Attitude control simulation method and system for launch vehicle

Abstract

An attitude control simulation method for a launch vehicle comprises: receiving a pulse accumulation sent by a laser strapdown inertial measurement unit (11); performing calculation processing according to the pulse accumulation to obtain target calculation results (12), wherein the target calculation results comprise a navigation calculation result, a guidance calculation result and an attitude control calculation result; and sending the navigation calculation result, the guidance calculation result and the attitude control calculation result to a simulation server side (13). By means of the attitude control simulation method for a launch vehicle, the problems whereby a real environment is difficult to be simulated, a verification method is not sufficiently intuitive and the safety risk is high when a launch vehicle is tested are solved, and the simulation and optimization of the launch vehicle on a computer under different working conditions are realized without experimental verification, thereby saving on the time and cost, and accurately predicting attitude responses and control effects of the launch vehicle under different working conditions.

Description

A launch vehicle attitude control simulation method and system CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] This application claims priority to Chinese Patent Application No. 2023106101236 filed in China on May 29, 2023, the entire contents of which are incorporated herein by reference. Technical Field

[0002] The present disclosure relates to the technical field of launch vehicle simulation, and in particular to a launch vehicle attitude control simulation method and system. Background Art

[0003] Digital twin technology refers to the use of computer technology to create virtual, one-to-one digital models of physical systems. Digital twin technology is widely used in the launch vehicle sector, helping to better understand and predict rocket behavior, optimize rocket design and operation, and improve rocket reliability and safety. Within the launch vehicle sector, digital twins can be applied in many areas, including optimizing rocket design and manufacturing processes, facilitating virtual testing and simulation on computers to verify and optimize design parameters, predicting rocket performance and diagnosing faults, and simulating and predicting rocket behavior under different circumstances to better understand and analyze rocket performance and issues. However, the existing digital twin simulation of launch vehicles still has the following problems: 1) It is difficult to simulate the real environment. The verification process of the rocket attitude control algorithm usually needs to be carried out in a real environment. However, in the traditional semi-physical simulation experiment before the rocket launch, it is difficult to fully simulate the real flight environment, which will cause problems in the actual use of the algorithm; 2) The verification method is not intuitive enough. Traditional simulation experiments interpret data by observing the changing curves. This data interpretation method has limitations and requires professional personnel to interpret it. It is not intuitive enough and cannot fully see the entire rocket's trajectory and structural changes; 3) Poor repeatability. Traditional simulation experiments usually involve multiple environmental parameters and external interference factors, which makes it difficult to ensure the consistency and repeatability of the verification results; 4) High safety risks. Traditional simulation verification methods need to be tested and tested during the actual launch process, which increases the safety risks during the flight and the trial and error costs are high; 5) Long verification cycle. Traditional verification methods involve multiple environmental parameters and external interference factors. The verification process is relatively cumbersome and may require multiple experiments and tests to ensure the accuracy of the verification results. Therefore, the verification cycle is relatively long. Technical issues

[0004] The present disclosure provides a management method and device for virtual reality applications to solve the problems of difficulty in simulating a real environment during launch vehicle testing, lack of intuitive verification methods, and high safety risks. Technical Solutions

[0005] To solve the above technical problems, the technical solutions disclosed in this disclosure are as follows:

[0006] Disclosed is a launch vehicle attitude control simulation method, which is applied to a flight control module and includes:

[0007] The full amount of pulses sent by the laser strapdown inertial system is received; a solution process is performed according to the full amount of pulses to obtain a target solution result; the target solution result includes a navigation solution result, a guidance solution result, and an attitude control solution result; and the navigation solution result, the guidance solution result, and the attitude control solution result are sent to a simulation server.

[0008] Optionally, performing a solution process based on the full pulse amount to obtain a target solution result includes:

[0009] Tool error compensation is performed on the full pulse amount to obtain apparent velocity increments and angular increments in the first direction, the second direction, and the third direction; navigation solution processing is performed on the apparent velocity increments and the angular increments to obtain a navigation solution result; the navigation solution result includes at least one of an attitude angle, an altitude, a relative velocity inclination, an apparent acceleration mean, and an angular velocity mean; guidance solution processing is performed on the navigation solution result to obtain a guidance solution result; the guidance solution result includes a separation timing and a program angle; attitude control solution processing is performed based on the guidance solution result and the navigation solution result to obtain an attitude control solution result.

[0010] Optionally, performing attitude control solution processing according to the guidance solution result and the navigation solution result to obtain an attitude control solution result includes:

[0011] Determine a target core stage to be attitude controlled at the current moment; the target core stage includes at least one of a booster stage, a core stage 1, a core stage 2, a core stage 3 and a final stage; perform attitude control solution processing on the target core stage according to the guidance solution result and the navigation solution result to obtain an attitude control solution result.

[0012] Optionally, when the target core stage is a booster stage and a core stage, performing attitude control solution processing on the target core stage according to the guidance solution result and the navigation solution result to obtain an attitude control solution result includes:

[0013] Determine a target transfer function; process the program angle in the guidance solution result through the target transfer function to obtain a first target booster stage swing angle; perform swing angle limiting calculation on the first target booster stage swing angle to obtain a second target booster stage swing angle after limiting; determine a first target equivalent swing angle of the three channels of the booster stage based on the second target booster stage swing angle; perform instruction allocation based on the first target equivalent swing angle to determine a first servo swing angle instruction; obtain an attitude control solution result for the booster stage based on the guidance timing corresponding to the core first stage at the current moment; wherein, if the guidance timing is not the ignition moment, the first servo swing angle instruction is output as the attitude control solution result, otherwise, the second servo swing angle instruction corresponding to the core first stage is output as the attitude control solution result.

[0014] Optionally, when the target core stage is core one, core two, or core three, performing attitude control solution processing according to the guidance solution result and the navigation solution result to obtain an attitude control solution result includes:

[0015] The operation phase of the target core stage is determined according to the separation timing in the guidance solution result; and the attitude control solution result is determined according to the operation phase.

[0016] Optionally, when the target core stage is the final stage, performing attitude control solution processing according to the guidance solution result and the navigation solution result to obtain an attitude control solution result includes:

[0017] The program angle in the guidance solution is processed through a digital correction network to obtain a correction value; the correction value is threshold-controlled to determine a fusion communication control instruction and output the fusion communication control instruction as a posture control solution result.

