A launch vehicle aerodynamic heat rapid calculation method, device and electronic equipment

Abstract

The present application relates to a kind of launch vehicle aerodynamic heat rapid calculation method, according to different launch vehicle aerodynamic heat calculation state establishes several aerodynamic heat calculation programs, forms aerodynamic heat calculation APP by encapsulating several aerodynamic heat calculation program source code, extract variable parameter in code to operation interface, realize the quick calculation aerodynamic heat by setting interface parameter.For the aerodynamic heat response of different positions of rocket calculation, the APP currently encapsulated code establishes eight programs, such as the program of the point of a layer of non-metallic material calculation, i.e.the aerodynamic heat calculation source code based on MATLAB, above program can basically include all the launch vehicle aerodynamic heat calculation state, subsequent if new calculation state appears, can be increased to aerodynamic heat calculation APP by similar method New program.

Description

Technical Field

[0001] This invention belongs to the field of aerodynamic calculation of launch vehicles, and specifically relates to a method, device, electronic equipment and storage medium for rapid calculation of aerodynamics and heat of launch vehicles. Background Technology

[0002] The thermal protection design of launch vehicles is crucial to the success or failure of flight test missions. Traditional aerodynamic and thermal calculations using the CFD finite volume method are resource-intensive. Pre-processing mesh generation requires significant manual intervention, and because aerodynamic and thermal calculations demand sufficiently fine meshes, the mesh size is often enormous, on the order of tens of millions. Even on high-performance clusters, completing a single calculation cycle can take several weeks. During the launch vehicle design and feasibility study phase, the iteration speed of aerodynamic and thermal calculations impacts the project's development schedule; therefore, it is necessary to establish a rapid aerodynamic and thermal calculation method. Summary of the Invention

[0003] In the launch vehicle design demonstration phase, to achieve rapid iteration of aerodynamic and thermal calculations, this invention provides a rapid aerodynamic and thermal calculation method for launch vehicles. Several aerodynamic and thermal calculation programs are established according to different launch vehicle aerodynamic and thermal calculation states. The source code of several aerodynamic and thermal calculation programs is encapsulated to form an aerodynamic and thermal calculation APP. Variable parameters in the code are extracted to the operation interface. By setting the interface parameters, rapid aerodynamic and thermal calculations can be achieved.

[0004] Furthermore, the aerodynamic and thermal calculation program established based on the aerodynamic and thermal calculation state of the launch vehicle includes...

[0005] Calculation program for non-metallic materials on the first floor of the station.

[0006] A calculation program for a non-metallic layer of material in the fairing curve segment.

[0007] Calculation program for two layers of non-metallic materials in the fairing curve segment.

[0008] Calculation program for non-metallic materials in a single layer of a column segment.

[0009] Calculation program for two layers of non-metallic materials in column segments

[0010] Calculation program for the metal material of a single layer of column segment.

[0011] Grid rudder leading edge calculation program,

[0012] Grid rudder surface calculation program.

[0013] Furthermore, the packaging steps of the aerodynamic thermal calculation APP include:

[0014] S1. Parse the code of each program and extract the variable parameters to the operation interface, mainly including the calculation start time, calculation end time, calculation section position, thermal protection layer type and thickness parameters.

[0015] S2. Design an input box for calculating the number of points. Through this input box, the number of rows on the input interface can be changed, thereby changing the number of calculation states.

[0016] S3. Input interface design: Create an empty table with several columns, including compartment, interface position, number of material layers, calculation end time, thickness of the first material layer, material of the second layer, and thickness of the second material layer; bind the input parameter array to this table to achieve the effect of displaying the input parameters on the interface.

[0017] S4. Output interface design: Create an empty table and several columns, including the compartment name, interface location, outer wall temperature, intermediate layer temperature, and inner wall temperature; bind the calculated output parameter array to this table and display the calculation results on the interface.

