CN110304270B - Omnibearing launch control method and device for carrier rocket and computer equipment - Google Patents
Omnibearing launch control method and device for carrier rocket and computer equipment Download PDFInfo
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Abstract
The invention is suitable for the technical field of rocket control, and provides an omnibearing launch control method and device for a carrier rocket, computer equipment and a storage medium, wherein the method comprises the following steps: acquiring an initial attitude angle and calculating an initial quaternion, wherein the initial attitude angle comprises an initial rolling angle; acquiring a real-time angle increment of a current control period, and calculating a real-time quaternion; calculating a standard quaternion according to the initial rolling angle; and calculating attitude angle deviation based on the real-time quaternion and the standard quaternion, and acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism. The initial quaternion is calculated through the initial attitude angle, the standard quaternion and the real-time quaternion are obtained, the attitude angle deviation is obtained through calculation, the target attitude control signal for controlling the rocket is obtained after calculation and calculation, the omnibearing launching control of the rocket is realized without rotating the indexing mechanism, the cost and the time for rotating the rocket for alignment are saved by the indexing mechanism, and the quick response capability and the automation level of the rocket are improved.
Description
Technical Field
The invention belongs to the technical field of rocket control, and particularly relates to an omnibearing launch control method and device for a carrier rocket, computer equipment and a storage medium.
Background
At present, most of the traditional liquid rocket or solid rocket navigation adopts platform inertial navigation (inertial navigation), and with the technical progress, the mode of adopting strapdown inertial navigation gradually becomes a main navigation control mode of the rocket, particularly the solid rocket. The inertial navigation system is an autonomous navigation system independent of any external information, and is mainly divided into a platform type inertial navigation system and a strapdown type inertial navigation system. A Strapdown Inertial Navigation System (SINS) is developed on the basis of a platform type inertial navigation system, is a frameless system and consists of a rate gyro, a linear accelerometer and a computer. The gyroscope and the accelerometer are directly connected to the carrier and are used for measuring angular motion information and linear motion information of the carrier respectively, and the computer calculates the attitude, the speed and the position of the carrier according to the measurement information. The strapdown inertial navigation system saves a complex electromechanical platform, and has the advantages of simple structure, small volume, light weight, low cost, simple maintenance and high reliability.
For a solid launch vehicle, strapdown inertial navigation is usually used for navigation control, but before the launch of the rocket, azimuth alignment (manual or automatic) is required, so that the actual shooting direction of the rocket in the vertical state is coincident with the theoretical shooting direction or controlled at a smaller angle, namely, the launching azimuth angle of the rocket is turned to the theoretical shooting direction. For a rocket with tens of tons or even hundreds of tons, the manual or automatic rotation adjustment process takes a certain time, and an indexing mechanism is further added, which increases the complexity and equipment cost of the launching system, and affects the response capability and automation level of the system.
Disclosure of Invention
The embodiment of the invention provides an omnibearing launch control method for a carrier rocket, and aims to solve the problems of long time and high cost caused by the fact that a rotary indexing mechanism is required to carry out azimuth alignment before launching the existing rocket.
The embodiment of the invention is realized in such a way, and provides an omnibearing launch control method for a carrier rocket, which comprises the following steps:
acquiring an initial attitude angle and calculating an initial quaternion, wherein the initial attitude angle comprises an initial rolling angle;
acquiring a real-time angle increment of a current control period, and calculating a real-time quaternion;
calculating a standard quaternion according to the initial rolling angle;
and calculating attitude angle deviation based on the real-time quaternion and the standard quaternion, and acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism.
Still further, after the step of calculating an attitude angle deviation based on the real-time quaternion and the standard quaternion, and obtaining a guidance control signal according to the attitude angle deviation and outputting the guidance control signal to an actuator, the method further comprises the steps of:
and if the time sum of the control period is smaller than a preset time threshold, iteratively calculating the real-time quaternion and the standard quaternion of the next control period and recalculating the attitude angle deviation.