[0018] The present disclosure also provides a launch vehicle attitude control simulation method, which is applied to a simulation server and includes:

[0019] The method comprises the steps of: receiving a start signal sent by a simulation client; sending a master command signal to a flight control module according to the start signal; sending simulated position information to a laser strapdown inertial system (LSINS) so that the LINS sends a full pulse quantity calculated according to the simulated position information to the flight control module; receiving a navigation solution result, a guidance solution result, and an attitude control solution result sent by the flight control module; performing model solution processing according to the navigation solution result, the guidance solution result, and the attitude control solution result to obtain a model solution result; sending the model solution result to the simulation client and the LINS respectively so that the simulation client performs a three-dimensional display; and the model solution result including the latest simulated position information.

[0020] The present disclosure provides a launch vehicle attitude control simulation system, comprising:

[0021] A flight control module; a simulation client electrically connected to the flight control module; a simulation server communicatively connected to the simulation client and electrically connected to the flight control module; a laser strapdown inertial system (LSINS) unit (LSU) communicatively connected to the simulation server and the flight control module; wherein the flight control module receives a full pulse quantity sent by the LSI; the flight control module performs a solution processing based on the full pulse quantity to obtain a target solution result; the target solution result includes a navigation solution result, a guidance solution result, and an attitude control solution result; and the flight control module sends the navigation solution result, the guidance solution result, and the attitude control solution result to the simulation server.

[0022] Optionally, the flight control module includes:

[0023] An inertial navigation module, the inertial navigation module is communicatively connected to the laser strapdown inertial group; a guidance module, the guidance module is communicatively connected to the inertial navigation module; and an attitude control module, the attitude control module is communicatively connected to the inertial navigation module and the guidance module respectively; wherein the inertial navigation module is used to perform navigation solution processing on the full pulse amount to obtain a navigation solution result; the navigation solution result includes at least one of an attitude angle, an altitude, a relative velocity inclination angle, an apparent acceleration mean, and an angular velocity mean; the guidance module is used to perform guidance solution processing on the navigation solution result to obtain a guidance solution result; the guidance solution result includes a separation timing and a program angle; the attitude control module is used to perform attitude control solution processing based on the guidance solution result and the navigation solution result to obtain an attitude control solution result.

[0024] Optionally, the simulation server is connected to the simulation client via Ethernet communication; the laser strapdown inertial group is connected to the simulation server and the flight control module via a communication bus. Beneficial effects

[0025] The above solution of the present disclosure includes at least the following beneficial effects:

[0026] The above-mentioned scheme disclosed in the present invention receives the full amount of pulses sent by the laser strapdown inertial group; performs solution processing according to the full amount of pulses to obtain a target solution result; the target solution result includes a navigation solution result, a guidance solution result and an attitude control solution result; and sends the navigation solution result, the guidance solution result and the attitude control solution result to a simulation server. This solves the problems of difficulty in simulating the real environment during launch vehicle testing, the lack of intuitiveness in the verification method and the high safety risk, and realizes the simulation and optimization of the launch vehicle under different working conditions on a computer without the need for experimental verification, saving time and cost, and accurately predicting the attitude response and control effect of the rocket under different working conditions. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] FIG1 is a flow chart of a method for simulating attitude control of a launch vehicle applied to a flight control module according to an embodiment of the present disclosure;

[0028] FIG2 is a schematic diagram of a framework of a launch vehicle attitude control simulation system according to a specific embodiment of the present disclosure;

[0029] FIG3 is a flow chart of attitude control solution processing when the target core stage is the core stage and the booster stage in a specific embodiment provided by the present disclosure;

[0030] FIG4 is a flow chart of attitude control solution processing when the target core stage is the first core stage in a specific embodiment provided by the present disclosure;

[0031] FIG5 is a flow chart of attitude control solution processing when the target core stage is a core stage 2 in a specific embodiment provided by the present disclosure;

[0032] FIG6 is a flow chart of attitude control solution processing when the target core stage is a core three stage in a specific embodiment provided by the present disclosure;

[0033] 7 is a schematic diagram of a flow chart of attitude control solution processing when the target core stage is the final stage in a specific embodiment provided by the present disclosure;

[0034] FIG8 is a flow chart of a method for simulating the attitude control of a launch vehicle applied to a simulation server according to an embodiment of the present disclosure;

[0035] FIG9 is a schematic diagram of a system flow chart of a method for simulating attitude control of a launch vehicle in a specific embodiment of the present disclosure;

[0036] FIG10 is a schematic diagram of the overall implementation architecture of the attitude control simulation method for a launch vehicle in a specific embodiment provided by the present disclosure. Best Mode for Carrying Out the Invention

[0037] As shown in FIG1 , an embodiment of the present disclosure provides a launch vehicle attitude control simulation method, which is applied to a flight control module and includes:

[0038] Step 11, receiving the full amount of pulses sent by the laser strapdown inertial system;

[0039] Step 12: performing a calculation based on the full pulse quantity to obtain a target calculation result; the target calculation result includes a navigation calculation result, a guidance calculation result, and an attitude control calculation result;

[0040] Step 13: Send the navigation solution result, the guidance solution result, and the attitude control solution result to a simulation server.

[0041] In this embodiment, a method for simulating the attitude control of a carrier rocket involves a flight control module, a simulation server, a laser strapdown inertial system (LSINS) and a simulation client, wherein the simulation client is used to provide control instructions to the flight control module and the simulation server, and can also visually display the flight status of the carrier rocket; the simulation server can perform model solution according to the instructions of the simulation client, and provide simulated position information to the laser strapdown inertial system; the laser strapdown inertial system is used to determine the full pulse amount based on the simulated position information; the flight control module is used to perform navigation solution, guidance solution and attitude control solution based on the full pulse amount provided by the laser strapdown inertial system, and provide the solution results to the simulation server, so that the simulation server can control the simulation client to visually display the flight status of the carrier rocket based on the solution results, solving the problems of difficulty in simulating the real environment during carrier rocket testing, lack of intuitive verification methods and high safety risks, and realizing the simulation and optimization of carrier rockets under different working conditions on a computer without the need for experimental verification, saving time and cost, and accurately predicting the attitude response and control effect of the rocket under different working conditions.

