A Mission-Decommissioning Design Method for Launch Vehicles

Abstract

The present invention relates to a method for de-missionizing the design of a launch vehicle. Starting from the problems encountered in the contract fulfillment design process of the launch vehicle, the present invention proposes, from an overall perspective, the idea of considering de-missionizing design during the development of the launch vehicle, that is, de-missionizing design methods are given in aspects such as the trajectory calculation of the launch vehicle, the design of the aerodynamic heat environment, the load environment, the space heat environment, the design of the satellite-vehicle separation scheme, the design of the measurement and control scheme, the design of the guidance system scheme, the design of the attitude control scheme, and the design of the flight control software, etc., so as to minimize the customized design work for different launch missions during the contract fulfillment design process, shorten the development cycle, reduce the launch cost, and meet the de-missionizing requirements during the contract fulfillment process of the launch vehicle.

Description

Technical Field

[0001] The present invention relates to the technical field of launch vehicles, and particularly relates to a method for de-missionizing the design of launch vehicle launches. Background Art

[0002] During the overall design process of a launch vehicle, it is impossible to take into account all launch missions. For different launch missions, due to changes in the mass and quantity of satellites and the requirements of the satellite party, the takeoff mass of each stage of the launch mission, the ballistic trajectory, the TT&C (Telemetry, Tracking and Command) plan, the satellite-rocket separation plan, the control stability plan, and the flight control software, etc. will all change. A large amount of customized design work needs to be carried out for each system of the launch vehicle. Therefore, in the process of commercial launches, the design and launch cycles are long, and a large amount of human and material resources are required, making it difficult to meet the current requirements for rapid launches and high-density launches. In order to adapt to different launch missions, reduce the requirements for human and material resources during the commercial launch stage of launch vehicles, shorten the development cycle, reduce the launch cost, meet the requirements for rapid launches and high-density launches, and enhance the market competitiveness of products, it is necessary to consider the idea of de-missionizing design during the development process of launch vehicles. Summary of the Invention

[0003] To solve the above technical problems, the present invention proposes a method for de-missionizing the design of launch vehicles, which can quickly give a reasonable launch plan for different launch missions and realize the de-missionizing design requirements of launch vehicles.

[0004] The technical solution provided by the present invention is specifically as follows:

[0005] A method for de-missionizing the design of launch vehicle launches, comprising

[0006] Selecting a corresponding landing area based on the characteristics of the target orbit, and then designing and optimizing the ballistic trajectory according to this landing area to make the landing area meet the safety requirements, and pre-designing a ballistic calculation input interface according to the existing target orbit parameters;

[0007] Performing ballistic envelope calculations according to existing and future possible launch missions; on this basis, selecting the most severe working conditions of the aerodynamic heat environment, the load environment, and the space heat environment, and respectively carrying out the design of the aerodynamic heat protection plan, the strength design, and the space heat protection plan;

[0008] Statistically analyzing the space dimensions, installation interfaces, masses, and separation plans of existing satellites, and designing satellite adapters of different standards;

[0009] Automatically planning a ground-based TT&C station, a space-based TT&C station, and a rocket roll angle adjustment plan that meet the TT&C requirements according to the position and attitude of the rocket and the position and data transmission capacity information of the TT&C station;

[0010] According to different orbit injection methods and different satellite-rocket separation schemes, conduct envelope design for the guidance system; after the ballistic calculation is completed, generate the control timing sequence of the guidance system;

[0011] Design attitude control schemes for each stage:

[0012] For the first-stage flight segment, analyze the dynamic characteristic envelope of the first-stage rocket body under different launch points, correct the network gain scheduling according to the launch point altitude, and use a fixed filter of the envelope elastic mode for attitude control;

[0013] For the second-stage flight segment, optimize the structure of the second-stage correction network, broaden the notch range of the filter, and use fixed correction network parameters for attitude control;

[0014] For the last-stage flight segment, according to the different influences of different satellite shape and quality parameters on the control ability, pre-design the corresponding gain scheduling table, and generate the correction network parameters according to the satellite shape and quality parameters before launch;

[0015] Design the flight control software. The flight control software includes general modules and special modules. Design the software architecture to implement the control logic and call relationship between the general modules and special modules, and realize the adaptability to different tasks by reserving configurable interfaces.

[0016] Furthermore, the target orbit parameters include target orbit altitude, target orbit inclination, local time of descending node / geographical longitude of ascending node, eccentricity, argument of perigee.

[0017] Furthermore, the ballistic envelope calculation conducts ballistic calculations for working conditions of different launch sites, different orbit altitudes, and different combinations of orbit inclinations.