Further, the step of obtaining the initial attitude angle and calculating the initial quaternion specifically includes:
acquiring a theoretical directive angle and an initial actual azimuth angle;
obtaining the initial rolling angle through the theoretical shooting angle and the initial actual azimuth angle;
and calculating to obtain the initial quaternion based on the initial attitude angle.
Further, the step of obtaining the real-time angle increment of the current control period and calculating the real-time quaternion specifically includes:
acquiring real-time angle increment acquired by an inertial navigation system and a real-time quaternion of a previous control period;
and calculating to obtain the real-time quaternion of the current control period based on the real-time angle increment and the real-time quaternion of the previous control period.
Further, the step of calculating a standard quaternion according to the initial roll angle specifically includes:
acquiring a standard pitch angle and a standard yaw angle of a current control period;
setting a standard roll angle to the initial roll angle;
and calculating to obtain a standard quaternion of the current control period according to the standard rolling angle, the standard pitch angle and the standard yaw angle.
Further, the step of calculating an attitude angle deviation based on the real-time quaternion and the standard quaternion, and obtaining a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuator specifically includes:
calculating to obtain an attitude angle deviation based on the real-time quaternion and the standard quaternion of the current control period;
inputting the attitude angle deviation into an attitude control system for resolving to obtain the target attitude control signal;
and outputting the target attitude control signal to an executing mechanism for execution.
Further, before the step of outputting the target attitude control signal to an actuator for execution, the method further comprises:
and superposing the omnidirectionally transformed transverse normal guidance control signal to the target attitude control signal.
An embodiment of the present invention further provides a rocket launching control device, including:
the system comprises a first acquisition module, a second acquisition module and a fourth acquisition module, wherein the first acquisition module is used for acquiring an initial attitude angle and calculating an initial quaternion, and the initial attitude angle comprises an initial rolling angle;
the first calculation module is used for acquiring the real-time angle increment of the current control period and calculating a real-time quaternion;
the second calculation module is used for calculating a standard quaternion according to the initial rolling angle;
and the second acquisition module is used for calculating the attitude angle deviation based on the real-time quaternion and the standard quaternion, acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism.
The embodiment of the invention also provides computer equipment which comprises a memory and a processor, wherein the memory stores a computer program, and the processor realizes the steps of the omnibearing launch control method for the carrier rocket when executing the computer program.
An embodiment of the present invention further provides a computer-readable storage medium, where a computer program is stored on the computer-readable storage medium, and when the computer program is executed by a processor, the computer program implements the steps of the omni-directional launch control method for a launch vehicle.
In the embodiment of the invention, an initial attitude angle is obtained and an initial quaternion is calculated, wherein the initial attitude angle comprises an initial rolling angle; acquiring a real-time angle increment of a current control period, and calculating a real-time quaternion; calculating a standard quaternion according to the initial rolling angle; and calculating attitude angle deviation based on the real-time quaternion and the standard quaternion, and acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism. Because the initial quaternion is calculated through any initial attitude angle obtained by placing the rocket at any position of the launching system, and then the standard quaternion and the real-time quaternion are obtained, so that the attitude angle deviation is calculated and the target attitude control signal for controlling the rocket is obtained after calculation, the omnibearing launching control of the rocket is realized without rotating the indexing mechanism, the cost and the time for rotating the rocket for alignment are saved by the indexing mechanism, and the quick response capability and the automation level of the rocket are improved.
Drawings
FIG. 1 is a schematic view of the orientation of each axis of an rocket coordinate system O1X1Y1Z1 of a rocket provided by an embodiment of the invention and erected on a launching pad;
fig. 2 is a directional diagram of an rocket coordinate system O1X1Y1Z1 and a rocket navigation coordinate system oyx Z, in which a rocket is erected on a launching pad, under an omnidirectional launching condition provided by an embodiment of the invention;
FIG. 3 is a flow chart of an omni-directional launch control method for a launch vehicle according to an embodiment of the present invention;
FIG. 4 is a schematic structural diagram of a rocket launch control device according to an embodiment of the present invention;
fig. 5 is a schematic structural diagram of an electronic device according to an embodiment of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is described in further detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not intended to limit the invention.
