Rapid model selection method and device for carrier rocket power
Through statistical historical data, the load factor, thrust-to-weight ratio and surface-to-face thrust ratio are determined, and the active engine is screened based on the probability of mission success, which solves the problems of launch vehicle power selection time and resource consumption, and achieves rapid selection and reasonable spectrum design.
Patent Information
- Application Number
- CN202411840746.3
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2024-12-13
- Publication Date
- 2025-03-11
AI Technical Summary
When designing the carrier rocket type spectrum, the existing technology requires a lot of time and computing resources, and it is difficult to quickly carry out the selection of carrier rocket power with reasonable capability gradients.
The carrying coefficient, thrust-to-weight ratio and surface thrust-to-face ratio are determined through statistical historical data, and the active engine is screened based on the probability of mission success, and the launch vehicle power is quickly selected.
It has achieved rapid selection of launch vehicle power, supported initial development of launch vehicle spectral design with reasonable capability gradients, laying a theoretical foundation for refined design.
Smart Images
Figure CN119598612A_ABST
Abstract
Description
Technical Field
[0001] The invention relates to the technical field of overall launch vehicle design, and in particular to a launch vehicle power rapid selection method and device. Background Art
[0002] In order to coordinate the development of launch vehicles and avoid the problems of unreasonable gradient of carrying capacity and duplication of capacity construction, it is particularly important to carry out the design of launch vehicle spectrum. The main content of the design of launch vehicle spectrum is to determine the power selection of launch vehicle under the gradient carrying capacity, mainly including the rocket diameter, the type and number of the first-stage engine. For the overall professionals of launch vehicles, the rapid development of the preliminary spectrum design of launch vehicles plays a great role in forming the design capability of the launch vehicle system, and can lay a solid foundation for the development of refined power selection. However, there is an extremely complex nonlinear relationship between the main parameters of carrying capacity and power selection. At present, the normal development of the launch vehicle spectrum design requires detailed launch vehicle configuration design, ballistic design and aerodynamic calculation, which requires a lot of time cost and computing resources, which is not conducive to the generation of professional system design capabilities. Summary of the invention
[0003] In view of the deficiencies in the above-mentioned prior art, the present invention provides a method and device for quickly selecting launch vehicle power, which can quickly select launch vehicle power, support the initial design of a launch vehicle spectrum with a reasonable capability gradient, and lay a theoretical foundation for carrying out refined launch vehicle spectrum design.
[0004] To achieve the above object, the present invention provides a method for rapid selection of launch vehicle power, comprising the following steps: Step 1, obtaining the carrying capacity of the launch vehicle, and obtaining the carrying coefficient based on statistical historical data, thereby calculating the launch vehicle takeoff mass; Step 2, obtaining a thrust-to-weight ratio based on statistical historical data, and calculating the thrust of the launch vehicle based on the thrust-to-weight ratio and the launch vehicle takeoff mass, and then selecting an existing engine; Step 3, calculating the number of engines based on the sea level thrust of the current engines and the thrust of the launch vehicle, and obtaining the surface-to-thrust ratio based on statistical historical data; Step 4, based on the number of engines, the sea level thrust of the engines in service and the surface-to-thrust ratio, the diameter of the launch vehicle is calculated, and it is determined whether the number of engines matches the diameter of the launch vehicle: If yes, proceed to step 5; Otherwise, adjust the surface-to-push ratio and return to step 4; Step 5, calculating the mission success probability of a single-core launch vehicle based on the number of engines, and determining whether the mission success probability meets the requirements: If yes, output the selection results, including the types and number of engines in service and the diameter of the launch vehicle; Otherwise, replace the current engine and return to step 3.
[0005] In one embodiment, in step 1, the calculation process of the launch vehicle takeoff mass is: Obtain the load factors of all launch vehicle historical samples, perform density clustering on all the load factors, and use the maximum and minimum values in the load factor cluster as the upper and lower limits of the load factor range; A loading factor is selected within the loading factor range, and the launch mass of the carrier rocket is calculated based on the selected loading factor and the carrying capacity, which is:
[0006] in, is the launch vehicle takeoff mass, For carrying capacity, is the load factor.