[0042] The following is a detailed description of the functions of the flight control module, simulation server, laser strapdown inertial system, and simulation client:

[0043] The simulation client and the simulation server together constitute a digital twin simulation system; wherein, the simulation server uses Veristand software (software for real-time testing and simulation, which can quickly build real-time test systems, hardware-in-the-loop HIL test systems and embedded software verification systems) and runs in the PharLap real-time operating system; the simulation server can receive data, solve models and output data, wherein data reception refers to receiving the attitude control solution results transmitted by the flight control module through the 1553B bus, and receiving the ignition and separation status of each core stage through the IO (Input / Output) level signal. The ignition and separation status of the core stage can be determined by the separation timing in the guidance solution result; the model solution is solved by the cooperation of multiple mathematical models. Specifically, the numerical values ​​between multiple mathematical models are transferred to form a closed loop of the simulation process to realize the solution based on the attitude control solution result and obtain the model solution result; in a preferred implementation example, the mathematical model includes a rocket body model, a six-degree-of-freedom simulation calculation model, an inertial group model, and an RCS (Rich Communication Suite, integrated communication) control model, engine model, and environment model; data output refers to sending part of the model solution results to the simulation client through Ethernet (Transmission Control Protocol TCP) for 3D display, and sending the other part of the model solution results to the laser strapdown inertial system through the 1553B bus for further solution.

[0044] The simulation client preferably uses CAD (computer-aided design) engineering drawing software and Unity (creation engine) software; the simulation client is equipped with a simulated rocket body model, simulation software scene and a visual two-dimensional display interface; among them, the rocket body model constructed by simulation is preferably a three-dimensional rocket body model, which is generated by digital twin technology to achieve high-precision simulation and prediction; specifically, the three-dimensional model can be composed of four booster stage servos, core first stage servo, core second stage servo, core third stage servo, fairing and satellite, etc., and by receiving instructions sent by the simulation server, the rocket body model is controlled to perform booster stage ignition separation, core first stage ignition separation, core second stage ignition separation, core third stage ignition, satellite-rocket separation and RCS nozzle ignition and other state changes; the rocket body model includes rotation The motion model and translational motion model, that is, rotation and translation can be realized in three-dimensional space, provide necessary input data for the simulation process; the simulation software scenarios include sky box scenarios, land scenarios, ocean scenarios and space scenarios, etc., and of course other scenarios can also be used, but this application is not limited to this; based on the simulation software scenarios, launch scenarios can be selected for simulated carrier rockets, such as land launch or ocean launch; in a specific implementation example, after the carrier rocket is launched, during the booster stage and the core stage, the carrier rocket operates in the atmosphere, so the simulation software scenario where the booster stage and the core stage are located is a land scene or an ocean scene, and in the core stage, the carrier rocket operates outside the atmosphere, so the simulation software scenario corresponding to the core stage is a space scene.

[0045] The two-dimensional interface is used to intuitively display the current status of the launch vehicle on the simulation client. The software sets up a two-dimensional display interface. The displayed values ​​include the rocket's current longitude, latitude and altitude, the rocket's pitch, heading, roll attitude angle, the rocket's current stage, the rocket's servo actuator swing angle value and at least one of the rocket's RCS nozzle switch status. The specific display data can be adjusted according to actual conditions. It is worth noting that other modules can also be set up in the simulation client, such as satellite models, flame smoke particles, etc., which can be adjusted according to actual conditions.

[0046] As shown in Figure 2, in a specific implementation example, the flight control module is provided with on-rocket software, which is used for flight control solution. The simulation server is provided with a rocket body model, an environment model, a six-degree-of-freedom model, an engine model and a fusion communication model, all of which are mathematical models; the simulation client is provided with a three-dimensional rocket body model, a sky and earth model (i.e., the above-mentioned simulation software scenario), a satellite model, a two-dimensional numerical interface and flame and smoke particles; the instructions or data of the on-rocket software are transmitted to the simulation server after data conversion processing. Specifically, the data messages transmitted by the on-rocket software in the 1553B or 422 serial port communication mode are converted into transmission data of optical fiber communication and sent to the simulation server for model solution. The simulation server and the simulation client are connected through Ethernet communication, and the data transmission protocol is TCP protocol. Modes for Carrying Out the Invention

[0047] In an optional embodiment of the present disclosure, step 12 includes:

[0048] Step 121, performing tool error compensation on the full pulse amount to obtain apparent velocity increments and angular increments in the first direction, the second direction, and the third direction;

[0049] Step 122: performing navigation solution processing on the apparent velocity increment and the angular increment to obtain a navigation solution result; the navigation solution result includes at least one of an attitude angle, an altitude, a relative velocity inclination angle, an apparent acceleration mean, and an angular velocity mean;

[0050] Step 123, performing guidance solution processing on the navigation solution result to obtain a guidance solution result; the guidance solution result includes a separation timing and a program angle;

[0051] Step 124 : performing attitude control solution processing according to the guidance solution result and the navigation solution result to obtain an attitude control solution result.

[0052] In an embodiment of the present disclosure, a full pulse amount is provided by a laser strapdown inertial system (LISUS). The LSSUS obtains longitude, latitude, altitude, angular velocity, and acceleration information (i.e., simulated position information) from a simulation server via a 1553B bus. The LSSUS solves the simulated position information and outputs a full pulse amount. The full pulse amount is preferably the full pulse amount of positive and negative directions of the accelerometer in the X, Y, and Z directions, and the full pulse amount of positive and negative directions of the gyroscope in the X, Y, and Z directions, totaling 12 full pulse amounts. After receiving the full pulse amount via 1553B, the flight control module performs tool error compensation on the full pulse amount to obtain apparent velocity increments and angular increments in the first, second, and third directions. The first direction is preferably the aforementioned X direction, the second direction is preferably the Y direction, and the third direction is preferably the Z direction. Since the flight control module includes an inertial navigation module, a guidance module, and an attitude control module, each module in the flight control module processes the apparent velocity increments and angular increments in the three directions in turn. Specifically:

[0053] The apparent velocity increment and angular increment are transmitted to the inertial navigation module for navigation solution processing to obtain a navigation solution result, which includes a three-channel attitude angle, altitude, relative velocity inclination angle, apparent acceleration mean, and angular velocity mean. The navigation solution result is transmitted to the guidance module for guidance solution processing, and the separation timing and program angle of the target core stage are output. The separation timing can specifically include the ignition timing and separation timing of the booster stage, core stage 1, core stage 2, and core stage 3, or the separation timing of the fairing and / or satellite, or the separation timing of the camera and matching resistor switching, and can be adjusted according to actual needs. The program angle is preferably a three-channel program angle, which is used for the attitude control module to perform attitude control solution processing. The attitude control module receives the apparent acceleration mean, angular velocity mean, and quaternion in the navigation solution result output by the inertial navigation module, receives the timing control signal (for indicating the separation timing) and the three-channel program angle from the guidance module, and then performs attitude control solution processing for the target core stage, outputting the attitude control solution result.