And calculating an initial quaternion through any initial attitude angle obtained by placing the rocket at any position of the launching system, and further obtaining a standard quaternion and a real-time quaternion, thereby obtaining the attitude angle deviation and resolving to obtain a target attitude control signal of the rocket, and realizing the omnibearing launching control of the rocket.
Referring to fig. 1, a schematic view of pointing directions of axes of an rocket coordinate system O1X1Y1Z1 in which a rocket stands on a launching platform is shown, where O1 is an origin of coordinates, inertial navigation employs strapdown inertial navigation, and a three-axis direction of the inertial navigation coincides with the rocket coordinate system.
Referring to fig. 2, the directional diagram is an arrow coordinate system O1X1Y1Z1 and a rocket navigation coordinate system oyx when the rocket is erected on the launching pad under the condition of omnibearing launching. The rocket navigation coordinate system, oyx, is a launch inertial coordinate system, which is defined as: the origin of coordinates O coincides with O1, the OX axis points to the theoretical emission point A0 on the horizontal plane, the OY axis is the reverse direction of the perpendicular line of the emission point, the OZ axis is determined by the right-hand rule, and the solidification is carried out at the moment of the emission point.
the-Y1 direction of inertial navigation is the actual azimuth A1 of the rocket; generally, a launching system (a launching vehicle and a launching tower) erects the rocket from the horizontal, and after the rocket is erected in place, the rocket is subjected to related tests, and the A1 can be obtained by optical aiming in the + Y1 direction, and the angle can also be obtained in a self-aligning (self-north-seeking) mode. The orientation of A1 is uncertain when the missile is upright and in most cases it is not coincident with the theoretical heading A0. Conventional rockets generally rotate the rocket through an indexing mechanism, so that the A1 is overlapped with the A0 or has small deviation, and the method is equivalent to the I quadrant pointing theory direction of the rocket. If the rocket does not rotate around the X1 axis before launching, the included angle gamma 0 between A1 and A0 is A0-A1 which can reach 180 degrees at most. After the included angle between the current azimuth angle of the rocket and the theoretical direction is obtained, the included angle is sent to a control computer through a ground measurement and launch control system, and the control computer executes the omnibearing launch control method for the carrier rocket, so that omnibearing launch control is realized without rotating an indexing mechanism.
Example one
As shown in fig. 3, fig. 3 is an omnidirectional launch control method for a launch vehicle according to an embodiment of the present invention, including the following steps:
101. an initial attitude angle is obtained and an initial quaternion is calculated, the initial attitude angle comprising an initial roll angle.
In this embodiment, the rocket is placed at any position of the launching system to obtain any initial attitude angle, including an initial roll angle, an initial yaw angle, and an initial pitch angle. The initial rolling angle is an included angle between an actual azimuth angle of the rocket and a theoretical direction, and an initial yaw angle and an initial pitch angle can be obtained through initial leveling. And then obtaining the initial quaternion through an initial quaternion calculation formula based on the initial attitude angle.
102. And acquiring the real-time angle increment of the current control period, and calculating a real-time quaternion.
Wherein, the control period is the time interval of the rocket control computer sending the control signal. And finally, calculating to obtain the real-time quaternion of the current control period based on the real-time quaternion of the previous control period and the real-time angle increment of the current control period.
103. And calculating a standard quaternion according to the initial rolling angle.
In this embodiment, the standard quaternion is a theoretical value of a standard flight trajectory of the rocket, and can be obtained by calculating a standard roll angle, a standard yaw angle and a standard pitch angle. In the process of realizing the omnibearing launching control without rotating the indexing mechanism, the standard rolling angle can be set as the initial rolling angle, and the standard yaw angle and the standard pitch angle can be obtained by the control computer according to theoretical values calculated by the standard rocket flight path and stored in the control computer in a form of a table, and then the theoretical values are obtained by looking up the table according to needs.
104. And calculating attitude angle deviation based on the real-time quaternion and the standard quaternion, and acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism.
In this embodiment, the real-time quaternion of the current control period may be obtained in step 102, the standard quaternion may be obtained in step 103, the attitude angle deviation of the current control period is calculated by the attitude angle deviation calculation formula based on the real-time quaternion and the standard quaternion, and then the attitude angle deviation is input to the attitude control system for calculation, so as to obtain the target attitude control signal and output to the actuator for attitude control.
