CN111176310B - Test method, device and system for carrier rocket attitude control system - Google Patents
Test method, device and system for carrier rocket attitude control system Download PDFInfo
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- CN111176310B CN111176310B CN201911414863.2A CN201911414863A CN111176310B CN 111176310 B CN111176310 B CN 111176310B CN 201911414863 A CN201911414863 A CN 201911414863A CN 111176310 B CN111176310 B CN 111176310B
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Abstract
The invention discloses a test method, a device and a system of a carrier rocket attitude control system, wherein the carrier rocket attitude control system is used for carrying out attitude control on a preset carrier rocket model, and the method comprises the following steps: inputting fault parameters in a carrier rocket attitude control system; acquiring first flight state data of the carrier rocket after fault injection according to the input fault parameters; judging whether the fault causes the operation error of the carrier rocket or not according to the first flight state data; and when the fault causes the carrier rocket to have an error in operation, recording fault parameters and first flight state data. By implementing the method, the problem that each system of the carrier rocket is defaulted to normally operate and the fault condition of the carrier rocket is ignored in the conventional test method of the attitude control system of the carrier rocket is solved by combining the fault type in the attitude control system and the operation data of the carrier rocket, the flight safety performance of the carrier rocket is ensured, and the test method of the carrier rocket is more comprehensive and more accurate.
Description
Technical Field
The invention relates to the field of aerospace equipment, in particular to a method, a device and a system for testing a carrier rocket attitude control system.
Background
The carrier rocket control system is a core system for ensuring the normal flight of the carrier rocket, the attitude control system is an important component of the carrier rocket control system, the working condition of the attitude control system can directly influence the success and failure of the launch task of the carrier rocket, and the flight quality of the carrier rocket can be directly influenced. And the safety of people's lives and properties and the precision of load orbit entering along the flight trajectory of the carrier rocket are closely related.
The existing test method of the carrier rocket attitude control system mainly depends on frequency domain analysis, mathematical simulation and semi-physical simulation, and the test methods are all established on a mathematical model reflecting the kinematics and dynamics environment of the carrier rocket. Based on the mathematical models, deviation can be added for certain parameters, and the attitude control system can be examined by matching with a simulated flight test. However, the existing test method for the attitude control system of the carrier rocket defaults that each system of the carrier rocket normally operates, and influences caused by faults of the carrier rocket are ignored.
Disclosure of Invention
Therefore, the invention aims to overcome the defects that each system of a default carrier rocket normally operates and the fault condition of the carrier rocket is ignored in the existing carrier rocket test technology, and provides a test method, a device and a system of a carrier rocket attitude control system.
According to a first aspect, the embodiment of the invention discloses a test method of a carrier rocket attitude control system, wherein the carrier rocket attitude control system is used for carrying out attitude control on a preset carrier rocket model and comprises the following steps: inputting fault parameters in a carrier rocket attitude control system; acquiring first flight state data of the carrier rocket after fault injection according to input fault parameters; judging whether the fault causes the carrier rocket to operate mistakenly or not according to the first flight state data; and when the fault causes the carrier rocket to have an error in operation, recording the fault parameter and the first flight state data.
With reference to the first aspect, in a first embodiment of the first aspect, the inputting the fault parameter in the posture control system of the launch vehicle includes: inputting a fault parameter at a flight phase of the launch vehicle, or inputting fault parameters at a flight phase of the launch vehicle, or inputting a fault parameter at flight phases of the launch vehicle, or inputting fault parameters at flight phases of the launch vehicle.
With reference to the first aspect, in a second embodiment of the first aspect, the method for testing a launch vehicle attitude control system further includes: when the fault causes that the carrier rocket does not have an error in operation, increasing the fault parameter; acquiring second flight state data of the carrier rocket after fault injection according to the increased fault parameters; judging whether the fault causes the carrier rocket to operate mistakenly or not according to the second flight state data; if the fault causes the carrier rocket to operate mistakenly, recording the increased fault parameters and the second flight state data; and if the fault causes the carrier rocket to operate without errors, continuously increasing the fault parameters until the fault causes the carrier rocket to operate with errors.
With reference to the first aspect, in a third implementation of the first aspect, the flight state data includes one or more of flight speed, or flight angle, or altitude, or flight attitude.
With reference to the first aspect, in a fourth implementation of the first aspect, the input fault parameter is implemented by changing a control parameter in an attitude control system.