[0054] In an optional embodiment of the present disclosure, step 124 includes:

[0055] Step 1241, determining the target core stage to be attitude controlled at the current moment; the target core stage includes at least one of a booster stage, a core stage 1, a core stage 2, a core stage 3, and a final stage;

[0056] Step 1242: Perform attitude control solution processing on the target core stage according to the guidance solution result and the navigation solution result to obtain an attitude control solution result.

[0057] In an embodiment of the present disclosure, a target core stage to be subjected to attitude control at a current moment is located; each target core stage is provided with a servo controller, specifically including at least one of a booster stage, a core stage 1, a core stage 2, a core stage 3, and a final stage; the attitude of the target core stage is the control object, and the target core stage can be adjusted according to the physical model of the actual simulated launch vehicle; that is, it is possible to simulate attitude control of the launch vehicle according to its actual composition and structure;

[0058] Based on the separation timing in the guidance solution result and the average apparent acceleration, average angular velocity and quaternion in the navigation solution result, the target core stage is subjected to attitude control solution processing to obtain the attitude control solution result, which includes servo command information (such as servo swing angle command) and RCS attitude control command.

[0059] In an optional embodiment of the present disclosure, when the target core stage is a booster stage and a core stage, step 1242 includes:

[0060] Step 1242a1, determining the target transfer function;

[0061] Step 1242a2: Process the program angle in the guidance solution result using the target transfer function to obtain a first target booster stage swing angle.

[0062] Step 1242a3: performing angle limiting calculation on the first target booster stage angle to obtain a second target booster stage angle after limiting.

[0063] Step 1242a4: Determine a first target equivalent swing angle of the three channels of the booster stage based on the second target booster stage swing angle;

[0064] Step 1242a5: performing command allocation based on the first target equivalent swing angle to determine a first servo swing angle command;

[0065] Step 1242a6, based on the guidance timing corresponding to the core stage at the current moment, obtain the attitude control solution result for the booster stage; wherein, if the guidance timing is not the ignition moment, the first servo swing angle instruction is output as the attitude control solution result; otherwise, the second servo swing angle instruction corresponding to the core stage is output as the attitude control solution result.

[0066] In the embodiment of the present disclosure, when the target core stage is a booster stage, the launch vehicle preferably has four booster stages, each of which is equipped with a servo controller, and the four booster stages are evenly cross-strapped to the core stage. An attitude control solution is performed based on the pitch, yaw, and roll program angles (i.e., three-channel program angles) of the guidance module, as well as the timing of booster stage ignition, core stage ignition, and booster stage separation (i.e., separation timing), the average axial apparent acceleration of the inertial navigation module, and the average angular velocity of the rocket body:

[0067] Determine a target transfer function, which describes the dynamic characteristics of the swing angle control system, that is, is used to express the relationship between the input signal and the output signal. Specifically, the process of determining the target transfer function is as follows: step a11, calculating a first static amplification coefficient based on the guidance solution result and the navigation solution result; step a12, correcting the first static amplification coefficient through a digital correction network to obtain a second static amplification coefficient; step a13, determining the target transfer function based on the second static amplification coefficient; wherein the static method coefficient is preferably a fixed value used to adjust the sensitivity of the system, and the digital correction network is used to filter and correct the data to reduce errors and noise.

[0068] The program angle in the guidance solution is processed by the target transfer function to obtain the first target booster stage swing angle, and the first target booster stage swing angle is subjected to the swing angle limiting calculation to obtain the second target booster stage swing angle after limiting; the swing angle limiting calculation is used to avoid the attitude control instability caused by excessive swing of the booster and to limit the swing angle, that is, to ensure that the swing angle is within a safe range; according to the second target booster stage swing angle, the first target equivalent swing angle of the three channels of the booster stage is determined; the equivalent value is the swing angle value determined by comprehensively considering the information of each channel of the booster stage; based on the first target equivalent swing angle, the corresponding control instruction (first servo swing angle instruction) is generated, and then the first servo swing angle instruction is assigned to the booster. Different servo controllers corresponding to the booster stage are used to achieve precise attitude control; since the core stage needs to be ignited in advance in the booster stage, it is necessary to judge the guidance timing of the core stage. If the ignition time point is reached, the booster stage performs attitude control consistent with the core stage (the second servo swing angle instruction), and the stage outputs the servo control instruction corresponding to the core stage. Specifically, according to the guidance timing corresponding to the core stage at the current moment, the attitude control solution result of the booster stage is obtained; among them, if the guidance timing is not the ignition moment, the first servo swing angle instruction is output as the attitude control solution result, otherwise, the second servo swing angle instruction corresponding to the core stage is output as the attitude control solution result.

[0069] As shown in FIG3 , in a specific embodiment, the attitude control solution processing process for the booster stage and the core stage is as follows:

[0070] Step 31, determining whether the current stage is in the booster stage, that is, whether the target core stage to be controlled at the current moment is the booster stage;

[0071] Step 32: If yes, calculate the first static amplification factor based on the guidance solution result and the navigation solution result;

[0072] Step 33, correcting the first static amplification factor through a digital correction network to obtain a second static amplification factor;

[0073] Step 34: Determine a target transfer function based on the second static amplification factor, and based on the target transfer function, process the program angle in the booster stage guidance solution and perform angle limiting calculations to obtain a booster stage servo angle after limiting. This booster stage servo angle is then assigned a command to determine a booster stage servo command.

[0074] Step 35: Determine the guidance timing (i.e., whether the core stage, which is evenly cross-bound with the booster stage, is ignited) corresponding to the current moment.

[0075] Step 36: If the guidance timing is the ignition timing, the second servo swing angle instruction corresponding to the core level 1 is output as the attitude control solution result; otherwise, the first servo swing angle instruction is output as the attitude control solution result.

[0076] In an optional embodiment of the present disclosure, when the target core level is core level 1, core level 2, or core level 3, step 1242 includes:

[0077] Step 1242b1, determining the operating phase of the target core stage according to the separation timing in the guidance solution result;

[0078] Step 1242b2: Determine the attitude control solution result according to the operation stage.