[0018] Furthermore, the automatic planning of ground-based measurement and control stations, space-based measurement and control stations, and rocket roll angle adjustment schemes that meet the measurement and control requirements according to the rocket position and attitude, as well as the position and data transmission capacity information of the measurement and control stations, includes:

[0019] Determine ground-based and space-based measurement and control stations that meet the measurement and control needs according to the data transmission requirements of the mission, and then calculate the roll angle that meets the measurement and control requirements in the current attitude based on the beam angle constraints of the measurement and control antennas according to the current position, attitude of the rocket, and the position of the measurement and control stations.

[0020] Furthermore, for the first-stage flight segment, the attitude control scheme for the de-tasked scheme of correcting network gain scheduling according to the launch point altitude and the fixed filter of the envelope elastic mode includes:

[0021] Before launch, generate a corresponding gain interpolation table that changes with time according to the altitude of the launch point as the attitude control system parameters during flight;

[0022] The design method of the fixed filter for the envelope elastic mode is as follows: First, determine the envelope range of the frequency, and then design the filter according to the envelope range of the frequency to achieve the task-independent design of the filter.

[0023] Furthermore, the shape and property parameters of the satellite include the mass, center of mass, and moment of inertia data of the satellite.

[0024] Furthermore, the general module includes a navigation calculation module, an attitude angle calculation module, a stage transition control module, and a bus data processing module; the special module includes an attitude control module and a guidance control module.

[0025] Furthermore, the control logic and call relationship between the general module and the special module implemented by the designed software architecture include:

[0026] The flight control software sequentially calls the general module or the special module corresponding to the flight timing according to the flight timing; among them, before calling the special module, first perform a logical judgment according to the flight mission, and then call the special module matching the flight mission.

[0027] Furthermore, the flight control software includes a mission planning system. The mission planning system performs calculations and data processing. After inputting the input parameters required for the calculations to the mission planning system, the mission planning system outputs relevant configuration parameters;

[0028] The special module uses the configuration parameters generated by the mission planning system for dynamic combination according to different missions.

[0029] Furthermore, the input parameters include the target orbit, launch point, launch time, and orbit injection method;

[0030] The configuration parameters include the weights of each stage of the rocket, statistical wind field data, attitude control consumption of each stage, and flight timing.

[0031] Compared with the prior art, the present invention has the following beneficial effects:

[0032] Starting from the problems encountered in the performance fulfillment design process of the launch vehicle, the present invention proposes the idea of task-independent design that needs to be considered in the development process of the launch vehicle from an overall perspective, that is, in aspects such as the ballistic calculation of the launch vehicle, the design of the aerodynamic heat environment, the load environment, the space heat environment scheme, the satellite-rocket separation scheme, the measurement and control scheme, the guidance system scheme, the attitude control scheme, and the flight control software design, etc., gives the task-independent design method, as much as possible reducing the customized design work for different launch missions in the performance fulfillment design process, shortening the development cycle, reducing the launch cost, and meeting the task-independent requirements in the performance fulfillment process of the launch vehicle. Specific Embodiments

[0033] In view of the many defects in the existing design process of launch vehicles, the present invention provides a method for de-tasked design of launch vehicles. During the early design process of launch vehicles, the de-tasked design concept is adopted for various design tasks associated with the mission, meeting the de-tasked requirements during the performance of the launch vehicle contract. The following details the specific implementation methods of the present invention.

[0034] A method for de-tasked design of launch vehicle launches, comprising the steps

[0035] S1. Since the target orbit requirements, orbit insertion methods, and satellite-rocket separation schemes corresponding to different launch missions are different, the required ballistic calculation methods are also different.

[0036] According to the de-tasked requirements, the landing areas of each stage of the separated body under different mission conditions (at the four major launch sites and under different target orbit inclinations) are known before the rocket launch. Based on the characteristics of the target orbit, the corresponding landing area is selected, and then the trajectory is designed and optimized according to this landing area to ensure the safety requirements of the landing area, avoiding the process of landing area survey before each launch mission. According to the existing target orbit parameters, the input interface for ballistic calculation is pre-designed to meet the needs of different satellite parties. In addition, different orbit insertion methods can be selected according to manual selection or automatic interpretation methods. The satellite-rocket separation scheme is designed according to the installation method of the satellite, the far-field safety between the upper stage and the satellite, and between satellites to meet the de-tasked requirements.

[0037] Target orbit parameters: including target orbit altitude, target orbit inclination, local time of descending node / geographical longitude of ascending node, eccentricity, argument of perigee.