According to a second aspect, an embodiment of the present invention discloses a test apparatus for a posture control system of a launch vehicle, where the posture control system of the launch vehicle is used to perform posture control on a preset launch vehicle model, and includes: the input module is used for inputting fault parameters in the carrier rocket attitude control system; the first acquisition module is used for acquiring first flight state data of the carrier rocket after fault injection according to input fault parameters; the first judging module is used for judging whether the fault causes the carrier rocket to operate mistakenly or not according to the first flight state data; and the first recording module is used for recording the fault parameters and the first flight state data when the fault causes the carrier rocket to operate in error.
With reference to the second aspect, in a first implementation manner of the second aspect, the input module specifically includes: a first input submodule for inputting a fault parameter during a flight phase of said launch vehicle; the first input sub-module is used for inputting a plurality of fault parameters in one flight stage of the carrier rocket, the first input sub-module is used for inputting one fault parameter in a plurality of flight stages of the carrier rocket, and the first input sub-module is used for inputting a plurality of fault parameters in a plurality of flight stages of the carrier rocket.
With reference to the second aspect, in a second embodiment of the second aspect, the apparatus further comprises: the increasing module is used for increasing the fault parameter when the fault causes that the carrier rocket does not have fault in operation; the second acquisition module is used for acquiring second flight state data of the carrier rocket after fault injection according to the increased fault parameters; the second judgment module is used for judging whether the fault causes the carrier rocket to operate mistakenly or not according to the second flight state data; the second recording module is used for recording the increased fault parameters and the second flight state data if the fault causes the carrier rocket to have errors in operation; the third recording module is used for continuously increasing the fault parameters until the fault causes the carrier rocket to operate mistakenly if the fault causes the carrier rocket to operate not mistakenly; and recording fault parameters causing the operation error of the carrier rocket and corresponding flight state data thereof. According to a third aspect, an embodiment of the present invention discloses a test system for a posture control system of a launch vehicle, including: at least one control device configured to perform the steps of the method for testing a posture control system of a launch vehicle as set forth in the first aspect or any one of the embodiments of the first aspect, acquire a flight state of the launch vehicle according to a type of an injected fault, and determine whether the fault causes the launch vehicle to fail in flight.
According to a fourth aspect, an embodiment of the present invention discloses a computer-readable storage medium, on which a computer program is stored, which, when executed by a processor, implements the steps of the method for testing a launch vehicle attitude control system as described in the first aspect or any one of the embodiments of the first aspect.
The technical scheme of the invention has the following advantages:
1. the invention discloses a test method, a device and a system of a carrier rocket attitude control system, wherein the carrier rocket attitude control system is used for carrying out attitude control on a preset carrier rocket model, and the method comprises the following steps: inputting fault parameters in a carrier rocket attitude control system; acquiring first flight state data of the carrier rocket after fault injection according to the input fault parameters; judging whether the fault causes the operation error of the carrier rocket or not according to the first flight state data; and when the fault causes the carrier rocket to have an error in operation, recording fault parameters and first flight state data. By implementing the method, the problem that each system of the carrier rocket is defaulted to normally operate and the fault condition of the carrier rocket is ignored in the conventional test method of the attitude control system of the carrier rocket is solved by combining the fault type in the attitude control system and the operation data of the carrier rocket, the flight safety performance of the carrier rocket is ensured, and the test method of the carrier rocket is more comprehensive and more accurate.
2. The invention discloses a test method of a carrier rocket attitude control system, which can simulate various faults by inputting various fault parameters and test the flight condition of the carrier rocket, can inject the same fault in one flight stage or a plurality of flight stages of the carrier rocket, and can inject various faults in one flight stage or a plurality of flight stages, and reasonably designs and sets the type and the number of fault injection in the test of the carrier rocket attitude control system, so that the fault injection can correctly reflect the possible problems of the carrier rocket attitude control system in an actual flight test, comprehensively simulates the fault condition of the carrier rocket attitude control system, and the test result of the carrier rocket is more accurate.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.
Fig. 1 is a structural block diagram of a mathematical simulation environment in a test method of a carrier rocket attitude control system in embodiment 1 of the present invention;
fig. 2 is a block diagram of a semi-physical simulation environment in a test method of a carrier rocket attitude control system in embodiment 1 of the present invention;
fig. 3 is a flowchart of a specific example of a test method of a launch vehicle attitude control system according to embodiment 1 of the present invention;
fig. 4 is a flowchart of another specific example of a test method of a launch vehicle attitude control system according to embodiment 1 of the present invention;
fig. 5 is a block diagram showing a specific example of a test apparatus for a posture control system of a launch vehicle in embodiment 1 of the present invention;
fig. 6 is a block diagram showing a specific example of a test system of a posture control system of a launch vehicle in embodiment 1 of the present invention;
fig. 7 is a block diagram showing a specific example of a first controller in a test system of a posture control system of a launch vehicle in embodiment 1 of the present invention.