In the embodiment of the invention, an initial attitude angle is obtained and an initial quaternion is calculated, wherein the initial attitude angle comprises an initial rolling angle; acquiring a real-time angle increment of a current control period, and calculating a real-time quaternion; calculating a standard quaternion according to the initial rolling angle; and calculating attitude angle deviation based on the real-time quaternion and the standard quaternion, and acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism. The initial quaternion is calculated through any initial attitude angle obtained by placing the rocket at any position of the launching system, and then the standard quaternion and the real-time quaternion are obtained, so that the attitude angle deviation is calculated and the target attitude control signal for controlling the rocket is obtained after calculation, the omnibearing launching control of the rocket is realized without rotating the indexing mechanism, the cost is saved by the aid of the indexing mechanism, the time for rotating the rocket for alignment is saved, and the quick response capability and the automation level of the rocket are improved.
Example two
Further, after the step of calculating the attitude angle deviation based on the real-time quaternion and the standard quaternion, and obtaining the guidance control signal according to the attitude angle deviation and outputting the guidance control signal to the actuator, the method further comprises the steps of:
and if the time sum of the control period is smaller than a preset time threshold, iteratively calculating the real-time quaternion and the standard quaternion of the next control period and recalculating the attitude angle deviation.
In this embodiment, before the rocket enters the standard flight trajectory, the attitude needs to be adjusted, and the adjustment time may be a time threshold preset according to the actual situation, such as 40 s; within the preset time threshold, the control computer sends a control signal at regular intervals (i.e. a control period, such as 10ms) to gradually adjust the flight attitude so as to finally satisfy the condition of entering the orbit, and therefore, it is necessary to iteratively calculate the real-time quaternion and the standard quaternion and recalculate the attitude angle deviation according to the above embodiment, and obtain the guidance control signal of each control period, so as to continuously control the rocket within the preset time threshold.
EXAMPLE III
Further, the step of obtaining the initial attitude angle and calculating the initial quaternion specifically includes:
acquiring a theoretical directive angle and an initial actual azimuth angle;
obtaining the initial rolling angle through the theoretical shooting angle and the initial actual azimuth angle;
and calculating to obtain the initial quaternion based on the initial attitude angle.
In this embodiment, as shown in fig. 1 and fig. 2, the theoretical direction of the rocket is a0, which is calculated by ballistic calculation software according to a standard flight trajectory; the initial actual azimuth angle a1 is obtained by optical aiming in the + Y1 direction when the rocket is placed at any position of the launching system, and can also be obtained by self-alignment (self-north-seeking).
The initial attitude angle comprises an initial rolling angle, an initial yaw angle and an initial pitch angle. Wherein the initial rolling angle gamma 0 is an included angle between the actual azimuth angle of the rocket and the theoretical direction, namely gamma 0 is A0-A1; the initial yaw angle can be obtained through initial leveling
Initial pitch angle psi 0, wherein initial yaw angle
Is + Y1 from the horizontal plane, and the initial pitch angle ψ 0 is + Z1 from the horizontal plane. The initial quaternion Q0 may then be derived based on the initial attitude angle by the following initial quaternion calculation formula:
example four
Further, the step of obtaining the real-time angle increment of the current control period and calculating the real-time quaternion specifically includes:
acquiring real-time angle increment acquired by an inertial navigation system and a real-time quaternion of a previous control period;
and calculating to obtain the real-time quaternion of the current control period based on the real-time angle increment and the real-time quaternion of the previous control period.
In this embodiment, the three-axis sensor of the inertial navigation system collects real-time angular increments Δ θ of the current control period of three channels (yaw, pitch, roll) in real timeX1、ΔθY1、ΔθZ1Then the real-time quaternion and the real-time quaternion of the previous control period are calculated by the following calculation formula:
the subscript t0 represents a previous control period, the subscript h represents a current control period, and the subscript t0+ h represents that the latest real-time quaternion Qj is obtained by using inertial navigation output in the previous control period and the current control period.