[0007] In one embodiment, in step 2, the calculation process of the launch vehicle thrust is: Obtain the thrust-to-weight ratios of all launch vehicle historical samples, perform density clustering on all thrust-to-weight ratios, and use the maximum and minimum values in the thrust-to-weight ratio clusters as the upper and lower limits of the thrust-to-weight ratio range; A thrust-to-weight ratio is selected within the thrust-to-weight ratio range, and the thrust of the carrier rocket is calculated based on the selected thrust-to-weight ratio and the take-off mass of the carrier rocket, which is:
[0008] in, To push the launch vehicle, is the gravitational acceleration, take , For thrust-to-weight ratio.
[0009] In one embodiment, in step 3, the calculation process of the number of engines is:
[0010] in, is the number of engines, is the sea level thrust of the current engine, To round up.
[0011] In one embodiment, in step 3 and step 4, the process of selecting the thrust ratio, calculating the launch vehicle diameter, and adjusting the thrust ratio is as follows: Obtain the thrust-to-surface ratios of all launch vehicle historical samples, perform density clustering on all thrust-to-surface ratios, and use the maximum and minimum values in the thrust-to-surface ratio clusters as the upper and lower limits of the thrust-to-surface ratio range; Select the upper limit of the surface push ratio range as the current surface push ratio , and based on the number of engines, the sea level thrust of the engines in service and the current thrust ratio The diameter of the launch vehicle is calculated as:
[0012] in, is the launch vehicle diameter; Determine the number of engines The calculated launch vehicle diameter Match: If yes, proceed to step 5; Otherwise, let or Then, return to step 4, where is the reduction coefficient of the surface-to-thrust ratio, is the reduction constant of the surface-to-thrust ratio.
[0013] In one embodiment, step 6 includes: The mission success probability of a single-core launch vehicle is calculated based on the number of engines:
[0014] in, is the probability of mission success, is the probability of success of other systems except the primary power system, is the number of engines, k is the number of redundant engines, p is the success probability of a single engine, is the probability of explosion of a single engine, For Select from engines Number of combinations of engines; Determine the calculated probability of mission success Whether it exceeds the set threshold: If yes, output the selection results, including the types and number of engines in service and the diameter of the launch vehicle; Otherwise, replace the current engine and return to step 3.
[0015] To achieve the above object, the present invention further provides a launch vehicle power rapid selection device, which adopts the above method to perform launch vehicle power rapid selection, and the launch vehicle power rapid selection device comprises: The launch vehicle takeoff mass calculation unit is used to obtain the carrying capacity of the launch vehicle and obtain the carrying coefficient based on the statistical historical data, thereby calculating the launch vehicle takeoff mass; A launch vehicle thrust calculation unit, used to obtain a thrust-to-weight ratio according to statistical historical data, and to calculate the launch vehicle thrust based on the thrust-to-weight ratio and the launch vehicle takeoff mass; An engine number calculation unit, used to calculate the number of engines according to the sea level thrust of the engines currently in service and the thrust of the carrier rocket; A launch vehicle diameter calculation unit, used to obtain a surface-to-thrust ratio according to statistical historical data, and to calculate the launch vehicle diameter based on the number of engines, the sea level thrust of the engines currently in service, and the surface-to-thrust ratio; A surface-thrust ratio adjustment unit, used to determine whether the number of engines matches the diameter of the launch vehicle, and to adjust the surface-thrust ratio if they do not match; A mission success probability calculation unit, used to calculate the mission success probability of a single-core stage launch vehicle according to the number of engines; The success probability judgment unit is used to judge whether the success probability of the mission meets the requirements, and output the selection results when the requirements are met, including the type and number of engines in service and the diameter of the launch vehicle.
[0016] To achieve the above object, the present invention further provides a terminal device, wherein the terminal device is provided with: Memory, used to store programs; The processor is used to execute the program stored in the memory. When the program is executed, the processor is used to execute the method as described above.
[0017] To achieve the above object, the present invention further provides a computer-readable storage medium, wherein the computer-readable storage medium stores computer-executable instructions; when the computer-executable instructions are executed by a processor, they are used to implement the above method.
[0018] Compared with the prior art, the present invention has the following beneficial technical effects: The present invention determines the carrying coefficient, thrust-to-weight ratio and surface-to-thrust ratio by statistically analyzing historical data, and screens the existing engines according to the mission success probability of the launch vehicle. It can quickly select the power of the launch vehicle, support the initial design of the launch vehicle spectrum with a reasonable capability gradient, and lay a theoretical foundation for carrying out refined launch vehicle spectrum design. BRIEF DESCRIPTION OF THE DRAWINGS
[0019] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying creative work.