Detailed Description
The technical solutions of the present invention will be described clearly and completely with reference to the accompanying drawings, and it should be understood that the described embodiments are some, but not all embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
In the description of the present invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, and are only for convenience of description and simplicity of description, but do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
In the description of the present invention, it should be noted that, unless otherwise explicitly specified or limited, the terms "mounted," "connected," and "connected" are to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral connection; can be mechanically or electrically connected; the two elements may be directly connected or indirectly connected through an intermediate medium, or may be communicated with each other inside the two elements, or may be wirelessly connected or wired connected. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.
In addition, the technical features involved in the different embodiments of the present invention described below may be combined with each other as long as they do not conflict with each other.
In the method, the control system of the carrier rocket is the core for ensuring the carrier rocket to normally fly, and the attitude control system is an important component of the control system of the carrier rocket, so that the working condition of the attitude control system can directly influence the correctness or the mistake of the operation of the carrier rocket, namely the success or the failure of the launching task of the carrier rocket, and the flight quality of the carrier rocket is directly influenced. And these are closely related to the life and property safety of people along the flight trajectory of the carrier rocket and the load orbit-entering precision, so that the operation correctness and performance of the attitude control system are very necessary to be tested and tested in the research and development process of the carrier rocket.
However, in the existing test method of the attitude control system of the carrier rocket, all components of the default carrier rocket normally operate, and the influence of the fault of the attitude control system on the carrier rocket is not considered. The historical data of the launch vehicle flight over the target time period indicates that the fault is the underlying factor in the launch vehicle flight, but not all faults will cause the launch vehicle to operate incorrectly. On the other hand, the deviation added by the existing carrier rocket attitude control system testing method aiming at certain parameters is usually obtained by calculation according to the tolerance and reasonable error generated in the carrier rocket manufacturing process, and the situation that certain parameters exceed the set value under the condition of carrier rocket fault is not considered.
The test method of the attitude control system of the carrier rocket can also be realized through a semi-physical simulation experiment, at the moment, the hardware of the attitude control system participating in the simulation experiment may have fault conditions, but the fault data is generally regarded as invalid data to be ignored in the existing test method. Specifically, the existing test method for the attitude control system of the carrier rocket mainly depends on frequency domain analysis, mathematical simulation and semi-physical simulation, and the test methods are all established on a mathematical model reflecting the kinematics and dynamics environment of the carrier rocket. Based on the mathematical models, the evaluation on the attitude control system of the carrier rocket can be completed by adding deviation for certain parameters which have influence on the operation of the attitude control system and matching with a simulated flight test.
Example 1
The embodiment of the invention provides a test method of a carrier rocket attitude control system, which is implemented in a preset attitude control system environment. The launch vehicle is generally of the 2-4 class and can be used to bring artificial earth satellites, manned spacecraft, space stations or interplanetary probes, etc. into a predetermined orbit. The final stage has an instrument cabin in which a guidance and control system, a remote measuring system and a transmitting field safety system are arranged. The effective load is arranged on the instrument cabin, and a fairing is sleeved outside the effective load. Each stage of which includes an arrow body structure, a propulsion system, and a flight control system. The stages are connected with each other by stage intervals. The effective load is arranged on the instrument cabin, and a fairing is sleeved outside the effective load.
In order to obtain the influence of various flight faults on the attitude control system, the test can be carried out in a simulation mode. For example, the preset test environment of the attitude control system may include: mathematical simulation environment and semi-physical simulation environment. In a mathematical simulation environment, as shown in fig. 1, specifically includes: the simulation upper computer and the real-time simulation lower computer; the real-time simulation lower computer runs various mathematical simulation models and modules, wherein the mathematical simulation models comprise: a guidance, navigation and control (GNC) module, an attitude control System actuator mathematical model, a dynamic mathematical model, an Inertial Measurement Unit (IMU) mathematical model, and a digital virtual global navigation Satellite model (GNSS).