EXAMPLE five
Further, the step of calculating a standard quaternion according to the initial roll angle specifically includes:
acquiring a standard pitch angle and a standard yaw angle of a current control period;
setting a standard roll angle to the initial roll angle;
and calculating to obtain a standard quaternion of the current control period according to the standard rolling angle, the standard pitch angle and the standard yaw angle.
In the present embodiment, the standard quaternion QcxIs a theoretical value of a rocket standard flight path and can pass through a standard rolling angle gamma cx and a standard yaw angle
And calculating the standard pitch angle psi cx. In the attitude adjustment process for realizing the omnidirectional emission control without rotating the indexing mechanism, the standard roll angle γ cx may be set to the initial roll angle γ 0, that is, γ cx ═ γ 0, and the standard yaw angle
The standard pitch angle ψ cx can be obtained by looking up a table from a table stored in the control computer according to the time of flight or the altitude of flight, and then the standard quaternion Q of the current control period described above is obtained by calculation according to the following formulacx：
EXAMPLE six
Further, the step of calculating an attitude angle deviation based on the real-time quaternion and the standard quaternion, obtaining a target attitude control signal according to the attitude angle deviation, and outputting the target attitude control signal to an execution mechanism specifically includes:
calculating to obtain an attitude angle deviation based on the real-time quaternion and the standard quaternion of the current control period;
inputting the attitude angle deviation into an attitude control system for resolving to obtain the target attitude control signal;
and outputting the target attitude control signal to an executing mechanism for execution.
In this embodiment, the real-time quaternion of the current control period is recorded
Then the standard quaternion Q is usedcxAnd real-time quaternion QjThe attitude angle deviation can be calculated as follows:
and further obtaining attitude angle deviation of three channels:
wherein, Δ ZTX1、ΔZTY1、ΔZTZ1The attitude angle deviation of rolling, yaw and pitch are respectively.
Then the attitude angle deviation delta ZT of the three channelsX1、ΔZTY1、ΔZTZ1Inputting the target attitude control signal into an attitude control system for resolving to obtain a target attitude control signal of the arrow system, and outputting the target attitude control signal to an execution mechanism for executionThereby controlling the arrow body. The attitude control system is the same as the attitude control system which does not use the omnibearing transmission control, and finally the attitude control is realized by controlling the rotation of the arrow body.
EXAMPLE seven
Further, before the step of outputting the target attitude control signal to an actuator for execution, the method further includes:
and superposing the omnidirectionally transformed transverse normal guidance control signal to the target attitude control signal.
In this embodiment, if the rocket adds the transverse normal steering control signal during the omnidirectional launching control, the transverse normal steering control signal needs to be transformed in an omnidirectional manner, that is, the transverse normal steering control signal is transformed from the navigation coordinate system to the rocket body coordinate system O1X1Y1Z1, and then is superimposed on the target attitude control signal, and then the superimposed target attitude control signal is output, where the transformation formula is as follows:
UY and UZ are transverse normal guidance control signals obtained by calculation under a navigation coordinate system and are used for controlling the position and the speed of the rocket, and if omnibearing launching is not carried out, the signals can be directly added to a yaw channel and a pitching channel.
It should be noted that the preset time threshold is the time (e.g. 40s) for performing attitude adjustment on the rocket before entering the standard flight trajectory, and within the preset time threshold, the control computer sends a control signal at regular intervals (i.e. a control period, e.g. 10ms) to perform continuous control on the rocket, i.e. to perform gradual adjustment on the flight attitude until the rocket enters the standard flight trajectory. If the time sum of the continuous control periods is greater than or equal to a preset time threshold value, which indicates that the posture of the rocket is close to the orbit entering condition, resetting the standard rolling angle gamma cx according to the following formula until the gamma cx is equal to 0:
γcx＝γ0+(0-γ0)*(t-T40)/15
wherein T is zero time for starting timing of rocket ignition, and T is40The time for the posture adjustment is 40s in the present embodiment, and may be set to other times according to actual situations. And when the gamma cx is 0, namely the included angle between the current azimuth angle of the rocket and the theoretical direction is 0, the current azimuth angle of the rocket is superposed with the theoretical direction, the rocket enters a standard flight orbit, and the omnibearing launching control is finished.
