Method for calculating flame propagation process of solid engine

Abstract

The invention relates to a method for calculating a flame propagation process of a solid engine, which comprises the following steps of: judging a grid boundary state of a solid propellant, performing source item mass addition on a grid when the grid meets an ignition criterion condition, simulating surface ignition of the propellant, establishing a multi-factor driving ignition criterion, and introducing flame propagation process correction to calculate the flame propagation process of the solid engine. According to the method, control over the flow field flame propagation process in the solid engine is corrected, the problem that the flame propagation speed in a traditional calculation method is higher than the actual situation is solved, accurate prediction of the solid engine ignition transient flame propagation process is achieved, and technical support is provided for related research of solid engine flame propagation and the like.

Description

Technical Field

[0001] This invention relates to the field of numerical calculation of solid rocket motor flame propagation, and provides a method for calculating the flame propagation process of a solid rocket motor to simulate the flame propagation process. Background Technology

[0002] The reliability of solid rocket motors is one of the key factors affecting the success or failure of missile launches. Among them, the reliability of motor ignition is an important component of motor operational reliability, and the study of motor ignition criteria and flame propagation mechanisms is closely related to motor ignition reliability. With the continuous improvement of the performance requirements of solid rocket motors, high-energy propellants, high mass ratio, high pressure, high fill ratio, and full load have placed new and higher demands on ignition matching design. This requires mastering motor ignition criteria and flame propagation mechanisms to provide important technical support for the reliable and stable operation of motors.

[0003] Current flame propagation simulation studies are based on the ignition temperature criterion. However, this method alone is insufficient to represent the actual flame propagation process. The criterion of this method is to capture the temperature state of the first layer of the propellant surface, which is located in the gas phase region. When the propellant surface ignites, a large amount of high-temperature gas flow will be added to the gas phase region in the form of mass source terms, momentum source terms, and energy source terms, causing the temperature of the gas on the nearby unignited propellant surface to rise instantaneously above the ignition temperature, meeting the ignition requirements and determining that the propellant surface is ignited. The flame propagation speed calculated by this method is usually faster than the actual situation. Therefore, a more accurate calculation method for predicting the flame propagation process is needed. Summary of the Invention

[0004] Based on this, a method for calculating the flame propagation process of a solid rocket motor is provided, which solves the problem that the flame propagation speed in existing calculation methods is faster than the actual situation, and improves the accuracy of the flame propagation calculation method.

[0005] The present invention is achieved through the following technical solution.

[0006] A method for calculating the flame propagation process of a solid rocket motor includes the following steps:

[0007] Step S1: Based on the boundary state of the solid propellant grid, if the conditions of the ignition criterion are met, the grid is determined to be in an ignition state.

[0008] Step S2, based on traditional ignition criteria, introduces propellant oxygen-fuel ratio and ignition environment pressure parameters to propose a multi-factor driven ignition criterion; specifically including:

[0009] Based on relevant experimental data, and considering factors such as ignition heat flux density, different propellant types, ignition ambient pressure, ignition ambient temperature, and changes in propellant thermophysical properties, a thermobaric ignition criterion is established, grounded in propellant thermophysical properties and integrating ignition heat flux, propellant ambient pressure, and initial temperature characteristics. The ignition criterion is expressed as follows:

[0010]

[0011] In the formula, The surface temperature of the propellant; This is the temperature threshold constant; The mass ratio of oxidant to fuel; For ambient pressure; This is the pressure effect constant; Critical ignition criterion constant;

[0012] Step S3: Based on the multi-factor driven ignition criterion, a flame propagation process correction is introduced to modify the control of the flame propagation process in the internal flow field of the solid rocket motor, obtain a reasonable flame propagation rate correction coefficient, and introduce it into the calculation method to complete the simulation of the flame propagation process of the solid rocket motor.