[0038] Orbit insertion method: According to the number of ignition times of the upper stage of the launch vehicle, it is divided into single ignition of the upper stage for orbit insertion and two ignitions of the upper stage for orbit insertion. Generally, for lower orbit altitudes (below 500 km), the single ignition of the upper stage of the launch vehicle for orbit insertion is selected, which can enter the target orbit faster. When the orbit altitude is higher, the two ignitions of the upper stage of the launch vehicle for orbit insertion are selected. The rocket flight time is longer, but the carrying capacity is higher.

[0039] The satellite-rocket separation scheme refers to the way the satellite separates. The satellite is pre-installed on the satellite adapter of the upper stage. According to the spatial size of the satellite, the satellite is generally arranged on the side or axial direction of the satellite adapter, resulting in different satellite separation methods: when the satellite is arranged on the side of the satellite adapter, it will separate laterally along the upper stage; when the satellite is arranged axially on the satellite adapter, it will separate axially along the upper stage.

[0040] Far-field safety: After the satellite separates from the upper stage, it is necessary to ensure the safety of the upper stage and the satellite during the subsequent flight process, preventing them from being too close or colliding.

[0041] S2. Due to the differences in the takeoff mass, flight trajectory, and flight duration of launch vehicles corresponding to different launch missions, the aerodynamic heat environment, load environment, and space heat environment during the flight of launch vehicles are all different. Based on market research, the existing and future possible launch missions are statistically analyzed, and then the ballistic envelopes are calculated for all possible launch missions; the ballistic envelope calculation is to perform ballistic calculations for different launch sites (the four major domestic launch sites: Jiuquan, Taiyuan, Xichang, Wenchang), different orbital altitudes, and different orbital inclination combinations.

[0042] On this basis, the working conditions with the most severe aerodynamic heat environment, load environment, and space heat environment are selected, and the aerodynamic heat protection scheme design, strength design, and space heat protection scheme design are carried out respectively; thus ensuring that all launch mission working conditions are considered in the design stage of the launch vehicle.

[0043] S3. Due to the different launch requirements of different satellite parties, the designs of the number of satellites, space dimensions, mass, installation interfaces, etc. are different, resulting in different satellite installation and separation schemes. The space dimensions, installation interfaces, mass, separation schemes, etc. of existing satellites are statistically analyzed, and satellite adapters with different standards are designed to form a mission manual for satellite parties to select and design satellite installation interfaces, so as to meet the demand of de-tasking.

[0044] A satellite adapter refers to a satellite fixing and separation mechanism that can fix the satellite on the launch vehicle and safely separate the satellite from the launch vehicle under the control of a separation command.

[0045] S4. Due to the different flight trajectories of launch vehicles corresponding to different launch missions, and the positions of ground-based and space-based measurement and control stations are generally fixed, the resulting measurement and control schemes will also be different.

[0046] Therefore, according to different launch missions, during the process of ballistic calculation, a measurement and control scheme for ground-based measurement and control stations, space-based measurement and control stations, and the roll angle of the rocket that meets the measurement and control requirements can be automatically planned based on information such as the position and attitude of the rocket, the position of the measurement and control stations, and the data transmission capacity.

[0047] The amount of data to be transmitted and the flight trajectory of the rocket are different under different missions, and the ground-based and space-based measurement and control stations that can meet different flight trajectories are different; therefore, it is necessary to first determine the ground-based and space-based measurement and control stations that meet the measurement and control requirements according to the data transmission requirements of the mission, and then calculate the roll angle that meets the measurement and control requirements at the current attitude based on the current position and attitude of the rocket and the position of the measurement and control stations, subject to the beam angle constraint of the measurement and control antenna.

[0048] Rocket roll angle adjustment scheme: Since the ground-based antenna and space-based antenna are generally located on the side of the carrier rocket body and have fixed positions, the antenna can only point to the ground-based measurement and control station or space-based measurement and control station by adjusting the rocket body attitude to achieve data transmission. Since adjusting the roll angle does not affect the flight trajectory of the carrier rocket, the measurement and control scheme is generally optimized by adjusting the roll angle.

[0049] S5. Due to different launch missions, different flight timings, different rocket orbit insertion methods, different satellite-rocket separation methods, the design of the guidance system is also different. Therefore, envelope design is carried out during the design process of the guidance system, considering the adaptability to different orbit insertion methods and different satellite-rocket separation schemes. At the same time, after the ballistic calculation is completed, the control timing of the guidance system is generated to meet the de-tasking requirements.