The above optional embodiment is a supplementary embodiment of the omnibearing launch control method for the launch vehicle in fig. 3, and the method in the above optional embodiment can achieve corresponding beneficial effects, and is not described here again to avoid repetition.
Example eight
Referring to fig. 4, fig. 4 is a schematic structural diagram of a rocket launch control device according to an embodiment of the present invention, and as shown in fig. 4, the device 200 includes:
a first obtaining module 201, configured to obtain an initial attitude angle and calculate an initial quaternion, where the initial attitude angle includes an initial roll angle;
a first calculating module 202, configured to obtain a real-time angle increment of a current control period, and calculate a real-time quaternion;
a second calculating module 203, configured to calculate a standard quaternion according to the initial roll angle;
and a second obtaining module 204, configured to calculate an attitude angle deviation based on the real-time quaternion and the standard quaternion, and obtain a target attitude control signal according to the attitude angle deviation and output the target attitude control signal to an execution mechanism.
Example nine
Referring to fig. 5, fig. 5 is a schematic structural diagram of an electronic device according to an embodiment of the present invention, as shown in fig. 5, including: a memory 302, a processor 301, and a computer program stored on the memory 302 and executable on the processor 301, wherein:
the processor 301 is configured to call the computer program stored in the memory 302, and execute the following steps:
acquiring an initial attitude angle and calculating an initial quaternion, wherein the initial attitude angle comprises an initial rolling angle;
acquiring a real-time angle increment of a current control period, and calculating a real-time quaternion;
calculating a standard quaternion according to the initial rolling angle;
and calculating attitude angle deviation based on the real-time quaternion and the standard quaternion, and acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism.
Optionally, the processor 301 further performs the following steps after the step of calculating the attitude angle deviation based on the real-time quaternion and the standard quaternion, and obtaining the guidance control signal according to the attitude angle deviation and outputting the guidance control signal to the actuator:
and if the time sum of the control period is smaller than a preset time threshold, iteratively calculating the real-time quaternion and the standard quaternion of the next control period and recalculating the attitude angle deviation.
Optionally, the step of executing by the processor 301 to obtain the initial attitude angle and calculate the initial quaternion specifically includes:
acquiring a theoretical directive angle and an initial actual azimuth angle;
obtaining the initial rolling angle through the theoretical shooting angle and the initial actual azimuth angle;
and calculating to obtain the initial quaternion based on the initial attitude angle.
Optionally, the step of executing, by the processor 301, the step of obtaining the real-time angular increment of the current control period and calculating the real-time quaternion specifically includes:
acquiring real-time angle increment acquired by inertial navigation equipment and a real-time quaternion of a previous control period;
and calculating to obtain the real-time quaternion of the current control period based on the real-time angle increment and the real-time quaternion of the previous control period.
Optionally, the step of executing, by the processor 301, the standard quaternion according to the initial roll angle specifically includes:
acquiring a standard pitch angle and a standard yaw angle of a current control period;
setting a standard roll angle to the initial roll angle;
and calculating to obtain a standard quaternion of the current control period according to the standard rolling angle, the standard pitch angle and the standard yaw angle.
Optionally, the step of executing, by the processor 301, the calculation of the attitude angle deviation based on the real-time quaternion and the standard quaternion, and obtaining the target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to the execution mechanism specifically includes:
calculating to obtain an attitude angle deviation based on the real-time quaternion and the standard quaternion of the current control period;
inputting the attitude angle deviation into an attitude control system for resolving to obtain the target attitude control signal;
and outputting the target attitude control signal to an executing mechanism for execution.
Optionally, the processor 301 further includes, before the step of outputting the target attitude control signal to an actuator for execution:
and superposing the transverse and normal guidance control signal to the target attitude control signal.
The Processor 301 may be a Digital Signal Processor (DSP), an Application Specific Integrated Circuit (ASIC), a Field Programmable Gate Array (FPGA) or other Programmable logic device, discrete Gate or transistor logic device, discrete hardware components, and so on.