[0018] S5. Save the temperature curves calculated each time, naming the file according to the section name + location for easy identification;

[0019] S6. Envelope analysis function design: compare the calculation results with the data from previous flight tests to determine whether the calculation results are within the successful envelope range.

[0020] Furthermore, the calculation start time in step S1 is determined based on the actual trajectory, and its value is the trajectory time when the flight Mach number Ma = 1. The corresponding time value is input.

[0021] The calculation end time is the aerodynamic and thermal calculation end time of this section;

[0022] The cross-sectional position is the calculated axial distance between the foremost face of the module and the theoretical apex of the rocket, and should be input according to the actual situation;

[0023] The heat protection layer type refers to the different procedures corresponding to one-layer and two-layer heat protection schemes, and the different procedures corresponding to metallic and non-metallic materials;

[0024] The thickness parameter refers to the input based on the actual thickness.

[0025] Furthermore, in step S2, the number of calculation points is related to the specific number of compartments. By inputting the data through the interface and clicking "Update Points," the number of rows in the calculation status will be automatically updated, ensuring that the number of rows on the interface matches the number of points being solved.

[0026] Furthermore, in step S3, for the first layer of material, the material type is selected from the drop-down list according to the actual scheme, and the drop-down list includes all materials in the material library; the input method for the second layer of material and thickness is the same as the method for the first layer.

[0027] The number of material layers can be freely input according to the thermal protection scheme;

[0028] Before step S3, design the material library maintenance interface, add material types, and change material parameters, specifically including the material's density, specific heat capacity, and thermal conductivity, in SI units. Click "Save to Material Library" to update the material library.

[0029] Furthermore, each aerodynamic and thermal calculation program established based on different launch vehicle aerodynamic and thermal calculation states calculates each state point. The specific calculation process is as follows:

[0030] First, click on ballistic analysis to process the upstream standard ballistic file. Then, extract the three columns of data (time, altitude, and Mach number) from the ballistic file separately, remove duplicate time points, and save it as a ballistic.dat file.

[0031] Then, atmospheric parameters are solved.

[0032] Then set the material parameters, including thickness, density, specific heat capacity, and thermal conductivity;

[0033] Then set the initial temperature for calculation;

[0034] Then perform von Neumann stability analysis iterative calculations and output the spatial step size of the difference scheme;

[0035] Solve using the difference method;

[0036] Finally, the temperature curve and the maximum temperature value are output.

[0037] The present invention also provides a rapid aerodynamic and thermal calculation device for launch vehicles, comprising:

[0038] Several aerodynamic thermal calculation program modules are used to perform aerodynamic thermal calculations based on different launch vehicle aerodynamic thermal calculation states;

[0039] The aerodynamic thermal calculation APP module is used to encapsulate several aerodynamic thermal calculation programs to form an aerodynamic thermal calculation APP with an operating interface.

[0040] The present invention also provides an electronic device, comprising: a processor;

[0041] The memory is communicatively connected to the processor.

[0042] At least one program, stored in the memory and configured to be executed by the processor, is configured to: implement the aforementioned rapid aerodynamic thermal calculation method for launch vehicles.

[0043] The present invention also provides a computer-readable storage medium having a computer program stored thereon, characterized in that the computer program, when executed by an electronic device, implements the aforementioned rapid aerodynamic and thermal calculation method for launch vehicles.

[0044] The present invention has the following advantages:

[0045] Several aerodynamic and thermal calculation programs are established based on different aerodynamic and thermal calculation states of launch vehicles. These programs are then encapsulated into an aerodynamic and thermal calculation app, with variable parameters extracted from the code and integrated into the user interface. By setting these interface parameters, a one-click aerodynamic and thermal calculation function is achieved. These programs can simultaneously calculate the aerodynamic and thermal properties of any number of modules. The number of thermal protection layers, material type, and material thickness for each module are fully visualized on the interface, and different calculation programs can be selected from a dropdown menu for different calculation states. A material library maintenance function is also included, making it very convenient to update material parameters and avoiding calculation errors caused by manually inputting material parameters. Data analysis functions are also provided, allowing direct interpretation of the calculation results after completion to determine whether the results meet design requirements. The types of different programs selected for different calculation states include: module type (stagnation point, cone section, column section), material type (metallic, non-metallic), and number of thermal protection layers (one layer, two layers). Different programs are selected based on these different types.