[0050] S6. Due to different launch missions, different launch sites, different launch times, different takeoff masses of each stage, different ballistic trajectories, and different roll angles of the carrier rocket caused by the change of the measurement and control scheme, the attitude control methods of each stage are different.

[0051] For the first-stage flight segment: By analyzing the envelope of the dynamic characteristics (static instability degree, rudder effectiveness) of the first-stage rocket body at different launch points, the attitude control scheme for the first-stage flight segment can be designed as a de-tasking scheme that adjusts the network gain scheduling according to the launch point altitude and fixes the filter of the envelope elastic mode. That is, before launch, a corresponding gain interpolation table that changes with time is generated according to the altitude of the launch point and used as the parameters of the attitude control system during flight, which can adapt to the changes of the dynamic characteristics of the rocket body in the dense atmosphere under different launch conditions;

[0052] The traditional filter design determines the frequency through vibration tests and then designs the filter. The fixed filter of the envelope elastic mode in this scheme means that the envelope range of the frequency is first determined, and then the filter is designed according to the envelope range of the frequency to achieve the de-tasking design of the filter. Therefore, the filter in this scheme has better adaptability.

[0053] For the second-stage flight segment: Optimize the structure of the second-stage correction network, broaden the notch range of the filter, and a set of fixed correction network parameters can be used for attitude control to achieve the de-tasking goal. The optimization of the structure of the second-stage correction network specifically has two aspects: on the one hand, it is the attitude adjustment threshold. Generally, the larger the threshold, the larger the design space of the correction network and the stronger its adaptability, but the control accuracy will decrease; on the other hand, it is to increase the order of the correction network to improve the flexibility of the design;

[0054] For the last-stage flight segment: According to the different influences of different satellite shape and quality parameters on the control ability, design the corresponding gain scheduling table, and generate the control system parameters according to the shape and quality parameters of the satellite before launch to achieve the de-tasking of the last stage.

[0055] The gain scheduling table contains different control system parameters designed according to the morphological parameters of the satellite. The morphological parameters of the satellite include the mass, center of mass, moment of inertia data, etc. of the satellite. For example, different intervals are divided according to the position of the center of mass of the satellite from the front end face, and different calibration network parameters (i.e., control system parameters) are designed for different intervals.

[0056] S7. Due to the changes in the TT&C scheme and the separation scheme of the satellite and the rocket brought about by different launch missions, the attitude angle adjustment method and the separation timing of the satellite and the rocket will also change, and the flight control software also needs to adapt to the impacts brought about by these changes.

[0057] Conduct a general and specific analysis of the flight control software, design the flight control software, which includes general modules and specific modules. Design a microkernel and hierarchical software architecture to implement the control logic and call relationship of the general modules and specific modules, and achieve adaptability to different tasks by reserving configurable interfaces.

[0058] The general modules include a navigation calculation module, an attitude angle calculation module, a stage transition control module, a bus data processing module, etc.; the specific modules include an attitude control module, a guidance control module, etc.

[0059] The flight control software sequentially calls the general module or specific module corresponding to the flight timing according to the flight timing; among them, before calling the specific module, first conduct a logical judgment according to the flight mission, and then call the specific module that matches the flight mission.

[0060] For example: The rocket's orbit injection method is generally divided into single-stage final ignition orbit injection and transfer orbit injection. The control strategies of the first three stages are the same for different orbit injection methods, that is, they are designed as general modules. The final stage orbit injection methods are different, and the control strategies adopted are also different. Set the different parts as specific modules, and select the corresponding specific module according to different tasks or orbit injection methods.

[0061] The flight control software also includes a mission planning system. The mission planning system performs calculations and data processing. After inputting the input parameters required for the calculations into the mission planning system, the mission planning system outputs relevant configuration parameters;

[0062] The configuration parameters include the configuration parameters in all mission states, and the configuration parameters required for different missions are different; the specific modules use the configuration parameters generated by the mission planning system according to different missions for dynamic combination.

[0063] Exemplarily, the input parameters include the target orbit, launch point, launch time, orbit injection method, etc.;

[0064] Exemplarily, the configuration parameters include the weights of each sub-stage of the rocket, statistical wind field data, attitude control consumption of each stage, flight timing, etc.