[0013] Furthermore, step S1 specifically includes:

[0014] A grid model of the internal flow field of a solid rocket engine is established. During the flame propagation calculation, the grid boundary state of the solid propellant is judged in real time. When a certain grid on the propellant boundary meets the ignition criterion, the grid is determined to be in an ignition state. The grid is then mass-added with a source term to simulate the ignition state of the solid propellant surface.

[0015] Furthermore, the source term priming of the mesh includes:

[0016] The solid rocket motor flame propagation control equation is expressed as follows:

[0017]

[0018] Quality source item is represented as

[0019]

[0020] The momentum source term is represented as

[0021]

[0022] The energy source term is represented as

[0023]

[0024] The empirical formulas for combustion chamber pressure and combustion rate are expressed as follows:

[0025]

[0026] Gas quality source item is represented as

[0027]

[0028] The momentum source term of the gas in the x-direction is expressed as:

[0029] ,

[0030] The momentum source term of the gas in the y-direction is represented as

[0031] ,

[0032] The momentum source term of the gas in the z-direction is expressed as:

[0033] ,

[0034] in, , , These represent the injection velocities of the gas in the x, y, and z directions, respectively.

[0035] The gas energy source item is represented as

[0036] .

[0037] Furthermore, step S3 specifically includes:

[0038] Set initial parameters, including the flame dwell time reference t. ref Real physical time t flow The grid scale dx and time step dt are used. After the initial flow field, a cyclic calculation phase begins. In each iteration, values ​​are taken from all computational cells on the propellant surface to obtain the physical quantities required for the ignition criterion under multi-factor driving, including temperature and pressure. These are then substituted into the ignition criterion to determine whether the ignition conditions are met. If the conditions are met and adjacent surfaces are not burning, the surface is considered ignited. If adjacent surfaces are already burning, the ignition criterion and flame residence time Δt must be determined simultaneously. When the propellant surface state meets the ignition criterion and the flame residence time Δt ≥ t... ref If the propellant surface is ignited, the process continues until all propellant surfaces are in a burning state, thus completing the simulation of the entire flame propagation process.

[0039] Furthermore, the flame propagation speed v was measured experimentally. ref The mesh scale dx is obtained from the mesh model, and the flame dwell time reference t is calculated from both. ref, represented as

[0040] .

[0041] Compared with the prior art, the advantages of this invention are: it establishes a multi-factor driven ignition criterion based on the traditional ignition criteria, establishes a thermo-baric ignition criterion, introduces a correction coefficient to correct the flame propagation rate, and improves the effectiveness of predicting the flame propagation process in the internal flow field of solid rocket motors and the accuracy of predicting the ignition pressure build-up process. Attached Figure Description

[0042] Figure 1 This is a schematic diagram of the calculation method for the flame propagation process of a solid rocket motor according to the present invention;

[0043] Figure 2 This is a schematic diagram of the numerical method for the flame propagation process of a solid rocket motor according to the present invention;

[0044] Figure 3 This is a flowchart illustrating the solid rocket motor flame propagation model correction method of the present invention;

[0045] Figure 4 Images of flame propagation obtained from experimental measurements and numerical simulations of the solid rocket motor flame propagation process of this invention. Detailed Implementation

[0046] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not intended to limit the invention.

[0047] In one embodiment, such as... Figure 1 As shown, a method for calculating the flame propagation process of a solid rocket motor is provided, including the following steps:

[0048] Step S1: Based on the boundary state of the solid propellant grid, if the conditions of the ignition criterion are met, the grid is determined to be in an ignition state.

[0049] In step S1, such as Figure 2 As shown, a grid model of the internal flow field of a solid rocket engine is established. During the flame propagation calculation, the grid boundary state of the solid propellant is judged in real time. When a certain grid of the propellant boundary meets the ignition criterion, the source term is added to the grid to simulate the ignition of the solid propellant surface.