[0020] Figure 1 This is a flow chart of a method for rapid selection of launch vehicle power in Embodiment 1 of the present invention; Figure 2 The present invention is a statistical diagram of the carrying coefficient of the current launch vehicle in Example 1; Figure 3 A statistical diagram of thrust-to-weight ratio of active launch vehicles in Example 1 of the present invention; Figure 4 This is a statistical diagram of the area-to-thrust ratio of active launch vehicles in Example 1 of the present invention; Figure 5 This is a structural block diagram of a launch vehicle power rapid selection device in Embodiment 2 of the present invention; Figure 6 This is a structural block diagram of the terminal device in Example 2 of the present invention.
[0021] The realization of the purpose, functional features and advantages of the present invention will be further explained in conjunction with embodiments and with reference to the accompanying drawings. DETAILED DESCRIPTION
[0022] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.
[0023] In addition, the technical solutions between the various embodiments of the present invention can be combined with each other, but it must be based on the fact that ordinary technicians in the field can implement it. When the combination of technical solutions is contradictory or cannot be implemented, it should be deemed that such combination of technical solutions does not exist and is not within the scope of protection required by the present invention.
[0024] Example 1 This embodiment discloses a method for rapid selection of launch vehicle power, which can determine the carrying coefficient, thrust-to-weight ratio and surface-to-thrust ratio by statistically analyzing historical data, and screen the existing engines by the mission success probability of the launch vehicle. It can quickly select the launch vehicle power, support the initial development of a launch vehicle spectrum design with a reasonable capability gradient, and lay a theoretical foundation for carrying out refined launch vehicle spectrum design.
[0025] refer to Figure 1 The method for rapid selection of launch vehicle power in this embodiment specifically includes the following steps: Step 1, obtaining the carrying capacity of the launch vehicle, and obtaining the carrying coefficient based on statistical historical data, thereby calculating the launch vehicle takeoff mass; Step 2: derive the thrust-to-weight ratio based on historical statistical data, calculate the launch vehicle thrust based on the thrust-to-weight ratio and the launch vehicle takeoff mass, and then select an existing engine; Step 3, the number of engines is calculated based on the sea level thrust of the current engines and the thrust of the launch vehicle, and the surface thrust ratio is obtained based on the statistical historical data; Step 4: Calculate the diameter of the launch vehicle based on the number of engines, the sea level thrust and surface-to-thrust ratio of the engines currently in service, and determine whether the number of engines matches the diameter of the launch vehicle: If yes, proceed to step 5; Otherwise, adjust the surface-to-push ratio and return to step 4; Step 5: Calculate the mission success probability of a single-core launch vehicle based on the number of engines, and determine whether the mission success probability meets the requirements: If yes, output the selection results, including the types and number of engines in service and the diameter of the launch vehicle; Otherwise, replace the current engine and return to step 3.
[0026] In the specific implementation process of step 1, the calculation process of the launch vehicle takeoff mass is: First, the load factors of all launch vehicle historical samples are obtained, and all load factors are clustered by density. The maximum and minimum values in the load factor clusters are used as the upper and lower limits of the load factor range. Select a load factor within the load factor range, and calculate the launch vehicle takeoff mass based on the selected load factor and carrying capacity, which is:
[0027] in, is the launch vehicle takeoff mass, For carrying capacity, is the load factor.
[0028] Of course, in the specific application process, it is not limited to using density clustering to determine the load factor range. You can also artificially define a load factor range so that the load factors of most launch vehicle historical samples are within this range, for example, more than 40% of the load factors are within this range. Figure 2 The figure shows the statistics of the payload coefficient of current launch vehicles, where the payload coefficient is generally between 2% and 4%. The payload coefficient range can be directly defined as When selecting a load factor within the load factor interval, one can either randomly select one within the load factor interval or directly select the median of the load factor interval.
[0029] In the specific implementation process of step 2, the calculation process of the launch vehicle thrust is: Obtain the thrust-to-weight ratios of all launch vehicle historical samples, perform density clustering on all thrust-to-weight ratios, and use the maximum and minimum values in the thrust-to-weight ratio clusters as the upper and lower limits of the thrust-to-weight ratio range; Select a thrust-to-weight ratio within the thrust-to-weight ratio range, and calculate the launch vehicle thrust based on the selected thrust-to-weight ratio and the launch vehicle takeoff mass, which is:
[0030] in, is the launch vehicle thrust, is the gravitational acceleration, take , For thrust-to-weight ratio.