[0079] In the embodiment of the present disclosure, when the target core stage is the core stage 1, the core stage 1 is divided into two operating stages during the operation of the carrier rocket: a powered flight stage (i.e., an operating stage from the completion of the separation of the booster stage to the exhaustion of the core stage 1 engine) and a core stage 1 glide stage (i.e., from the exhaustion of the core stage 1 engine to the completion of the separation of the core stage 1, and the separation of the fairing is also completed in this stage); when the target core stage 2 is the core stage 2, the core stage 2 is divided into three operating stages during the operation of the carrier rocket: a glide stage before the core stage 2 ignition (i.e., a stage from the completion of the separation of the core stage 1 to the ignition of the core stage 2), a core stage 2 powered stage (i.e., from the ignition of the core stage 2 to the exhaustion of the core stage 2), and a core stage 2 glide stage (i.e., from the exhaustion of the core stage 2 to the deflection of the core stage 2 nozzle); when the target core stage 3 is the core stage 3, the core stage 3 is divided into two operating stages during the operation of the carrier rocket: a glide stage before the core stage 3 ignition (i.e., from the completion of the separation of the core stage 2 to the ignition of the core stage 3 engine) and a core stage 3 powered flight stage (i.e., from the ignition of the core stage 3 engine to the exhaustion of the core stage 3 engine);

[0080] The operation phase of the core 1, core 2 or core 3 can be determined based on the separation timing in the guidance solution result, and the attitude control solution result is determined according to the operation phase;

[0081] As shown in FIG4 , in another specific embodiment, when the target core stage is core 1, the attitude control solution processing process for core 1 is as follows:

[0082] Step 41, determining whether the target core stage to be attitude controlled is core stage 1;

[0083] Step 42: If yes, determine whether the core stage is in a powered flight phase.

[0084] Step 431 , if the core stage is in the powered flight phase, then calculate the third static amplification factor of the core stage;

[0085] Step 432, correcting the third static amplification factor through a digital correction network to obtain a fourth static amplification factor;

[0086] Step 433: Determine a first transfer function of the first core stage based on the fourth static amplification factor. Based on the first transfer function, process the program angle in the guidance solution of the first core stage and perform swing angle limiting calculation to obtain a limited first core stage servo swing angle. Determine a second target equivalent swing angle for the three channels of the first core stage based on the first core stage servo swing angle. Allocate a command for the second target equivalent swing angle to determine a first core stage servo command.

[0087] Step 441: If the core stage 1 is in the core stage 1 glide phase, the program angle in the current core stage 1 guidance solution is input into the digital correction network for correction processing to obtain a corrected program angle.

[0088] Step 442 , the corrected program angle is processed through a three-channel threshold control to limit the amplitude of the corrected program angle to a certain range to prevent instability caused by excessive control instructions;

[0089] In step 443, the corrected program angle after threshold control processing is assigned to the servo controller of the core stage, and corresponding control instructions are output. These instructions are used to actually control the attitude of the launch vehicle to maintain a stable gliding state of the core stage.

[0090] As shown in FIG5 , in another specific embodiment, when the target core stage is the second core stage, the attitude control solution processing process for the second core stage is:

[0091] Step 51, determining whether the target core stage to be attitude controlled is the second core stage;

[0092] Step 52: If yes, determine the operating stage of the core secondary;

[0093] Step 531: If the core stage 2 is in the coasting phase before core stage 2 ignition, the program angle in the guidance solution result of the current coasting phase before core stage 2 ignition is input into the digital correction network for correction processing to obtain a corrected program angle.

[0094] Step 532: Process the corrected program angle through a three-channel threshold control to limit the amplitude of the corrected program angle to a certain range to prevent excessive control instructions from causing instability in the core secondary stage.

[0095] In step 533, the corrected program angle after threshold control is assigned to the servo controller for the glide phase before the core stage ignition, and the controller outputs corresponding control instructions. These instructions are used to actually control the attitude of the launch vehicle to maintain a stable glide state during the glide phase before the core stage ignition.

[0096] Step 541 , if the core stage 2 is in the core stage 2 power stage, then calculate the fifth static amplification factor of the core stage 2;

[0097] Step 542, correcting the fifth static amplification factor through a digital correction network to obtain a sixth static amplification factor;

[0098] Step 543: Determine a second transfer function of the core-stage 2 based on the sixth static amplification factor. Based on the second transfer function, process the program angle in the guidance solution of the core-stage 2 and perform swing angle limiting calculation to obtain a limited core-stage 2 servo swing angle. Determine a third target equivalent swing angle for the three channels of the core-stage 2 based on the core-stage 2 servo swing angle. Commands are assigned to the third target equivalent swing angle to determine a core-stage 2 servo command.

[0099] Step 551: If the core-stage 2 is in the core-stage 2 glide phase, the program angle in the guidance solution of the current core-stage 2 glide phase is input into the digital correction network for correction processing to obtain a corrected program angle.

[0100] Step 552: Process the corrected program angle through a three-channel threshold control to limit the amplitude of the corrected program angle to a certain range to prevent excessive control instructions from causing instability in the core secondary stage.

[0101] In step 553, the corrected program angle after threshold control processing is assigned to the servo controller of the core second stage glide segment, and the corresponding control instructions are output. These instructions are used to actually control the attitude of the launch vehicle to maintain a stable glide state of the core second stage glide segment.

[0102] As shown in FIG6 , in another specific embodiment, when the target core stage is the core three stage, the attitude control solution processing process for the core three stage is:

[0103] Step 61, determining whether the target core stage to be attitude controlled is the core three stage;

[0104] Step 62: If yes, determine whether the core stage 3 is in the coasting phase before core stage 3 ignition.

[0105] Step 631: If yes, then the program angle in the guidance solution result of the current core three-stage glide segment before ignition is input into the digital correction network for correction processing to obtain a corrected program angle;

[0106] Step 632: The corrected program angle is processed through the three-channel threshold control of the glide segment to limit the amplitude of the corrected program angle to a certain range to prevent excessive control instructions from causing instability in the core three stages.

[0107] In step 633, the corrected program angle after threshold control is assigned to the servo controller for the glide phase before the core stage ignition, and the controller outputs corresponding control instructions. These instructions are used to actually control the attitude of the launch vehicle to maintain a stable glide state during the glide phase before the core stage ignition.

[0108] Step 641, if not, calculate the seventh static amplification factor of the core three stage;

[0109] Step 642, correcting the seventh static amplification factor through a digital correction network to obtain an eighth static amplification factor;

[0110] Step 643, determine the third transfer function of the core-three-stage powered flight segment according to the eighth static amplification coefficient, and based on the third transfer function, process and calculate the swing angle limit according to the program angle in the guidance solution result of the core-three-stage powered flight segment to obtain the core-three-stage servo swing angle after limit, and determine the fourth target equivalent swing angle of the core-three-stage three-channel according to the core-three-stage servo swing angle, assign instructions to the fourth target equivalent swing angle, and determine the servo instruction of the core-three-stage powered flight segment.

[0111] In an optional embodiment of the present disclosure, when the target core stage is the final stage, step 1242 includes:

[0112] Step 1242c1, processing the program angle in the guidance solution result through a digital correction network to obtain a correction value;

[0113] Step 1242c2: perform threshold control processing on the correction value, determine the fusion communication control instruction, and output the fusion communication control instruction as the attitude control solution result.

