determining the ignition time of the core engine according to the time point and the time interval at which the core engine can provide preset control force; the method comprises the steps of obtaining a flight attitude control scheme, wherein the flight attitude control scheme comprises the steps of determining a static amplification coefficient of a boosting flight section of a core engine ignition time point, a format of a control network of a boosting stable flight section, and an engine nozzle swing angle amplitude limit and a core engine nozzle swing angle amplitude limit of the boosting stable flight section; obtaining the flight attitude control scheme according to the static amplification factor of the boosting flight section, the format of a control network of the boosting stable flight section, the swing angle amplitude limit of the engine spray pipe of the boosting stable flight section and the swing angle amplitude limit of the core engine spray pipe; and controlling the flight attitude of the carrier rocket according to the flight attitude control scheme according to the ignition time.
7. A launch vehicle flight control computer comprising:
one or more processors;
storage means for storing one or more programs that when executed by the one or more processors cause the one or more processors to implement the method of any of claims 1-5.
8. A computer readable storage medium, characterized in that the computer readable storage medium has stored therein a program which, when executed by a processor, implements the method according to any of claims 1-5.
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| CN202310429817.XA CN116185058B (en) | 2023-04-21 | 2023-04-21 | Carrier rocket attitude control method and device and flight control computer |
| PCT/CN2023/119878 WO2024216836A1 (en) | 2023-04-21 | 2023-09-20 | Carrier rocket attitude control method and device, and flight control computer |
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| CN202310429817.XA CN116185058B (en) | 2023-04-21 | 2023-04-21 | Carrier rocket attitude control method and device and flight control computer |
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| CN116185058B (en) * | 2023-04-21 | 2023-07-07 | 东方空间技术(山东)有限公司 | Carrier rocket attitude control method and device and flight control computer |
| CN117555227B (en) * | 2023-06-05 | 2024-03-29 | 东方空间技术(山东)有限公司 | Control method, computing equipment and storage medium of carrier rocket |
| CN116400723B (en) * | 2023-06-07 | 2023-09-01 | 东方空间技术(山东)有限公司 | Carrier rocket load shedding control method, computing equipment and storage medium |
| CN116859981B (en) * | 2023-09-05 | 2023-12-15 | 东方空间技术(山东)有限公司 | Carrier rocket attitude control method and device and computing equipment |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4901949A (en) * | 1988-03-11 | 1990-02-20 | Orbital Sciences Corporation Ii | Rocket-powered, air-deployed, lift-assisted booster vehicle for orbital, supraorbital and suborbital flight |
| CN104267733A (en) * | 2014-10-25 | 2015-01-07 | 哈尔滨工业大学 | Attitude control type direct lateral force and aerodynamic force composite missile attitude control method based on mixed forecasting control |
| CN115755646A (en) * | 2022-12-07 | 2023-03-07 | 北京中科宇航技术有限公司 | Carrier rocket attitude control system simulation method and system |
Family Cites Families (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP5366607B2 (en) * | 2009-03-25 | 2013-12-11 | 株式会社Ｉｈｉエアロスペース | Rocket release method and device |
| CN110794863B (en) * | 2019-11-20 | 2021-05-28 | 中山大学 | Heavy carrier rocket attitude control method capable of customizing control performance indexes |
| US11353301B2 (en) * | 2020-01-22 | 2022-06-07 | Raytheon Company | Kinetic energy vehicle with attitude control system having paired thrusters |
| DE102020126575B4 (en) * | 2020-10-09 | 2024-03-14 | Arianegroup Gmbh | Launch vehicle system with launcher and launch aid unit |
| CN112987767B (en) * | 2021-01-19 | 2023-07-25 | 中国人民解放军63921部队 | Carrier rocket attitude control method with integrated boosting and core level |
| CN114384799B (en) * | 2022-01-14 | 2023-11-28 | 北京中科宇航技术有限公司 | Combined thrust vector control method for boosting and core-level engine |
| CN114859956B (en) * | 2022-07-05 | 2022-11-18 | 星河动力(北京)空间科技有限公司 | Control method, device and equipment of carrier rocket and storage medium |
| CN115203963A (en) * | 2022-07-26 | 2022-10-18 | 航天科工火箭技术有限公司 | Method, device, equipment and medium for identifying equivalent offset of engine thrust line |
| CN116185058B (en) * | 2023-04-21 | 2023-07-07 | 东方空间技术(山东)有限公司 | Carrier rocket attitude control method and device and flight control computer |
-
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- 2023-04-21 CN CN202310429817.XA patent/CN116185058B/en active Active
- 2023-09-20 WO PCT/CN2023/119878 patent/WO2024216836A1/en unknown
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4901949A (en) * | 1988-03-11 | 1990-02-20 | Orbital Sciences Corporation Ii | Rocket-powered, air-deployed, lift-assisted booster vehicle for orbital, supraorbital and suborbital flight |
| CN104267733A (en) * | 2014-10-25 | 2015-01-07 | 哈尔滨工业大学 | Attitude control type direct lateral force and aerodynamic force composite missile attitude control method based on mixed forecasting control |
| CN115755646A (en) * | 2022-12-07 | 2023-03-07 | 北京中科宇航技术有限公司 | Carrier rocket attitude control system simulation method and system |
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| WO2024216836A1 (en) | 2024-10-24 |
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