[0050] Specifically, the source term priming of the grid includes:

[0051] The control equation for flame propagation in a solid rocket motor is expressed as follows:

[0052]

[0053] Quality source item is represented as

[0054]

[0055] The momentum source term is represented as

[0056]

[0057] The energy source term is represented as

[0058]

[0059] The empirical formulas for combustion chamber pressure and combustion rate are expressed as follows:

[0060]

[0061] Gas quality source item is represented as

[0062]

[0063] The momentum source term of the gas in the x-direction is expressed as:

[0064] ,

[0065] The momentum source term of the gas in the y-direction is represented as

[0066] ,

[0067] The momentum source term of the gas in the z-direction is expressed as:

[0068] ,

[0069] in, , , These represent the injection velocities of the gas in the x, y, and z directions, respectively.

[0070] The gas energy source item is represented as

[0071]

[0072] Step S2: Based on the traditional ignition criteria, the propellant oxygen-fuel ratio and ignition environment pressure parameters are introduced to propose a multi-factor driven ignition criterion.

[0073] In step S2, based on relevant experimental data, considering ignition heat flux density, different propellant types, ignition ambient pressure, ignition ambient temperature, and changes in propellant thermophysical properties, a thermobaric ignition criterion is established by comprehensively considering the propellant thermophysical properties, ignition heat flux, propellant ambient pressure, and initial temperature characteristics. The ignition criterion is expressed as follows:

[0074]

[0075] In the formula, For propellant density; For propellant burning rate, This is the combustion rate coefficient. For pressure, Stress index; The area of ​​the propellant ignited; The density of the gas; The specific heat capacity of the propellant gas at constant pressure; The gas temperature; The surface temperature of the propellant; This is the temperature threshold constant; The mass ratio of oxidant to fuel; For ambient pressure; This is the pressure effect constant; Critical ignition criterion constant.

[0076] Step S3: Based on the multi-factor driven ignition criterion, a flame propagation process correction is introduced to correct the control of the flame propagation process in the internal flow field of the solid rocket motor, obtain a reasonable flame propagation rate correction coefficient, and introduce it into the calculation method to complete the simulation of the flame propagation process of the solid rocket motor.

[0077] In step S3, such as Figure 3 As shown, the flame dwell time reference duration t is set. ref (Obtained from experiments), actual physical time t flow Initial parameters such as grid scale dx and time step dt are used. After the flow field is initialized, a cyclic calculation phase begins. In each iteration, values ​​are taken from all computational cells on the propellant surface to obtain the physical quantities required for the ignition criterion under multi-factor driving, including temperature and pressure. These are then substituted into the ignition criterion to determine whether the ignition conditions are met. If the conditions are met and adjacent surfaces are unburned, the surface is considered ignited. If adjacent surfaces are already burning, the ignition criterion and flame residence time Δt must be determined simultaneously. The ignition is considered complete when the propellant surface state meets the ignition criterion and the flame residence time Δt ≥ t. ref If the propellant surface is ignited, the simulation is complete. This process continues until all propellant surfaces are in a burning state, thus simulating the entire flame propagation process.

[0078] Specifically, the flame propagation speed v was measured through experiments.ref The mesh scale dx is obtained from the mesh model, and the flame dwell time reference t is calculated from both. ref , represented as

[0079]

[0080] Specifically, the accuracy of the calculation method for the flame propagation process is verified, such as... Figure 4 As shown, the images of flame propagation obtained from experimental measurements and numerical simulations are presented. The flame propagation time in the numerical simulation is quite close to the experimental observation, and the error percentage remains at a relatively low level in all regions. Even in region 6, where the error is the largest, it is 11.49%, while the prediction errors for most other flame propagation regions are less than 10%.

[0081] Specifically, the combustion surface region times in the numerical simulation and experiments are shown in Table 1.