[0046] To calculate the aerodynamic and thermal response of rockets at different locations, the currently encapsulated code has created eight programs, including a non-metallic material calculation program for the stationary layer. These are aerodynamic and thermal calculation source codes written in MATLAB. These programs can basically cover all aerodynamic and thermal calculation states of launch vehicles. If new calculation states appear in the future, similar methods can be used to add new programs to the aerodynamic and thermal calculation app. Attached Figure Description

[0047] Figure 1 This is a schematic diagram of the aerodynamic thermal calculation program of the method of the present invention;

[0048] Figure 2 This is a schematic diagram of the calculation input interface of the method of the present invention;

[0049] Figure 3 This is a schematic diagram of the calculation results interface of the method of the present invention;

[0050] Figure 4 This is a schematic diagram of the material library interface of the method of the present invention;

[0051] Figure 5This is a schematic diagram of the envelope analysis interface of the method of the present invention;

[0052] Figure 6 This is a schematic diagram of the arrow body wall model of the method of the present invention. Detailed Implementation

[0053] This invention provides a rapid aerodynamic and thermal calculation method for launch vehicles. Several aerodynamic and thermal calculation programs are established according to different aerodynamic and thermal calculation states of launch vehicles. The source code of several aerodynamic and thermal calculation programs is encapsulated to form an aerodynamic and thermal calculation APP. Variable parameters in the code are extracted to the operation interface. By setting the interface parameters, rapid aerodynamic and thermal calculation can be achieved.

[0054] The aerodynamic and thermal calculation program established based on the aerodynamic and thermal calculation state of the launch vehicle includes:

[0055] Calculation program for non-metallic materials on the first floor of the station.

[0056] A calculation program for a non-metallic layer of material in the fairing curve segment.

[0057] Calculation program for two layers of non-metallic materials in the fairing curve segment.

[0058] Calculation program for non-metallic materials in a single layer of a column segment.

[0059] Calculation program for two layers of non-metallic materials in column segments

[0060] Calculation program for the metal material of a single layer of column segment.

[0061] Grid rudder leading edge calculation program,

[0062] Grid rudder surface calculation program.

[0063] In terms of interface design, the sysware.IDE industrial APP development platform is used for development. The packaging and development steps of the aerothermal calculation APP include:

[0064] S1. Parse the code of each program and extract the variable parameters to the operation interface, mainly including the calculation start time, calculation end time, calculation section position, thermal protection layer type and thickness parameters.

[0065] The calculation start time is determined based on the actual trajectory, and its value is the trajectory time when the flight Mach number Ma = 1. Input the corresponding time value. Since the program is best suited for Mach numbers above 1, the aerodynamic heat at low Mach numbers is not significant and can be ignored. Therefore, the calculation starts from Mach 1 and the Mach number is always greater than 1 thereafter.

[0066] The calculation end time is the aerodynamic and thermal calculation end time of this section; for example, the fairing calculation end time is the fairing jettison time, and the first-stage engine calculation end time is the first-stage separation time.

[0067] The cross-sectional position is the calculated axial distance between the foremost face of the module and the theoretical apex of the rocket, and should be input according to the actual situation;

[0068] The heat protection layer type refers to the different procedures corresponding to one-layer and two-layer heat protection schemes, and the different procedures corresponding to metallic and non-metallic materials;

[0069] The number of material layers can be freely entered according to the thermal protection scheme.

[0070] The thickness parameter refers to the input based on the actual thickness.