[0031] Similarly, in the specific application process, it is not limited to using density clustering to determine the thrust-to-weight ratio range. A thrust-to-weight ratio range can also be artificially defined so that the thrust-to-weight ratio of most launch vehicle historical samples is within this range, for example, more than 60% of the thrust-to-weight ratio is within this range. Figure 3 The following is a statistical chart of thrust-to-weight ratios of current launch vehicles, where the thrust-to-weight ratio is generally between 1.2 and 1.3. The thrust-to-weight ratio range can be directly defined as When selecting a thrust-to-weight ratio within the thrust-to-weight ratio range, one can either randomly select one within the thrust-to-weight ratio range or directly select the median of the thrust-to-weight ratio range.
[0032] In the specific implementation process of step 3, the calculation process of the number of engines is:
[0033] in, is the number of engines, is the sea level thrust of the current engine, To round up.
[0034] In the specific implementation process of step 3 and step 4, the process of selecting the surface thrust ratio, calculating the launch vehicle diameter, and adjusting the surface thrust ratio is as follows: Obtain the thrust-to-surface ratios of all launch vehicle historical samples, perform density clustering on all thrust-to-surface ratios, and use the maximum and minimum values in the thrust-to-surface ratio clusters as the upper and lower limits of the thrust-to-surface ratio range; Select the upper limit of the surface push ratio range as the current surface push ratio , and based on the number of engines, the sea level thrust of the engines in service and the current thrust ratio The diameter of the launch vehicle is calculated as:
[0035] in, is the launch vehicle diameter; Determine the number of engines The calculated launch vehicle diameter Whether it matches, that is, judging the current launch vehicle diameter Can it be installed? The current engines in service are: If yes, proceed to step 5; Otherwise, let or Then, return to step 4, where is the reduction coefficient of the surface-to-thrust ratio, is the reduction constant of the surface-to-thrust ratio.
[0036] Similarly, in the specific application process, it is not limited to using density clustering to determine the area-to-thrust ratio range. An area-to-thrust ratio range can also be artificially defined so that the area-to-thrust ratio of most launch vehicle historical samples is within this range, for example, more than 50% of the area-to-thrust ratio is within this range. Figure 4 The figure shows the statistics of the thrust-to-area ratio of current launch vehicles. The thrust-to-area ratio is generally between 300 kN / ㎡ and 600 kN / ㎡. The thrust-to-area ratio range can be directly defined as .
[0037] In the specific implementation process of step 6, the process of calculating the mission success probability of a single-core stage launch vehicle based on the number of engines is as follows: Given number of redundant units k (i.e. k When an engine fails, the mission can still be successful), the success probability of a single engine p , the probability of a single engine exploding (i.e. a single point failure occurs, and if any engine explodes, the mission is considered a failure) q , the probability of success of other systems except the primary power system , combined with the calculated number of engines , the mission success probability of a single-core launch vehicle can be calculated as:
[0038] in, is the probability of mission success, For Select from engines Number of combinations of engines; Determine the calculated probability of mission success Whether it exceeds the set threshold: If yes, output the selection results, including the types and number of engines in service and the diameter of the launch vehicle; Otherwise, replace the current engine and recalculate the number of engines and the diameter of the launch vehicle until Meet the requirements.
[0039] It is worth noting that although this embodiment Figure 1 The steps in the process are shown in sequence as indicated by the arrows, but the steps are not necessarily executed in the order indicated by the arrows. Unless otherwise specified in this document, there is no strict order restriction for the execution of the steps, and the steps can be executed in other orders. Figure 1 At least part of the steps may include multiple sub-steps or multiple stages. These sub-steps or stages are not necessarily executed at the same time, but can be executed at different times. The execution order of these sub-steps or stages is not necessarily sequential, but can be executed in turn or alternately with other steps or at least part of the sub-steps or stages of other steps.