In a semi-physical simulation environment, as shown in FIG. 2, comprising: the simulation system comprises a simulation upper computer, a real-time simulation lower computer and a real object, wherein the simulation upper computer consists of a simulation control upper computer and a simulation remote measuring upper computer; the real-time simulation lower computer runs various mathematical simulation modules, and the real-time simulation lower computer can specifically comprise: the system comprises a dynamics mathematical model, an attitude control system execution mechanism mathematical model, an IMU mathematical model and a digital virtual GNSS model; the real object part can comprise an rocket-borne computer, an executing mechanism, a loading platform, a rotary table, a gyroscope, a GNSS simulator and a GNSS receiver. Specifically, a person skilled in the art may add or delete some modules of the physical part according to actual needs, and the present invention is not limited to this.
Specifically, in the test method of the posture control system of the launch vehicle provided in the embodiment of the present invention, for various mathematical simulation models and modules running in the real-time simulation lower computer of fig. 1 and 2, the following table 1 is specifically described:
TABLE 1
The test method of the attitude control system of the carrier rocket provided by the embodiment of the invention can be realized by the test environment, and the attitude control system of the carrier rocket is tested by combining mathematical simulation and semi-physical simulation. Specifically, the mathematical simulation environment can be established in Matlab software on a Windows platform, a dynamic model, an execution mechanism model and the like of the launch vehicle are modeled by the Matlab and Simulink in the Matlab, and the flight of the launch vehicle can be simulated in the mathematical environment by matching with the GNC module. Testing an attitude control system according to simulated flight data of the carrier rocket in a mathematical simulation environment; the semi-physical simulation environment is based on mathematical simulation, and can replace a mathematical model part which can be replaced by a physical object in a test with the physical object according to the test requirement. In a semi-physical simulation environment, the GNC module runs on an rocket-borne computer, and simulated flight data of the carrier rocket can be acquired through actions of a simulation telemetering upper computer, a dynamic model running on a real-time simulation lower computer, an executing mechanism and a rotary table, so that the attitude control system of the carrier rocket is tested.
The test method of the carrier rocket of the embodiment of the invention can be applied to a specific application scene of testing an attitude control system in a carrier rocket development test stage, wherein the carrier rocket attitude control system is used for carrying out attitude control on a preset carrier rocket model, and as shown in figure 3, the method comprises the following steps:
step S11: inputting fault parameters in a carrier rocket attitude control system; in this embodiment, the attitude control system of the launch vehicle includes: the attitude control system comprises a control mechanism and an execution mechanism, and specifically, in a real attitude control system of a carrier rocket, the execution mechanism of the attitude control system specifically comprises: the System comprises a Reaction Control System (RCS), an air/gas rudder, an engine jet pipe swinging mechanism and the like, wherein each actuating mechanism has multiple implementation modes, in a mathematical simulation model, different implementation modes are abstracted to form the same mathematical model, and the injection faults are basically the same, so that the types of actuating mechanisms of the carrier rocket attitude Control System do not need to be distinguished specifically.
Illustratively, in a test environment of mathematical simulation, various faults are simulated by modifying parameters and logic in a mathematical model, namely, faults in a carrier rocket attitude control system can be simulated by inputting different fault parameters. In this embodiment, changing the parameter in the mathematical model may be changing the instruction size corresponding to the control instruction, for example, it may be proportionally increasing, and the control instruction can only output the maximum value or the minimum value; or in inverse proportion, the control output and the actuating mechanism do work in inverse proportion; specifically, the fault of the steering engine of the actuating mechanism of the attitude control system can be simulated by changing the deflection angle of the steering engine or changing the response time of the steering engine; the attitude control system has various actuators including RCS, air, gas vane, engine nozzle swing mechanism, etc. and may be changed in any parameters as long as the actuators in the attitude control system can be out of order.
Illustratively, the logic structure in the mathematical model can be changed by modifying the decoding tables and polarities of the RCS and the grid rudder to make the above elements dislocated so as to simulate the RCS in the attitude control system to be in fault or simulate the grid rudder in the attitude control system to be in fault. In actual use, as long as the logic that can make the attitude control system break down can be changed, the fault can be simulated, specifically, the steering engine in the attitude control system can be simulated to break down by changing the deflection angle of the steering engine.
For example, in a semi-physical simulation testing environment, when an actuator in the physical part is required to simulate a fault, the actuator may be switched to a real corresponding mathematical model before simulating the fault.