[0071] S2. Design an input box for calculating the number of points. Through this input box, the number of rows on the input interface can be changed, thereby changing the number of calculation states.

[0072] The number of calculation points is determined, which is related to the specific number of compartments. The calculation points are entered through the interface, and clicking "Update Points" will automatically update the number of rows in the calculation status, ensuring that the number of rows on the interface matches the number of points to be solved.

[0073] S3. Input interface design: Create an empty table with several columns, including compartment, interface position, number of material layers, calculation end time, thickness of the first material layer, material of the second layer, and thickness of the second material layer; bind the input parameter array to this table to achieve the effect of displaying the input parameters on the interface.

[0074] For the first layer of materials, select the material type from the drop-down menu based on the actual solution. The drop-down list includes all materials in the material library. The input method for the second layer of materials and thickness is the same as the first layer.

[0075] The number of material layers can be freely input according to the thermal protection scheme;

[0076] Before step S3, design the material library maintenance interface, add material types, and change material parameters, specifically including the material's density, specific heat capacity, and thermal conductivity, in SI units. Click "Save to Material Library" to update the material library.

[0077] S4. Output interface design: Create an empty table and several columns (five columns), including the compartment name, interface location, outer wall temperature, intermediate layer temperature, and inner wall temperature; bind the calculated output parameter array to this table and display the calculation results on the interface.

[0078] S5. Save the temperature curves calculated each time. Name the file according to the section name + location for easy identification. For example, file name "Secondary Engine (8.339).png" or "Transfer Module (3.12).png" contains text and numbers. The text part is the section name and the data in parentheses is the cross-sectional location.

[0079] S6. Envelope analysis function design: compare the calculation results with the data from previous flight tests to determine whether the calculation results are within the successful envelope range.

[0080] A method for rapid aerodynamic thermal calculation using the APP of this invention:

[0081] After completing the input on each interface, click on the program selection. Select different calculation programs according to different thermal protection schemes. The main differences are related to the calculation location and the number of thermal protection layers. Different programs are required for the head stagnation point, the fairing curved section, and the rocket body column section. Different programs are required for one-layer and two-layer thermal protection schemes. Different programs are required for metallic and non-metallic materials.

[0082] Click "Calculate" again to automatically calculate each state point.

[0083] Specifically, each aerodynamic and thermal calculation program, established based on the different aerodynamic and thermal calculation states of the launch vehicle, calculates each state point. The specific calculation process is as follows:

[0084] First, click on "Ballistic Analysis" to process the upstream standard ballistic file. Extract the time, altitude, and Mach number data columns separately, remove duplicate time points, and perform ballistic interpolation (time step 0.01s). Save the result as a "ballistic.dat" file. (Interpolate first, then save the .dat file.)

[0085] Then, atmospheric parameters are solved.

[0086] Then set the material parameters, including thickness, density, specific heat capacity, and thermal conductivity;

[0087] Then set the initial temperature for calculation;

[0088] Then perform von Neumann stability analysis iterative calculations and output the spatial step size of the difference scheme;

[0089] Solve using the difference method;

[0090] Finally, the output calculation results include the peak temperatures of the inner and outer walls of each compartment; if there are two layers of thermal protection material, the temperature of the middle layer is also output; temperature curves and maximum temperature values ​​are also output.

[0091] Each aerothermal calculation program is written in MATLAB based on the QJ1734-1989 Tactical Missile Aerothermal Engineering Calculation Method. Taking a cylindrical section with a single-layer non-metallic wall as an example, the program writing idea is explained in detail:

[0092] The heat transfer coefficient is calculated using the formula for compressible flat plate heat transfer coefficient.

[0093]

[0094] In the formula c p Take 1017.4.