[0114] In the embodiment of the present disclosure, when the target core stage is the final stage, the final stage is divided into three operating phases during the operation of the launch vehicle: the terminal velocity correction phase (i.e., from the exhaustion of the core three-stage engine to the end of the terminal velocity correction), the satellite-rocket separation phase (i.e., from the end of the terminal velocity correction to the end of the satellite-rocket separation), and the deorbit passivation phase (i.e., from the end of the satellite-rocket separation to the completion of the final stage passivation).

[0115] As shown in Figure 7, when the target core stage is the last stage, the attitude control solution processing process for the last stage is:

[0116] Step 71, determining whether the target core stage to be attitude controlled is the final stage;

[0117] Step 72: If yes, then the program angle in the current final-stage guidance solution is input into the digital correction network for correction processing to obtain a corrected program angle (i.e., correction value);

[0118] Step 73: The corrected program angle (i.e., correction value) is processed through a three-channel threshold control to limit the amplitude of the corrected program angle (i.e., correction value) to a certain range while meeting the preset control requirements of the final stage to prevent instability of the final stage caused by excessive control instructions. The preset control requirements can be adjusted according to actual conditions.

[0119] In step 74, the corrected program angle after threshold control processing is assigned to the servo controller of the final stage, and a corresponding RCS control instruction is output. The RCS control instruction is used to actually control the attitude of the launch vehicle to maintain a stable gliding state of the final stage.

[0120] The aforementioned launch vehicle attitude control simulation method using the flight control module addresses challenges such as the difficulty in simulating real-world environments, limited intuitive verification methods, and high safety risks during launch vehicle testing. Furthermore, in a specific implementation example, experimental simulations demonstrated that a Unity3D rocket model could visually display changes in the rocket's pitch, yaw, and roll angles under complex flight conditions. The servo model visually displayed the swing angle of the servo actuator and the changes in the servo flame. The RCS model visually displayed the attitude control nozzle switch status. The six-degree-of-freedom model outputs information such as longitude, latitude, and altitude required by the inertial control unit (IMU). The closed-loop operation of the entire system quickly and effectively verified the launch vehicle's flight control algorithm.

[0121] The embodiment of the present disclosure receives the full pulse quantity sent by the laser strapdown inertial system; performs calculation processing based on the full pulse quantity to obtain a target calculation result; the target calculation result includes a navigation calculation result, a guidance calculation result, and an attitude control calculation result; and sends the navigation calculation result, the guidance calculation result, and the attitude control solution result to a simulation server, thereby solving the problems of difficulty in simulating a real environment during launch vehicle testing, a less intuitive verification method, and high safety risks, and has the following beneficial effects:

[0122] 1) High-precision attitude control simulation and prediction: Using digital twin technology to generate a rocket attitude control model, it has high-precision simulation and prediction capabilities, and can more accurately predict the rocket's attitude response and control effect under different working conditions.

[0123] 2) Save time and cost. Digital twin technology can be used to perform simulation and optimization on computers without the need for experimental verification, saving time and cost. At the same time, digital twin technology can generate attitude control models of rockets under different working conditions, further reducing the cost of experimental testing.

[0124] 3) High practicality. The rocket attitude control simulation method based on digital twin can be widely used in the field of rocket attitude control, such as rocket design and optimization, rocket testing and verification, and rocket real-time control.

[0125] As shown in FIG8 , an embodiment of the present disclosure provides a method for simulating attitude control of a launch vehicle, which is applied to a simulation server and includes:

[0126] Step 81, receiving a start signal sent by the simulation client;

[0127] Step 82: Send a master command signal to the flight control module according to the start signal;

[0128] Step 83, sending simulated position information to the laser strapdown inertial system, so that the laser strapdown inertial system sends a full pulse quantity calculated according to the simulated position information to the flight control module;

[0129] Step 84, receiving the navigation solution result, guidance solution result, and attitude control solution result sent by the flight control module;

[0130] Step 85, performing model solution processing according to the navigation solution result, the guidance solution result, and the attitude control solution result to obtain a model solution result;

[0131] Step 86 : Send the model solution result to the simulation client and the laser strapdown inertial system respectively, so that the simulation client can perform three-dimensional display; the model solution result includes the latest simulated position information.

[0132] In an embodiment of the present disclosure, a simulation client sends a start signal to a simulation server. After receiving the start signal, the simulation server sends an IO master command signal to the flight control module, causing the flight control module to enter the flight control phase and begin flight control solution. The simulation server sends longitude, latitude, altitude, angular velocity, and acceleration information (i.e., simulated position information) to the laser strapdown inertial system via a 1553B bus. The laser strapdown inertial system calculates a full pulse based on the simulated position information and sends the full pulse to the flight control module, so that the flight control module on the rocket can perform solution processing based on the full pulse to obtain a target solution result. The target solution result (navigation solution result, guidance solution result, and attitude control solution result) of the flight control module is returned to the simulation server. The target solution result includes: servo command information, RCS attitude control switch, timing control signal, attitude information, etc. The simulation server performs model solution processing on the target solution result to obtain a model solution result. The model solution processing here includes rocket model solution, environment model solution, engine model solution, and six-degree-of-freedom simulation calculation model, etc.

[0133] Specifically, the model solution results obtained by solving the rocket model include: gravity, aerodynamic force, aerodynamic torque Mach number and other information; the model solution results obtained by solving the environment model include: dynamic pressure, wind speed and other information; the model solution results obtained by solving the engine model include: thrust and mass and other information; the model solution results obtained by the six-degree-of-freedom simulation calculation model include: longitude and latitude, angular velocity and angular acceleration and other information; the simulation client receives the model solution results and flight control module data information calculated by the simulation server for three-dimensional display. At the same time, the simulation server sends the calculated model solution results to the laser strapdown inertial system through the 1553B bus, forming a closed-loop operation of the data flow; it is worth noting that the laser strapdown inertial system is preferably a physical laser strapdown inertial system, that is, semi-physical simulation is performed, of course, it can also be a simulated laser strapdown inertial system, that is, full-physical simulation is performed, and this application is not limited to this.

[0134] As shown in Figures 9 and 10, in a specific embodiment, Unity software is used to build a three-dimensional rocket body model of the launch vehicle, and the launch vehicle is divided into five stages: booster stage, core stage 1, core stage 2, core stage 3 and final stage, to control the stable flight of the rocket in each stage.