Step S12: acquiring first flight state data of the carrier rocket after fault injection according to the input fault parameters; in this embodiment, after the simulated attitude control system fails, the flight status data of the launch vehicle can be obtained in the mathematical simulation environment, specifically, the flight status data includes the value of the flight index, for example, flight criteria may include three degrees of freedom speed, three degrees of freedom pose, attitude angle rate, fly height, mach number, dynamic pressure, Alpha-Q, angle of attack, sideslip angle, ballistic error, and in particular, in the mathematical model, the IMU of the launch vehicle can be simulated according to the IMU mathematical model, various angular velocities and accelerations of the launch vehicle can be directly obtained through the IMU, the angles and the velocities can be obtained through the calculation of the inertia measurement unit, namely, the integral operation, in addition, through the IMU mathematical model and the information provided by the digital virtual GNSS, the values of the flight parameters can be calculated through an algorithm in the GNC module.
Step S13: judging whether the fault causes the operation error of the carrier rocket or not according to the first flight state data; in this embodiment, according to the flight data of the carrier rocket obtained in the above steps, it can be determined whether the carrier rocket can normally fly or can normally execute a launch task of the carrier rocket after the attitude control system fails, specifically, the tolerance of the carrier rocket to the failure is determined; namely, whether the influence of the fault on the task execution of the carrier rocket does not influence the successful task execution or causes the task failure is judged.
For example, when a steering engine in an actuator of an attitude control system is simulated to have a fault, during normal operation of the launch vehicle, a normal deflection angle of the steering engine is 10 degrees, and when a control command is modified to make the steering engine deflect only 9 degrees, the steering engine has a fault, but normal flight of the launch vehicle at the moment is judged through acquired flight data of the launch vehicle, such as three-degree-of-freedom speed, three-degree-of-freedom pose, attitude angle rate, height, mach number, dynamic pressure, Alpha-Q, attack angle, sideslip angle, trajectory error and the like, by using values of flight parameters calculated by an algorithm in a GNC module, the fault of 9 degrees of deflection of the steering engine of the attitude control system can not cause operation errors of the launch vehicle, and at the moment, the fault can be tolerated by the launch vehicle.
When the control instruction is modified to enable the steering engine to deflect by 0.1 degree, the steering engine breaks down, the carrier rocket is judged to be incapable of flying normally through the acquired flight data of the carrier rocket, such as three-degree-of-freedom speed, three-degree-of-freedom pose, attitude angle rate, height, Mach number, dynamic pressure, Alpha-Q, attack angle, sideslip angle, trajectory error and the like, by the values of flight parameters solved through the algorithm in the GNC module, and then the fault, caused by the 0.1-degree-of-deflection fault of the steering engine, of the attitude control system can cause the running fault of the carrier rocket, and the fault can not be tolerated by the carrier rocket.
Therefore, the specific relation between the normal flight of the carrier rocket and the command fault of the steering engine is obtained through the method. By implementing the method, the fault type which does not influence the normal operation of the carrier rocket can be obtained, the fault type which can directly cause the operation failure of the carrier rocket can also be obtained, the flight condition of the carrier rocket under the fault condition of the attitude control system can be analyzed, and the test coverage range of the carrier rocket simulation test is expanded.
Step S14: and when the fault causes the carrier rocket to have an error in operation, recording fault parameters and first flight state data. In this embodiment, after the fault parameter is input, that is, after the parameter or logic of the attitude control system is changed and the simulated fault occurs, the flight state of the launch vehicle changes, at this time, the flight state data of the launch vehicle recorded in the mathematical simulation environment may be flight data of each flight phase, specifically, the flight data may be data when the launch vehicle executes a launch task, and may be carrying capacity, orbit accuracy, and adaptability and reliability of the launch vehicle to payloads of different weights. For example, launch capacity refers to the weight of the payload that a launch vehicle can deliver into a predetermined trajectory. When the orbit types of the effective load are different, the required energy is also different, so that the conditions of a low orbit, a sun synchronous orbit, a geosynchronous satellite transition orbit, a planetary detector orbit and the like are distinguished when the carrying capacity is marked; or the carrying capacity may be expressed by the payload weight at which the rocket reaches a certain characteristic speed.
Illustratively, by simulating the occurrence of the fault, the influence of a specific fault combination of the attitude control system of the carrier rocket on the flight of the carrier rocket can be obtained, and then the injected fault is associated with corresponding flight data and an operation result, namely, fault parameters and flight state data are recorded, so that the corresponding relation between different injected faults and the flight state of the carrier rocket is obtained.