It should be noted that, since the processor 301 executes the computer program stored in the memory 302 to implement the steps of the omnidirectional launch control method for a launch vehicle, all embodiments of the omnidirectional launch control method for a launch vehicle are applicable to the electronic device and can achieve the same or similar advantages.
In addition, the embodiment of the present invention further provides a computer-readable storage medium 302, where the computer-readable storage medium 302 stores a computer program, and the computer program, when executed by a processor, implements the steps of the above-mentioned all-directional launch control method for a launch vehicle.
That is, in the embodiment of the present invention, when the computer program of the computer-readable storage medium is executed by the processor, the steps of the above-described method for controlling omnidirectional launching of a launch vehicle can be implemented, the omnidirectional launching control of a rocket can be implemented without rotating the indexing mechanism, the cost and the time for rotating the rocket for alignment can be saved by omitting the indexing mechanism, and the fast response capability and the automation level of the rocket can be improved.
Illustratively, the computer program of the computer-readable storage medium comprises computer program code, which may be in the form of source code, object code, an executable file or some intermediate form, and the like. The computer-readable medium may include: any entity or device capable of carrying the computer program code, recording medium, usb disk, removable hard disk, magnetic disk, optical disk, computer Memory, Read-Only Memory (ROM), Random Access Memory (RAM), electrical carrier wave signals, telecommunications signals, software distribution medium, and the like.
It should be noted that all the embodiments of the omnidirectional launch control method for a launch vehicle described above are applicable to the computer-readable storage medium, and can achieve the same or similar advantageous effects, since the computer program of the computer-readable storage medium implements the steps of the omnidirectional launch control method for a launch vehicle described above when executed by the processor.
It will be understood by those skilled in the art that all or part of the processes of the methods of the embodiments described above can be implemented by a computer program, which can be stored in a computer-readable storage medium, and when executed, can include the processes of the embodiments of the methods described above.
It should be noted that, for simplicity of description, the above-mentioned method embodiments are described as a series of acts or combination of acts, but those skilled in the art will recognize that the present application is not limited by the order of acts described, as some steps may occur in other orders or concurrently depending on the application. Further, those skilled in the art should also appreciate that the embodiments described in the specification are exemplary embodiments and that the acts and modules referred to are not necessarily required in this application.
In the foregoing embodiments, the descriptions of the respective embodiments have respective emphasis, and for parts that are not described in detail in a certain embodiment, reference may be made to related descriptions of other embodiments.
In the embodiments provided in the present application, it should be understood that the disclosed apparatus may be implemented in other manners. For example, the above-described embodiments of the apparatus are merely illustrative, and for example, the division of the units is only one type of division of logical functions, and there may be other divisions when actually implementing, for example, a plurality of units or components may be combined or may be integrated into another system, or some features may be omitted, or not implemented. In addition, the shown or discussed mutual coupling or direct coupling or communication connection may be an indirect coupling or communication connection of some interfaces, devices or units, and may be an electric or other form.
The units described as separate parts may or may not be physically separate, and parts displayed 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 can be selected according to actual needs to achieve the purpose of the solution of the embodiment.
In addition, functional units in the embodiments of the present application may be integrated into one processing unit, or each unit may exist alone physically, or two or more units are integrated into one unit. The integrated unit may be implemented in the form of hardware, or may be implemented in the form of a software program module.
The integrated units, if implemented in the form of software program modules and sold or used as stand-alone products, may be stored in a computer readable memory. Based on such understanding, the technical solution of the present application may be substantially implemented or a part of or all or part of the technical solution contributing to the prior art may be embodied in the form of a software product stored in a memory, and including several instructions for causing a computer device (which may be a personal computer, a server, or a network device) to execute all or part of the steps of the method described in the embodiments of the present application. And the aforementioned memory comprises: a U-disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a removable hard disk, a magnetic or optical disk, and other various media capable of storing program codes.
Those skilled in the art will appreciate that all or part of the steps in the methods of the above embodiments may be implemented by associated hardware instructed by a program, which may be stored in a computer-readable memory, which may include: flash Memory disks, Read-Only memories (ROMs), Random Access Memories (RAMs), magnetic or optical disks, and the like.