[0135] The attitude control simulation system of the launch vehicle includes a flight control module, a simulation server, a laser strapdown inertial system and a simulation client. The simulation client sends a start instruction to the simulation server via Ethernet to indicate the start of the simulation experiment. The simulation server includes a three-dimensional rocket model obtained by Unity3D, which includes a nozzle model, a body model and a servo model. The simulation server sends simulated position information (including longitude, latitude, altitude, angular velocity, acceleration, etc.) to the laser strapdown inertial system through the 1553B bus. The laser strapdown inertial system calculates 12 full pulses based on the simulated position information and sends the full pulses, apparent velocity increment and angular increment to the flight control module. The flight control module consists of an inertial navigation module, a guidance module and an attitude control module. The inertial module outputs angular velocity, axial apparent acceleration, attitude angle, altitude, average apparent acceleration, etc. The guidance module outputs timing instructions (separation timing) and program angle, etc. The attitude control module outputs position information, angle information, servo nozzle swing information and RCS control instructions (target solution results); before the stop timing is reached, the target solution results are returned to the simulation server through Ethernet, and the launch vehicle model of the simulation server updates the flight status according to the input information; the simulation client receives the information solved by the simulation server model and the data information output by the flight control module for three-dimensional display. At the same time, the simulation server sends the solved data information to the laser strapdown inertial group through the 1553B bus, forming a closed-loop operation of the data flow.

[0136] The above-mentioned carrier rocket attitude control simulation method can control the servo actuator to swing to the corresponding position under extreme pulling conditions to achieve stable flight of the carrier rocket. At the same time, a three-dimensional simulation model consistent with the carrier rocket is proposed for the attitude control system. Through the Unity engine, when the rocket attitude information changes, the model can make corresponding attitude and position changes in real time; it should be noted that the method of the simulation server is a method corresponding to the method of the above-mentioned flight control module. All implementation methods in the method embodiments of the above-mentioned flight control module are applicable to the embodiments of the method of the simulation server, and can also achieve the same technical effects.

[0137] An embodiment of the present disclosure further provides a launch vehicle attitude control simulation system, comprising:

[0138] A flight control module; a simulation client electrically connected to the flight control module; a simulation server communicatively connected to the simulation client and electrically connected to the flight control module; a laser strapdown inertial system (LSINS) unit (LSU) communicatively connected to the simulation server and the flight control module; wherein the flight control module receives a full pulse quantity sent by the LSI; the flight control module performs a solution processing based on the full pulse quantity to obtain a target solution result; the target solution result includes a navigation solution result, a guidance solution result, and an attitude control solution result; and the flight control module sends the navigation solution result, the guidance solution result, and the attitude control solution result to the simulation server.

[0139] Optionally, the flight control module includes:

[0140] An inertial navigation module, the inertial navigation module is communicatively connected to the laser strapdown inertial group; a guidance module, the guidance module is communicatively connected to the inertial navigation module; and an attitude control module, the attitude control module is communicatively connected to the inertial navigation module and the guidance module respectively; wherein the inertial navigation module is used to perform navigation solution processing on the full pulse amount to obtain a navigation solution result; the navigation solution result includes at least one of an attitude angle, an altitude, a relative velocity inclination angle, an apparent acceleration mean, and an angular velocity mean; the guidance module is used to perform guidance solution processing on the navigation solution result to obtain a guidance solution result; the guidance solution result includes a separation timing and a program angle; the attitude control module is used to perform attitude control solution processing based on the guidance solution result and the navigation solution result to obtain an attitude control solution result.

[0141] Optionally, the simulation server is connected to the simulation client via Ethernet communication; the laser strapdown inertial group is connected to the simulation server and the flight control module via a 1553B bus.

[0142] It should be noted that the system is a system corresponding to the above-mentioned method, and all implementation methods in the above-mentioned method embodiments are applicable to the embodiments of the system and can achieve the same technical effects.

[0143] Those skilled in the art will appreciate that the units and algorithm steps of each example described in conjunction with the embodiments disclosed herein can be implemented in electronic hardware, or a combination of computer software and electronic hardware. Whether these functions are performed in hardware or software depends on the specific application and design constraints of the technical solution. Professionals and technicians can use different methods to implement the described functions for each specific application, but such implementation should not be considered beyond the scope of this disclosure.

[0144] Those skilled in the art will clearly understand that, for the convenience and brevity of description, the specific working processes of the systems, devices and units described above can refer to the corresponding processes in the aforementioned method embodiments and will not be repeated here.

[0145] In the embodiments provided in the present disclosure, it should be understood that the disclosed devices and methods can be implemented in other ways. For example, the device embodiments described above are merely schematic. For example, the division of the units is merely a logical function division. In actual implementation, there may be other division methods, such as multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the mutual coupling or direct coupling or communication connection shown or discussed can be through some interfaces, indirect coupling or communication connection of devices or units, which can be electrical, mechanical or other forms.

[0146] The units described as separate components may or may not be physically separate, and the components shown as units may or may not be physical units, that is, they may be located in one place or distributed across multiple network units. Some or all of these units may be selected to achieve the purpose of this embodiment according to actual needs.

[0147] In addition, each functional unit in each embodiment of the present disclosure may be integrated into one processing unit, or each unit may exist physically separately, or two or more units may be integrated into one unit.

[0148] If the functions are implemented in the form of software functional units and sold or used as independent products, they can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present disclosure, or the part that contributes to the prior art, or the part of the technical solution, can be embodied in the form of a software product. The computer software product is stored in a storage medium and includes several instructions for enabling a computer device (which can be a personal computer, server, or network device, etc.) to execute all or part of the steps of the method described in each embodiment of the present disclosure. The aforementioned storage medium includes various media that can store program codes, such as a USB flash drive, a mobile hard disk, a ROM, a RAM, a magnetic disk, or an optical disk.

[0149] In addition, it should be pointed out that in the apparatus and method of the present invention, it is obvious that each component or each step can be decomposed and / or recombined. These decompositions and / or recombinations should be regarded as equivalent schemes of the present invention. Moreover, the steps of performing the above-mentioned series of processing can naturally be performed in chronological order according to the order of description, but it is not necessary to perform them in chronological order, and some steps can be performed in parallel or independently of each other. For those of ordinary skill in the art, it is understandable that all or any steps or components of the method and apparatus of the present invention can be implemented in hardware, firmware, software or a combination thereof in any computing device (including a processor, storage medium, etc.) or a network of computing devices, which can be achieved by those of ordinary skill in the art using their basic programming skills after reading the description of the present invention.