The test method of the carrier rocket attitude control system provided by the embodiment of the invention comprises the following steps: simulating the carrier rocket attitude control system to have a fault by inputting fault parameters into the carrier rocket attitude control system, and then acquiring first flight state data of the carrier rocket after the fault is injected according to the input fault parameters; judging whether the fault causes the operation error of the carrier rocket or not according to the first flight state data; and when the fault causes the carrier rocket to have an error in operation, recording fault parameters and first flight state data. By implementing the test method of the invention and combining the fault type in the attitude control system and the running data of the carrier rocket, the problems that each system of the default carrier rocket normally runs and the fault condition of the carrier rocket is ignored in the existing test method of the attitude control system of the carrier rocket are solved, the flight safety performance of the carrier rocket is ensured, and the test method of the carrier rocket is more comprehensive and more accurate. That is, the fault type information of the carrier rocket which is fatal to the carrier rocket can be obtained by simulating the flight state of the carrier rocket after the fault occurs in the attitude control system, the fault injection controllability based on the simulation environment is high, the fault mode switching is simple, the cost is low, the fault test performed in the real-time simulation is high in fidelity and is more real and credible, and the test requirements of the attitude control system of the carrier rocket are met through actual flight inspection in the development process of a certain commercial carrier rocket.
As an optional embodiment of the present application, when step S11 is executed and the fault parameter is input in the posture control system of the launch vehicle, the method specifically includes: inputting a fault parameter in a flight phase of the launch vehicle, or inputting fault parameters in a flight phase of the launch vehicle, or inputting a fault parameter in multiple flight phases of the launch vehicle, or inputting fault parameters in multiple flight phases of the launch vehicle. In this embodiment, the phase of flight of the launch vehicle may include: the method comprises an ignition stage, a takeoff stage, a first-level flight stage, a second-level separation stage, a second-level flight stage, a third-level separation stage, a fairing separation stage, a fourth-level flight stage, a satellite-rocket separation stage and a track passivation stage. The input fault types may be: response time delay, thrust decline, unable change-over switch state, thrust direction skew, no response in the dead band, steering wheel control surface card is dead, control surface efficiency reduces, response angle error, steering wheel card are dead etc. Specifically, one or more types of faults can be input in multiple flight stages of the launch vehicle, or one or more types of faults can be input in a certain flight stage of the launch vehicle, and different fault permutation and combination can form a fault matrix; for example, it may be that a response time delay fault is input during the takeoff phase; inputting a response time delay fault and a response angle error fault in a takeoff stage; inputting response time delay faults in a takeoff phase and an ignition phase; response time delay faults and response angle error faults are input in a takeoff phase and an ignition phase.
The test method of the attitude control system of the carrier rocket provided by the embodiment of the invention can also simulate various faults by inputting various fault parameters to test the flight condition of the carrier rocket, can inject the same fault in one flight stage or a plurality of flight stages of the carrier rocket, and can inject various faults in one flight stage or a plurality of flight stages, so that the type and the number of fault injection in the test of the attitude control system of the carrier rocket can be reasonably designed and set, the problem which possibly occurs in the actual flight test of the attitude control system of the carrier rocket can be correctly reflected, the fault condition of the attitude control system of the carrier rocket can be comprehensively simulated, and the test result of the carrier rocket is more accurate.
As an optional embodiment of the present application, the method for testing a launch vehicle attitude control system, as shown in fig. 4, further includes:
step S41: when the carrier rocket is not in error due to faults, increasing fault parameters; in this embodiment, increasing the fault parameter may be increasing the severity of the fault by changing parameters or logic information of the model in the mathematical simulation environment.
Step S42: and acquiring second flight state data of the carrier rocket after fault injection according to the increased fault parameters.
Step S43: and judging whether the fault causes the operation error of the carrier rocket or not according to the second flight state data.
Step S44: and if the fault causes the carrier rocket to operate mistakenly, recording the increased fault parameters and the second flight state data.
Step S45: if the fault causes no error in the operation of the carrier rocket, the fault parameters are continuously increased until the fault causes the error in the operation of the carrier rocket.
The test method of the carrier rocket attitude control system provided by the embodiment of the invention can test the robustness and performance boundary of the attitude control system by gradually increasing the fault parameters of the faults, namely gradually amplifying the severity of the faults and circularly iterating the test process aiming at the fault types which do not directly cause the flight failure of the carrier rocket, thereby achieving the aim of comprehensively testing the carrier rocket attitude control system and ensuring the accuracy of the carrier rocket test.