[0040] Example 2 Based on the launch vehicle power rapid selection method in Example 1, this embodiment discloses a launch vehicle power rapid selection device, referring to Figure 5 The launch vehicle power rapid selection device includes a launch vehicle takeoff mass calculation unit, a launch vehicle thrust calculation unit, an engine number calculation unit, a launch vehicle diameter calculation unit, a surface-to-thrust ratio adjustment unit, a mission success probability calculation unit, and a success probability judgment unit. Specifically: The launch vehicle takeoff mass calculation unit is used to obtain the carrying capacity of the launch vehicle and obtain the carrying coefficient based on statistical historical data, thereby calculating the launch vehicle takeoff mass; The launch vehicle thrust calculation unit is used to obtain the thrust-to-weight ratio according to the statistical historical data, and to calculate the launch vehicle thrust based on the thrust-to-weight ratio and the launch vehicle takeoff mass; The engine number calculation unit is used to calculate the number of engines according to the sea level thrust of the current engines and the thrust of the launch vehicle; The launch vehicle diameter calculation unit is used to obtain the surface-to-thrust ratio based on statistical historical data, and calculate the launch vehicle diameter based on the number of engines, the sea level thrust of the engines in service, and the surface-to-thrust ratio; The thrust-to-surface ratio adjustment unit is used to determine whether the number of engines matches the diameter of the launch vehicle, and adjust the thrust-to-surface ratio if they do not match; The mission success probability calculation unit is used to calculate the mission success probability of a single-core stage launch vehicle according to the number of engines; The success probability judgment unit is used to judge whether the mission success probability meets the requirements, and output the selection results when the requirements are met, including the type and number of engines in service and the diameter of the launch vehicle.
[0041] In this embodiment, the specific working process and working principle of the launch vehicle takeoff mass calculation unit, the launch vehicle thrust calculation unit, the engine number calculation unit, the launch vehicle diameter calculation unit, the surface-to-thrust ratio adjustment unit, the mission success probability calculation unit and the success probability judgment unit are the same as those in the method in Example 1, so they are not repeated in this embodiment. Each unit module can be implemented in whole or in part by software, hardware and a combination thereof, and each unit module can be embedded in or independent of the processor in the computer device in the form of hardware, or can be stored in the memory in the computer device in the form of software, so that the processor can call and execute the operations corresponding to each of the above unit modules.
[0042] Example 3 like Figure 6 The terminal device disclosed in this embodiment includes a transmitter, a receiver, a memory, and a processor. The transmitter is used to send instructions and data, the receiver is used to receive instructions and data, the memory is used to store computer-executable instructions, and the processor is used to execute the computer-executable instructions stored in the memory to implement the method in the above-mentioned embodiment 1.
[0043] It should be noted that the above memory can be independent or integrated with the processor. When the memory is independently provided, the terminal device further includes a bus for connecting the memory and the processor.
[0044] Example 4 This embodiment discloses a computer-readable storage medium, in which computer-executable instructions are stored. When a processor executes the computer-executable instructions, the method in the above-mentioned embodiment 1 is implemented.
[0045] The above description is only a preferred embodiment of the present invention, and does not limit the patent scope of the present invention. All equivalent structural changes made by using the contents of the present invention specification and drawings under the inventive concept of the present invention, or directly / indirectly applied in other related technical fields are included in the patent protection scope of the present invention.
Claims
1. A method for rapid selection of launch vehicle power, characterized in that: The steps include: Step 1, obtaining the carrying capacity of the launch vehicle, and obtaining the carrying coefficient based on statistical historical data, thereby calculating the launch vehicle takeoff mass; Step 2, obtaining a thrust-to-weight ratio based on statistical historical data, and calculating the thrust of the launch vehicle based on the thrust-to-weight ratio and the launch vehicle takeoff mass, and then selecting an existing engine; Step 3, calculating the number of engines based on the sea level thrust of the current engines and the thrust of the launch vehicle, and obtaining the surface-to-thrust ratio based on statistical historical data; Step 4, based on the number of engines, the sea level thrust of the engines in service and the surface-to-thrust ratio, the diameter of the launch vehicle is calculated, and it is determined whether the number of engines matches the diameter of the launch vehicle: If yes, proceed to step 5; Otherwise, adjust the surface-to-push ratio and return to step 4; Step 5, calculating the mission success probability of a single-core launch vehicle based on the number of engines, and determining whether the mission success probability meets the requirements: If yes, output the selection results, including the types and number of engines in service and the diameter of the launch vehicle; Otherwise, replace the current engine and return to step 3.
2. The method for rapid selection of launch vehicle power according to claim 1, characterized in that: In step 1, the calculation process of the launch vehicle takeoff mass is: Obtain the load factors of all launch vehicle historical samples, perform density clustering on all the load factors, and use the maximum and minimum values in the load factor cluster as the upper and lower limits of the load factor range; A loading factor is selected within the loading factor range, and the launch mass of the carrier rocket is calculated based on the selected loading factor and the carrying capacity, which is: in, is the launch vehicle takeoff mass, For carrying capacity, is the load factor.