Claims

1. A de - missioning design method for launch vehicle launches, characterized in that, including selecting a corresponding landing area based on the target orbit characteristics, and then designing and optimizing the trajectory according to the landing area; pre-designing a ballistic calculation input interface according to the existing target orbit parameters; performing ballistic envelope calculations according to the existing and future possible launch missions; on this basis, selecting the most severe working conditions of the aerodynamic heat environment, load environment, and space heat environment, and respectively carrying out the design of the aerodynamic heat protection scheme, strength design, and space heat protection scheme design; conducting statistical analysis on the space dimensions, installation interfaces, mass, and separation schemes of existing satellites, and designing satellite adapters with different standards; automatically planning a ground-based measurement and control station, a space-based measurement and control station, and a rocket roll angle adjustment scheme that meet the measurement and control requirements according to the position and attitude of the rocket, as well as the position and data transmission capacity information of the measurement and control station; carrying out envelope design for the guidance system according to different orbit injection methods and different satellite-rocket separation schemes; after the ballistic calculation is completed, generating the control timing sequence of the guidance system; designing attitude control schemes for each stage: For the first-stage flight segment, analyze the dynamic characteristic envelope of the first-stage rocket body at different launch points, correct the network gain scheduling according to the launch point altitude, and use a fixed filter with envelope elastic modes for attitude control; For the second-stage flight segment, optimize the structure of the second-stage correction network, broaden the filter notch range, and use fixed correction network parameters for attitude control; For the final-stage flight segment, pre-design a corresponding gain scheduling table according to the different influences of different satellite shape and quality parameters on the control ability, and generate correction network parameters according to the satellite's shape and quality parameters before launch; designing flight control software, the flight control software includes general modules and special modules, designing the software architecture to realize the control logic and call relationship between the general modules and special modules, and realizing the adaptability to different tasks through reserved configurable interfaces.

2. A method for de - tasking the design of a launch vehicle launch mission according to claim 1, characterized in that, The target orbit parameters include target orbit altitude, target orbit inclination, local time of descending node / geographical longitude of ascending node, eccentricity, and argument of perigee.

3. A de-missioning design method for a launch vehicle according to claim 1, characterized in that The ballistic envelope calculation is to perform ballistic calculations for working conditions with different launch sites, different orbit altitudes, and different combinations of orbit inclinations.

4. A de-missioning design method for a launch vehicle according to claim 1, characterized in that, The automatically planning a ground-based measurement and control station, a space-based measurement and control station, and a rocket roll angle adjustment scheme that meet the measurement and control requirements according to the position and attitude of the rocket, as well as the position and data transmission capacity information of the measurement and control station includes: determining ground-based and space-based measurement and control stations that meet the measurement and control needs according to the data transmission requirements of the mission, and then based on the current position and attitude of the rocket and the position of the measurement and control station, calculating the roll angle that meets the measurement and control requirements in the current attitude based on the beam angle constraint of the measurement and control antenna.

5. A method for de - tasking the design of a launch vehicle launch mission according to claim 1, characterized in that For the first-stage flight segment, the attitude control scheme for correcting the network gain scheduling according to the launch point altitude and the de-tasked scheme of the fixed filter with envelope elastic modes includes: generating a corresponding gain interpolation table that changes with time according to the launch point altitude before launch, as the attitude control system parameters during flight; The design method of the fixed filter with envelope elastic modes is to first determine the envelope range of the frequency, and then design the filter according to the envelope range of the frequency to realize the de-tasked design of the filter.

6. A de-missioning design method for a launch vehicle according to claim 1, characterized in that, The shape and quality parameters of the satellite include the mass, center of mass, and moment of inertia data of the satellite.

7. A method for mission de - tasking design of a launch vehicle according to claim 1, characterized in that, The general module includes a navigation calculation module, an attitude angle calculation module, a stage transition control module, and a bus data processing module; the special module includes an attitude control module and a guidance control module.

8. A method for de-tasking the design of a launch vehicle launch mission according to claim 1, characterized in that, The control logic and call relationship of the general module and the special module implemented by the designed software architecture include: The flight control software sequentially calls the general module or the special module corresponding to the flight timing according to the flight timing; among them, before calling the special module, logical judgment is first performed according to the flight mission, and then the special module matching the flight mission is called.

9. A de-missioning design method for a launch vehicle according to claim 1, characterized in that, The flight control software includes a mission planning system. The mission planning system performs calculations and data processing. After inputting the input parameters required for calculation into the mission planning system, the mission planning system outputs relevant configuration parameters; The special module dynamically combines according to the different missions using the configuration parameters generated by the mission planning system.

10. A method for de-missionizing the design of a launch vehicle according to claim 9, characterized in that, The input parameters include the target orbit, the launch point, the launch time, and the orbit injection method; The configuration parameters include the weights of each sub-stage of the rocket, the statistical wind field data, the attitude control consumption of each stage, and the flight timing.