[0150] Therefore, the purpose of the present disclosure can also be achieved by running a program or a group of programs on any computing device. The computing device can be a well-known general-purpose device. Therefore, the purpose of the present disclosure can also be achieved simply by providing a program product containing program code that implements the method or device. That is to say, such a program product also constitutes the present disclosure, and the storage medium storing such a program product also constitutes the present disclosure. Obviously, the storage medium can be any well-known storage medium or any storage medium developed in the future. It should also be pointed out that in the device and method of the present disclosure, it is obvious that each component or each step can be decomposed and / or recombined. These decompositions and / or recombinations should be regarded as equivalent schemes of the present disclosure. In addition, the steps of performing the above-mentioned series of processing can naturally be performed in chronological order according to the order of description, but do not necessarily need to be performed in chronological order. Certain steps can be performed in parallel or independently of each other.

[0151] The above is only a preferred embodiment of the present disclosure. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present disclosure. These improvements and modifications should also be considered within the scope of protection of the present disclosure.

Claims

1. A launch vehicle attitude control simulation method, characterized in that: Applied to the flight control module, it includes: receiving the full amount of pulses sent by the laser strapdown inertial group; performing solution processing according to the full amount of pulses to obtain a target solution result; the target solution result includes a navigation solution result, a guidance solution result and an attitude control solution result; sending the navigation solution result, the guidance solution result and the attitude control solution result to the simulation server.

2. The method for simulating the attitude control of a launch vehicle according to claim 1, characterized in that: A solution is performed based on the full pulse to obtain a target solution result, including: performing tool error compensation on the full pulse to obtain apparent velocity increments and angular increments in the first direction, the second direction and the third direction; performing navigation solution processing on the apparent velocity increments and angular increments to obtain a navigation solution result; the navigation solution result includes at least one of an attitude angle, an altitude, a relative velocity inclination, an apparent acceleration mean and an angular velocity mean; performing guidance solution processing on the navigation solution result to obtain a guidance solution result; the guidance solution result includes a separation timing and a program angle; performing attitude control solution processing based on the guidance solution result and the navigation solution result to obtain an attitude control solution result.

3. The method for simulating the attitude control of a launch vehicle according to claim 2, characterized in that: Performing attitude control solution processing according to the guidance solution result and the navigation solution result to obtain an attitude control solution result, including: determining a target core stage to be attitude controlled at the current moment; the target core stage includes at least one of a booster stage, a core stage 1, a core stage 2, a core stage 3 and a final stage; Performing attitude control solution processing on the target core stage according to the guidance solution result and the navigation solution result to obtain an attitude control solution result.

4. The method for simulating the attitude control of a launch vehicle according to claim 3, characterized in that: When the target core stage is a booster stage and a core stage 1, an attitude control solution is performed on the target core stage according to the guidance solution result and the navigation solution result to obtain an attitude control solution result, including: determining a target transfer function; processing the program angle in the guidance solution result by means of the target transfer function to obtain a first target booster stage swing angle; performing a swing angle limiting calculation on the first target booster stage swing angle to obtain a second target booster stage swing angle after limiting; determining a first target equivalent swing angle of the three channels of the booster stage according to the second target booster stage swing angle; performing instruction allocation based on the first target equivalent swing angle to determine a first servo swing angle instruction; obtaining an attitude control solution result for the booster stage according to the guidance timing corresponding to the core stage 1 at the current moment; wherein, if the guidance timing is not the ignition moment, the first servo swing angle instruction is output as the attitude control solution result, otherwise, the second servo swing angle instruction corresponding to the core stage 1 is output as the attitude control solution result.

5. The method for simulating the attitude control of a launch vehicle according to claim 3, characterized in that: When the target core stage is core one, core two or core three, attitude control solution processing is performed according to the guidance solution result and the navigation solution result to obtain an attitude control solution result, including: determining the operating stage of the target core stage according to the separation timing in the guidance solution result; determining the attitude control solution result according to the operating stage.

6. The method for simulating the attitude control of a launch vehicle according to claim 3, characterized in that: When the target core stage is the last stage, attitude control solution processing is performed according to the guidance solution result and the navigation solution result to obtain an attitude control solution result, including: processing the program angle in the guidance solution result through a digital correction network to obtain a correction value; performing threshold control processing on the correction value to determine a fused communication control instruction and outputting the fused communication control instruction as an attitude control solution result.

7. A launch vehicle attitude control simulation method, characterized in that: Applied to a simulation server, the system includes: receiving a start signal sent by a simulation client; sending a main command signal to a flight control module according to the start signal; sending simulated position information to a laser strapdown inertial group so that the laser strapdown inertial group sends a full pulse amount obtained by solving the simulated position information to the flight control module; receiving navigation solution results, guidance solution results and attitude control solution results sent by the flight control module; performing model solution processing according to the navigation solution results, guidance solution results and attitude control solution results to obtain a model solution result; sending the model solution result to the simulation client and the laser strapdown inertial group respectively, so that the simulation client can perform a three-dimensional display; the model solution result includes the latest simulated position information.

8. A launch vehicle attitude control simulation system, characterized in that: include: Flight control module; A simulation client, the simulation client is electrically connected to the flight control module; a simulation server, the simulation server is communicatively connected to the simulation client and electrically connected to the flight control module; a laser strapdown inertial system group, the laser strapdown inertial system group is communicatively connected to the simulation server and communicatively connected to the flight control module; wherein the flight control module receives the full pulse amount sent by the laser strapdown inertial system group; the flight control module performs solution processing according to the full pulse amount to obtain a target solution result; the target solution result includes a navigation solution result, a guidance solution result and an attitude control solution result; the flight control module sends the navigation solution result, the guidance solution result and the attitude control solution result to the simulation server.

9. The launch vehicle attitude control simulation system according to claim 8, characterized in that: The flight control module includes: an inertial navigation module, which is communicatively connected to the laser strapdown inertial group; a guidance module, which is communicatively connected to the inertial navigation module; an attitude control module, which is communicatively connected to the inertial navigation module and the guidance module respectively; wherein the inertial navigation module is used to perform navigation solution processing on the full pulse amount to obtain a navigation solution result; the navigation solution result includes at least one of attitude angle, altitude, relative velocity inclination, apparent acceleration mean and angular velocity mean; the guidance module is used to perform guidance solution processing on the navigation solution result to obtain a guidance solution result; the guidance solution result includes a separation timing and a program angle; the attitude control module is used to perform attitude control solution processing according to the guidance solution result and the navigation solution result to obtain an attitude control solution result.

10. The launch vehicle attitude control simulation system according to claim 8, characterized in that: The simulation server is connected to the simulation client via Ethernet communication; the laser strapdown inertial unit is connected to the simulation server and the flight control module via a communication bus.

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