[0082] Table 1

[0083] area Test duration (ms) Simulation time (ms) error 1 0 0 0 2 2.4 2.5 4.17 % 3 5.6 5.3 -5.35 % 4 8.2 8 -2.44 % 5 11.2 10.4 -7.14 % 6 13.8 12.2 -11.59 % 7 15 13.6 -9.33 % 8 16 15.4 -3.75 %

[0084] The above description is merely a preferred embodiment of the present invention and does not limit the implementation and protection scope of the present invention. Those skilled in the art should realize that any equivalent substitutions and obvious changes made based on the description and illustrations of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for calculating the flame propagation process of a solid rocket motor, characterized in that, Includes the following steps: Step S1: Based on the boundary state of the solid propellant grid, if the conditions of the ignition criterion are met, the grid is determined to be in an ignition state. Step S2: Based on the traditional ignition criteria, the propellant oxygen-fuel ratio and ignition environment pressure parameters are introduced to propose a multi-factor driven ignition criterion. Specifically, it includes: Based on relevant experimental data, and considering factors such as ignition heat flux density, different propellant types, ignition ambient pressure, ignition ambient temperature, and changes in propellant thermophysical properties, a thermobaric ignition criterion is established, grounded in propellant thermophysical properties and integrating ignition heat flux, propellant ambient pressure, and initial temperature characteristics. The ignition criterion is expressed as follows: In the formula, The surface temperature of the propellant; This is the temperature threshold constant; The mass ratio of oxidant to fuel; Environmental pressure; This is the pressure effect constant; Critical ignition criterion constant; Step S3: Based on the multi-factor driven ignition criterion, a flame propagation process correction is introduced to modify the control of the flame propagation process in the internal flow field of the solid rocket motor, obtain a reasonable flame propagation rate correction coefficient, and introduce it into the calculation method to complete the simulation of the flame propagation process of the solid rocket motor.

2. The method for calculating the flame propagation process of a solid rocket motor according to claim 1, characterized in that, Step S1 specifically includes: A grid model of the internal flow field of a solid rocket engine is established. During the flame propagation calculation, the grid boundary state of the solid propellant is judged in real time. When a certain grid on the propellant boundary meets the ignition criterion, the grid is determined to be in an ignition state. The grid is then mass-added with a source term to simulate the ignition state of the solid propellant surface.

3. The method for calculating the flame propagation process of a solid rocket motor according to claim 2, characterized in that, The source term is primed on the grid. include: The solid rocket motor flame propagation control equation is expressed as follows: Quality source item is represented as The momentum source term is represented as The energy source term is represented as The empirical formulas for combustion chamber pressure and combustion rate are expressed as follows: Gas quality source item is represented as The momentum source term of the gas in the x-direction is expressed as: ， The momentum source term of the gas in the y-direction is represented as ， The momentum source term of the gas in the z-direction is expressed as: ， in, , , These represent the injection velocities of the gas in the x, y, and z directions, respectively. The gas energy source item is represented as 。 4. The method for calculating the flame propagation process of a solid rocket motor according to claim 1, characterized in that, Step S3 specifically includes: Set initial parameters, including the flame dwell time reference t. ref Real physical time t flow The grid scale dx and time step dt are used. After the initial flow field, the cyclic calculation phase begins. In each iteration, values ​​are taken for all computational cells on the propellant surface to obtain the physical quantities required for the ignition criterion under multi-factor driving, including temperature and pressure. These are substituted into the ignition criterion to determine whether the ignition conditions are met. If the conditions are met and adjacent surfaces are not burning, the surface is determined to be ignited. If adjacent surfaces are burning, the ignition criterion and flame residence time Δt need to be determined simultaneously. When the propellant surface state meets the ignition criterion and the flame residence time Δt ≥ t, the ignition criterion is satisfied. ref If the propellant surface is ignited, the process continues until all propellant surfaces are in a burning state, thus completing the simulation of the entire flame propagation process.

5. The method for calculating the flame propagation process of a solid rocket motor according to claim 1, characterized in that, The flame propagation speed v was measured experimentally. ref The mesh scale dx is obtained from the mesh model, and the flame dwell time reference t is calculated from both. ref , represented as 。

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