3. The method for rapid selection of launch vehicle power according to claim 1, characterized in that: In step 2, the calculation process of the launch vehicle thrust is: Obtain the thrust-to-weight ratios of all launch vehicle historical samples, perform density clustering on all thrust-to-weight ratios, and use the maximum and minimum values in the thrust-to-weight ratio clusters as the upper and lower limits of the thrust-to-weight ratio range; A thrust-to-weight ratio is selected within the thrust-to-weight ratio range, and the thrust of the carrier rocket is calculated based on the selected thrust-to-weight ratio and the take-off mass of the carrier rocket, which is: in, is the launch vehicle thrust, is the launch vehicle takeoff mass, is the acceleration due to gravity, For thrust-to-weight ratio.
4. The method for rapid selection of launch vehicle power according to claim 1, characterized in that: In step 3, the calculation process of the number of engines is: in, is the number of engines, is the launch vehicle thrust, is the sea level thrust of the current engine, To round up.
5. The method for rapid selection of launch vehicle power according to claims 1 to 4, characterized in that: In step 3 and step 4, the process of selecting the thrust ratio, calculating the launch vehicle diameter, and adjusting the thrust ratio is as follows: Obtain the thrust-to-surface ratios of all launch vehicle historical samples, perform density clustering on all thrust-to-surface ratios, and use the maximum and minimum values in the thrust-to-surface ratio clusters as the upper and lower limits of the thrust-to-surface ratio range; Select the upper limit of the surface push ratio range as the current surface push ratio , and based on the number of engines, the sea level thrust of the engines in service and the current thrust ratio The diameter of the launch vehicle is calculated as: in, is the launch vehicle diameter, is the number of engines, is the sea level thrust of the engines currently in service; Determine the number of engines The calculated launch vehicle diameter Match: If yes, proceed to step 5; Otherwise, let or Then, return to step 4, where is the reduction coefficient of the surface-to-thrust ratio, is the reduction constant of the surface-to-thrust ratio.
6. The method for rapid selection of launch vehicle power according to claims 1 to 4, characterized in that: Step 6 includes: The mission success probability of a single-core launch vehicle is calculated based on the number of engines: in, is the probability of mission success, is the probability of success of other systems except the primary power system, is the number of engines, k is the number of redundant engines, p is the success probability of a single engine, is the probability of explosion of a single engine, For Select from engines Number of combinations of engines; Determine the calculated probability of mission success Whether it exceeds the set threshold: If yes, output the selection results, including the types and number of engines in service and the diameter of the launch vehicle; Otherwise, replace the current engine and return to step 3.
7. A launch vehicle power rapid selection device, characterized in that: The method according to any one of claims 1 to 6 is used to quickly select the power of a launch vehicle, and the device for quickly selecting the power of a launch vehicle comprises: The launch vehicle takeoff mass calculation unit is used to obtain the carrying capacity of the launch vehicle and obtain the carrying coefficient based on the statistical historical data, thereby calculating the launch vehicle takeoff mass; A launch vehicle thrust calculation unit, used to obtain a thrust-to-weight ratio according to statistical historical data, and to calculate the launch vehicle thrust based on the thrust-to-weight ratio and the launch vehicle takeoff mass; An engine number calculation unit, used to calculate the number of engines according to the sea level thrust of the engines currently in service and the thrust of the carrier rocket; A launch vehicle diameter calculation unit, used to obtain a surface-to-thrust ratio according to statistical historical data, and to calculate the launch vehicle diameter based on the number of engines, the sea level thrust of the engines currently in service, and the surface-to-thrust ratio; A surface-thrust ratio adjustment unit, used to determine whether the number of engines matches the diameter of the launch vehicle, and to adjust the surface-thrust ratio if they do not match; A mission success probability calculation unit, used to calculate the mission success probability of a single-core stage launch vehicle according to the number of engines; The success probability judgment unit is used to judge whether the success probability of the mission meets the requirements, and output the selection results when the requirements are met, including the type and number of engines in service and the diameter of the launch vehicle.
8. A terminal device, characterized in that: The terminal device is provided with: Memory, used to store programs; A processor, configured to execute the program stored in the memory; when the program is executed, the processor is configured to execute the method according to any one of claims 1 to 6.
9. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores computer-executable instructions; when the computer-executable instructions are executed by a processor, they are used to implement the method according to any one of claims 1 to 6.