[0095]

[0096] In the formula, L is the length from the top of the rocket nose cone along the object surface to the calculation point;

[0097] When Re<10 7 hour,

[0098] c f =0.0592(Re * ) -0.2

[0099] When Re≥10 7 hour,

[0100]

[0101]

[0102]

[0103]

[0104] T * =T ∞ +0.5(T w -T ∞ )+0.22(T aw -T ∞ )

[0105]

[0106] Where r = 0.89;

[0107] Generally, the thickness of the arrow body's wall is much smaller than that of the other two directions. Therefore, the wall can be approximated as a flat plate for simplification. Within the wall, heat conduction is considered only in the direction perpendicular to the surface.

[0108] The arrow body wall can be divided into j layers of equal thickness, each layer having a thickness of Δy, such as... Figure 6 As shown;

[0109] The thermal balance equations and initial conditions for each layer of the thick-walled thermal structure are as follows.

[0110]

[0111]

[0112]

[0113] When t=0

[0114] T j =T n =T1=T0

[0115] By solving the difference equation and substituting the initial conditions, the temperature of each layer at any given time can be obtained.

[0116] External wall temperature:

[0117] Inner wall temperature:

[0118] Intermediate layer temperature:

[0119] The symbols are explained in Table 1 below.

[0120] Table 1 below

[0121]

[0122]

[0123] The explanation of the superscript of the symbols is shown in Table 2 below.

[0124] Table 2

[0125] Superscript significance * Reference conditions K The current time value can be omitted. K-1 Previous time interval value

[0126] The explanation of the subscripts for the symbols is shown in Table 3 below.

[0127] Table 3

[0128] subscript significance ∞ Free flow n The nth layer of thermal insulation material

[0129] The present invention also provides a rapid aerodynamic and thermal calculation device for launch vehicles, comprising:

[0130] Several aerodynamic thermal calculation program modules are used to perform aerodynamic thermal calculations based on different launch vehicle aerodynamic thermal calculation states;

[0131] The aerodynamic thermal calculation APP module is used to encapsulate several aerodynamic thermal calculation programs to form an aerodynamic thermal calculation APP with an operating interface.

[0132] The encapsulated aerodynamic thermal calculation APP module specifically includes:

[0133] The code parsing module is used to extract the variable parameters from each aerodynamic calculation program and display them on the user interface.

[0134] The input / output interface module is used for interface input and output, enabling the display of input parameters and calculation results on the interface.

[0135] The module for saving the calculated temperature curve is used to save the calculated temperature curve.

[0136] The envelope analysis module is used to compare the calculation results with the data from previous flight tests to determine whether the calculation results are within the successful envelope range.

[0137] The present invention also provides an electronic device, comprising: a processor;

[0138] The memory is communicatively connected to the processor.

[0139] At least one program, stored in the memory and configured to be executed by the processor, is configured to: implement the aforementioned rapid aerodynamic thermal calculation method for launch vehicles.

[0140] The present invention also provides a computer-readable storage medium having a computer program stored thereon, characterized in that the computer program, when executed by an electronic device, implements the aforementioned rapid aerodynamic and thermal calculation method for launch vehicles.

Claims

1. A method for rapid calculation of aerodynamic heat of a launch vehicle, characterized in that... Several aerodynamic and thermal calculation programs are established based on different launch vehicle aerodynamic and thermal calculation states. The source code of several aerodynamic and thermal calculation programs is encapsulated to form an aerodynamic and thermal calculation APP. Variable parameters in the code are extracted to the operation interface. By setting the interface parameters, aerodynamic and thermal calculations can be performed quickly. The packaging steps of the aerodynamic thermal calculation APP include: S1. Parse the code of each program and extract the variable parameters to the operation interface, including the calculation start time, calculation end time, calculation section position, thermal protection layer type and thickness parameters; S2. Design an input box for calculating the number of points. Through this input box, the number of rows on the input interface can be changed, thereby changing the number of calculation states. S3. Input interface design: Create an empty table with several columns, including section, section location, number of material layers, calculation end time, first layer material, first layer material thickness, second layer material, and second layer material thickness; bind the input parameter array to this table to achieve the effect of displaying the input parameter interface. S4. Output interface design: Create an empty table with several columns, including the compartment name, section location, outer wall temperature, intermediate layer temperature, and inner wall temperature; bind the calculated output parameter array to this table and display the calculation results on the interface. S5. Save the temperature curves calculated each time, naming the file according to the section name + location for easy identification; S6. Envelope analysis function design: compare the calculation results with the data from previous flight tests to determine whether the calculation results are within the successful envelope range.