Example 2
The test device of the carrier rocket of the embodiment of the invention can be applied to a specific application scene of testing an attitude control system in a development and test stage of the carrier rocket, wherein the attitude control system of the carrier rocket is used for carrying out attitude control on a preset carrier rocket model, and as shown in figure 5, the device comprises:
an input module 51 for inputting fault parameters in the launch vehicle attitude control system; for details of implementation, reference is made to the description of step S11 in the above method embodiment.
The obtaining module 52 is configured to obtain first flight state data of the carrier rocket after the fault is injected according to the input fault parameter; for details of implementation, reference is made to the description of step S12 in the above method embodiment.
The judging module 53 is configured to judge whether the fault causes an error in operation of the carrier rocket according to the first flight state data; for details of implementation, reference is made to the description of step S13 in the above method embodiment.
The recording module 54 is used for recording fault parameters and first flight state data when the fault causes the carrier rocket to have an error in operation; for details of implementation, reference is made to the description of step S14 in the above method embodiment.
The embodiment of the invention discloses a test device of a carrier rocket attitude control system, wherein the carrier rocket attitude control system is used for carrying out attitude control on a preset carrier rocket model, and the device comprises: inputting fault parameters in a carrier rocket attitude control system through an input module; acquiring first flight state data of the carrier rocket after fault injection according to the input fault parameters through an acquisition module; judging whether the fault causes the operation error of the carrier rocket or not through a judging module; and recording the fault parameters and the first flight state data by the recording module when the fault causes the carrier rocket to have an error in operation. By implementing the method, the problem that each system of the carrier rocket is defaulted to normally operate and the fault condition of the carrier rocket is ignored in the conventional test method of the attitude control system of the carrier rocket is solved by combining the fault type in the attitude control system and the operation data of the carrier rocket, the flight safety performance of the carrier rocket is ensured, and the test method of the carrier rocket is more comprehensive and more accurate.
Example 3
The embodiment of the invention provides a test system of a carrier rocket attitude control system, which comprises at least one control device 61, wherein the control device 61 is used for executing the steps of the test method of the carrier rocket attitude control system in any one of the above embodiments.
As shown in fig. 6, the control device 61 includes:
the first communication module 611: for transmitting data, receiving and transmitting fault parameters in the attitude control system of the launch vehicle, which are input according to the mathematical model, and acquiring flight state data of the guided launch vehicle. The first communication module can be a Bluetooth module and a Wi-Fi module, and then communication is carried out through a set wireless communication protocol.
The first controller 612: connected to the first communication module 611, as shown in fig. 7, includes: at least one processor 71; and a memory 72 communicatively coupled to the at least one processor 71; in the embodiment, the first communication module may be a wireless communication module, such as a bluetooth module, a Wi-Fi module, or the like, or a wired communication module, and the processor 71 and the memory 72 are connected through a bus 70, which is taken as an example in fig. 7. The transmission between the first controller 612 and the first communication module 611 is wireless transmission.
The memory 72 is a non-transitory computer readable storage medium, and can be used for storing non-transitory software programs, non-transitory computer executable programs, and modules, such as program instructions/modules corresponding to the test method of the launch vehicle attitude control system in the embodiments of the present application. The processor 71 executes various functional applications of the server and data processing by running non-transitory software programs, instructions and modules stored in the memory 72, namely, the test method of the launch vehicle attitude control system of the above-described method embodiment is realized.
The memory 72 may include a storage program area and a storage data area, wherein the storage program area may store an operating system, an application program required for at least one function; the storage data area may store data created according to use of a processing device operated by the server, and the like. Further, the memory 72 may include high speed random access memory, and may also include non-transitory memory, such as at least one magnetic disk storage device, flash memory device, or other non-transitory solid state storage device. In some embodiments, the memory 72 may optionally include memory located remotely from the processor 71, which may be connected to a network connection device via a network. Examples of such networks include, but are not limited to, the internet, intranets, local area networks, mobile communication networks, and combinations thereof.
One or more modules are stored in the memory 72 and, when executed by the one or more processors 71, perform the method described in any of the above embodiments.