The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents and improvements made within the spirit and principle of the present invention are intended to be included within the scope of the present invention.
Claims (10)
1. An omnibearing launch control method for a carrier rocket is characterized by comprising the following steps:
acquiring an initial attitude angle and calculating an initial quaternion, wherein the initial attitude angle comprises an initial rolling angle;
acquiring a real-time angle increment of a current control period, and calculating a real-time quaternion;
calculating a standard quaternion according to the initial rolling angle;
and calculating attitude angle deviation based on the real-time quaternion and the standard quaternion, and acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism.
2. The method as claimed in claim 1, wherein after the step of calculating an attitude angle deviation based on the real-time quaternion and the standard quaternion, and obtaining a pilot control signal according to the attitude angle deviation and outputting the pilot control signal to an actuator, the method further comprises the steps of:
and if the time sum of the control period is smaller than a preset time threshold, iteratively calculating the real-time quaternion and the standard quaternion of the next control period and recalculating the attitude angle deviation.
3. The method of claim 2, wherein the step of obtaining the initial attitude angle and calculating the initial quaternion specifically comprises:
acquiring a theoretical directive angle and an initial actual azimuth angle;
obtaining the initial rolling angle through the theoretical shooting angle and the initial actual azimuth angle;
and calculating to obtain the initial quaternion based on the initial attitude angle.
4. The method of claim 3, wherein the step of obtaining the real-time angular increment of the current control period and calculating the real-time quaternion specifically comprises:
acquiring real-time angle increment acquired by an inertial navigation system and a real-time quaternion of a previous control period;
and calculating to obtain the real-time quaternion of the current control period based on the real-time angle increment and the real-time quaternion of the previous control period.
5. The method according to claim 4, wherein the step of calculating a standard quaternion from the initial roll angle specifically comprises:
acquiring a standard pitch angle and a standard yaw angle of a current control period;
setting a standard roll angle to the initial roll angle;
and calculating to obtain a standard quaternion of the current control period according to the standard rolling angle, the standard pitch angle and the standard yaw angle.
6. The method of claim 5, wherein the step of calculating an attitude angle deviation based on the real-time quaternion and the standard quaternion, and obtaining a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuator specifically comprises:
calculating to obtain an attitude angle deviation based on the real-time quaternion and the standard quaternion of the current control period;
inputting the attitude angle deviation into an attitude control system for resolving to obtain the target attitude control signal;
and outputting the target attitude control signal to an executing mechanism for execution.
7. The method of claim 6, further comprising, prior to the step of outputting the target attitude control signal to an actuator for execution:
and superposing the omnidirectionally transformed transverse normal guidance control signal to the target attitude control signal.
8. A rocket launch control device, comprising:
the system comprises a first acquisition module, a second acquisition module and a fourth acquisition module, wherein the first acquisition module is used for acquiring an initial attitude angle and calculating an initial quaternion, and the initial attitude angle comprises an initial rolling angle;
the first calculation module is used for acquiring the real-time angle increment of the current control period and calculating a real-time quaternion;
the second calculation module is used for calculating a standard quaternion according to the initial rolling angle;
and the second acquisition module is used for calculating the attitude angle deviation based on the real-time quaternion and the standard quaternion, acquiring a target attitude control signal according to the attitude angle deviation and outputting the target attitude control signal to an actuating mechanism.
9. A computer arrangement, characterized by comprising a memory in which a computer program is stored and a processor which, when executing said computer program, carries out the steps of the method for controlling an omnidirectional launch for a launch vehicle according to any one of claims 1 to 7.
10. A computer-readable storage medium, characterized in that a computer program is stored thereon, which, when being executed by a processor, carries out the steps of the method for controlling an omnidirectional launch for a launch vehicle according to any one of claims 1 to 7.
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| CN117553631B (en) * | 2023-02-22 | 2024-04-09 | 东方空间技术(山东)有限公司 | Attitude control method, device and equipment for boosting flight section of carrier rocket |
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