2. The rapid aerodynamic and thermal calculation method for launch vehicles according to claim 1, characterized in that, The aerodynamic and thermal calculation program established based on the aerodynamic and thermal calculation state of the launch vehicle includes: Calculation program for non-metallic materials on the first floor of the station. A calculation program for a non-metallic layer of material in the fairing curve segment. Calculation program for two layers of non-metallic materials in the fairing curve segment. Calculation program for non-metallic materials in a single layer of a column segment. Calculation program for two layers of non-metallic materials in column segments Calculation program for the metal material of a single layer of column segment. Grid rudder leading edge calculation program, Grid rudder surface calculation program.

3. The rapid aerodynamic and thermal calculation method for launch vehicles according to claim 1, characterized in that, The calculation start time in step S1 is determined based on the actual trajectory, and its value is the trajectory time when the flight Mach number Ma = 1. Input the corresponding time value. The calculation end time is the aerodynamic and thermal calculation end time of this section; The cross-sectional position is the calculated axial distance between the foremost face of the module and the theoretical apex of the rocket, and should be input according to the actual situation; The heat protection layer type refers to the different procedures corresponding to one-layer and two-layer heat protection schemes, and the different procedures corresponding to metallic and non-metallic materials; The thickness parameter refers to the input based on the actual thickness.

4. The rapid aerodynamic and thermal calculation method for launch vehicles according to claim 1, characterized in that, In step S2, the number of calculation points is related to the specific number of compartments. By inputting the data through the interface and clicking "Update Points," the number of rows in the calculation status will be automatically updated, ensuring that the number of rows on the interface matches the number of points being solved.

5. The rapid aerodynamic and thermal calculation method for launch vehicles according to claim 1, characterized in that, In step S3, for the first layer of material, the material type is selected from the drop-down list according to the actual plan, and the drop-down list includes all materials in the material library; the input method for the second layer of material and thickness is the same as the first layer method. The number of material layers can be freely input according to the thermal protection scheme; Before step S3, design the material library maintenance interface, add material types, and change material parameters, specifically including the material's density, specific heat capacity, and thermal conductivity, in SI units. Click "Save to Material Library" to update the material library.

6. The rapid aerodynamic and thermal calculation method for launch vehicles according to claim 2, characterized in that, Each aerodynamic and thermal calculation program, established based on different aerodynamic and thermal calculation states of launch vehicles, calculates each state point. The specific calculation process is as follows: First, click on ballistic analysis to process the upstream standard ballistic file. Read out the three columns of data (time, altitude, and Mach number) from the ballistic file separately, remove duplicate time points, perform ballistic interpolation, and save it as a ballistic.dat file. Then, atmospheric parameters are solved. Then set the material parameters, including thickness, density, specific heat capacity, and thermal conductivity; Then set the initial temperature for calculation; Then perform von Neumann stability analysis iterative calculations and output the spatial step size of the difference scheme; Solve using the difference method; Finally, the temperature curve and the maximum temperature value are output.

7. An electronic device, characterized in that, include: processor; The memory is communicatively connected to the processor. At least one program is stored in the memory and configured to be executed by the processor, the at least one program being configured to: implement the rapid aerodynamic thermal calculation method for launch vehicles as described in any one of claims 1-6.

8. A computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by an electronic device, the computer program implements the rapid aerodynamic thermal calculation method for launch vehicles as described in any one of claims 1-6.

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