The embodiment of the present invention further provides a non-transitory computer readable medium, where the non-transitory computer readable storage medium stores a computer instruction, and the computer instruction is used to enable a computer to execute the method for testing the attitude control system of a launch vehicle described in any one of the above embodiments, where the storage medium may be a magnetic Disk, an optical Disk, a Read-only Memory (ROM), a Random Access Memory (RAM), a flash Memory (FlashMemory), a Hard Disk (Hard Disk Drive, abbreviated as HDD), or a Solid-State Drive (SSD); the storage medium may also comprise a combination of memories of the kind described above.
It should be understood that the above examples are only for clarity of illustration and are not intended to limit the embodiments. Other variations and modifications will be apparent to persons skilled in the art in light of the above description. And are neither required nor exhaustive of all embodiments. And obvious variations or modifications therefrom are within the scope of the invention.
Claims (8)
1. A test method of a carrier rocket attitude control system, wherein the carrier rocket attitude control system is used for carrying out attitude control on a preset carrier rocket model, and is characterized by comprising the following steps:
inputting fault parameters in a carrier rocket attitude control system;
acquiring first flight state data of the carrier rocket after fault injection according to input fault parameters;
judging whether the fault causes the carrier rocket to operate mistakenly or not according to the first flight state data;
when the fault causes the carrier rocket to have an error in operation, recording the fault parameter and the first flight state data;
when the fault causes the carrier rocket not to operate mistakenly, increasing the fault parameter;
acquiring second flight state data of the carrier rocket after fault injection according to the increased fault parameters;
judging whether the fault causes the carrier rocket to operate mistakenly or not according to the second flight state data;
if the fault causes the carrier rocket to operate mistakenly, recording the increased fault parameters and the second flight state data;
if the fault causes the carrier rocket to operate without errors, continuing to increase the fault parameters until the fault causes the carrier rocket to operate with errors; and recording fault parameters causing the operation error of the carrier rocket and corresponding flight state data thereof.
2. The method of claim 1, wherein the inputting fault parameters in the launch vehicle attitude control system comprises:
inputting a fault parameter during a flight phase of said launch vehicle, or
Inputting a plurality of fault parameters in a flight phase of the launch vehicle, or
Inputting a fault parameter during multiple flight phases of the launch vehicle, or
Inputting a plurality of fault parameters at a plurality of flight phases of the launch vehicle.
3. The method of claim 1, wherein the flight status data comprises one or more of a flight speed or a flight angle or an altitude or a flight attitude.
4. The method of claim 1, wherein the input fault parameter is implemented by changing a control parameter in an attitude control system.
5. A test device of a carrier rocket attitude control system, wherein the carrier rocket attitude control system is used for carrying out attitude control on a preset carrier rocket model, and is characterized by comprising the following components:
the input module is used for inputting fault parameters in the carrier rocket attitude control system;
the first acquisition module is used for acquiring first flight state data of the carrier rocket after fault injection according to input fault parameters;
the first judging module is used for judging whether the fault causes the carrier rocket to operate mistakenly or not according to the first flight state data;
the first recording module is used for recording the fault parameters and the first flight state data when the fault causes the carrier rocket to have an error in operation;
the increasing module is used for increasing the fault parameter when the fault causes that the carrier rocket does not have fault in operation;
the second acquisition module is used for acquiring second flight state data of the carrier rocket after fault injection according to the increased fault parameters;
the second judgment module is used for judging whether the fault causes the carrier rocket to operate mistakenly or not according to the second flight state data;
the second recording module is used for recording the increased fault parameters and the second flight state data if the fault causes the carrier rocket to have errors in operation;
the third recording module is used for continuously increasing the fault parameters until the fault causes the carrier rocket to operate mistakenly if the fault causes the carrier rocket to operate not mistakenly; and recording fault parameters causing the operation error of the carrier rocket and corresponding flight state data thereof.
6. The apparatus according to claim 5, wherein the input module specifically includes:
a first input submodule for inputting a fault parameter during a flight phase of said launch vehicle; or
Inputting a plurality of fault parameters in a flight phase of the launch vehicle, or
Inputting a fault parameter during multiple flight phases of the launch vehicle, or
Inputting a plurality of fault parameters at a plurality of flight phases of the launch vehicle.
7. A test system for a launch vehicle attitude control system, comprising:
at least one control device for performing the steps of the method of testing a posture control system of a launch vehicle according to any one of claims 1-4, acquiring a flight state of the launch vehicle according to the type of fault injected, and determining whether the fault causes the launch vehicle to fail.
8. A computer-readable storage medium, on which a computer program is stored, which, when being executed by a processor, carries out the steps of the method of testing a launch vehicle attitude control system according to any one of claims 1 to 4.
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