CN115730445A - Method, device and storage medium for predicting engine nozzle ablation - Google Patents

Abstract

The present application provides a prediction method, device and storage medium of engine nozzle ablation, the method comprising: calculating the initial inner ballistic performance parameter and the initial outer ballistic performance parameter of the spacecraft according to the initial nozzle ablation rate prediction model; The inner ballistic performance parameter and the initial outer ballistic performance parameter determine the parameter change range of the operating condition parameter during the flight of the spacecraft; the nozzle ablation rate simulation is carried out in the parameter change range to obtain the nozzle ablation rate approximate model; and the initial inner ballistic The ballistic performance parameters and the initial outer ballistic performance parameters are input into the nozzle ablation rate approximation model, and the target nozzle ablation prediction model of the spacecraft is obtained. The engine nozzle ablation prediction method, device and storage medium provided by this application calculate the inner ballistic parameters and outer ballistic parameters through a given initial nozzle ablation rate prediction model, and perform simulation based on these parameters, and finally obtain the target nozzle ablation rate prediction model. The tube ablation prediction model is closer to the ablation law of the nozzle during the actual flight of the spacecraft.

Description

Method, device and storage medium for predicting engine nozzle ablation

technical field

The present application relates to the field of engine technology, in particular, to a method, device and storage medium for predicting engine nozzle ablation.

Background technique

In the field of engines, especially in the field of spacecraft engines, with the continuous development of aerospace science and technology and the continuous improvement of human access to space, the design and development of advanced space vehicles has become the current research focus and hotspot. Solid rockets use solid rocket motors as the power system, which has many advantages such as simple structure, high reliability, short response time, and low difficulty in development. It is an important part of the launch vehicle system. Accurately evaluating the actual flight performance of solid rocket motors and fully exploiting the design potential of solid rocket motors are of great significance for improving the design level of solid rocket motors.

The traditional solid rocket motor design model usually includes drug shape design, internal ballistic performance calculation, component design and other components. The zero-dimensional internal ballistic equation is often used to calculate the change law of internal ballistic parameters such as engine combustion chamber pressure and thrust over time. Among them, the propulsion Key system parameters such as agent burning rate and nozzle ablation rate are given empirical values based on ground test results. This design process does not consider the impact of the rocket's outer ballistic flight process on the engine's inner ballistic process. Due to the deviation between the actual flight state of the rocket and the ground test state of the engine, there is also a certain deviation between the engine performance parameters calculated by the flight test and the calculation parameters of the engine design model based on the ground test. This deviation will affect the solid rocket design to a certain extent. Accurate evaluation of its engine performance and overall flight performance during the development phase.

As far as the prediction of nozzle throat ablation law is concerned, the throat ablation rate affects the change law of engine throat diameter, which leads to changes in parameters such as engine combustion chamber pressure, nozzle expansion ratio, and specific impulse efficiency, which in turn affects the engine internal pressure. Ballistic performance predicts accuracy. Relevant studies have shown that during actual flight, the ablation rate of the nozzle throat is significantly affected by the change of the outer ballistic flight overload. However, in the prior art, there is a certain deviation between the predicted results of the engine nozzle ablation law and the actual flight test results.

Contents of the invention

The purpose of the embodiments of the present application is a method, device and storage medium for predicting engine nozzle ablation. Through the prediction model of the initial nozzle ablation rate based on the prior art, the internal ballistic performance parameters and initial The outer ballistic performance parameters are then simulated based on the initial inner ballistic performance parameters and the initial outer ballistic performance parameters, so as to realize that the predicted results of the engine nozzle ablation law are closer to the actual flight test results.

In the first aspect, an embodiment of the present application provides a method for predicting engine nozzle ablation, including: calculating the initial inner ballistic performance parameters and initial outer ballistic performance parameters of the spacecraft according to the initial nozzle ablation rate prediction model; According to the initial internal ballistic performance parameter and the initial external ballistic performance parameter, the parameter variation range of the working condition parameter during the flight of the spacecraft is determined; the nozzle ablation rate simulation is carried out in the parameter variation range, and the nozzle ablation is obtained. rate approximation model; and inputting the initial inner ballistic performance parameters and initial outer ballistic performance parameters into the nozzle ablation rate approximation model to obtain the target nozzle ablation prediction model of the spacecraft.

The aforementioned predictive method of engine nozzle ablation, by calculating the initial internal ballistic performance parameters and initial external ballistic performance parameters, determines the variation range of the working condition parameters for simulating engine nozzle ablation, and then obtains the parameters through simulation The target nozzle ablation prediction model is obtained by using the nozzle ablation rate within the changing range, and introducing the initial inner ballistic performance parameters and the initial outer ballistic performance parameters. It solves the problem that in the prior art, through simple ground experiments, there is a deviation between the ablation law of the nozzle and the ablation law of the nozzle in the actual flight process due to the lack of consideration of the influencing factors in the external ballistic flight process. Furthermore, it provides more valuable guidance for the design of the engine.

With reference to the first aspect, optionally, wherein the operating condition parameters include engine parameters and flight overload parameters; determining the operating conditions of the spacecraft during flight according to the initial internal ballistic performance parameters and initial external ballistic performance parameters The parameter change range of the state parameters, including: determining the engine parameter change range during the flight process of the spacecraft according to the initial internal ballistic performance parameters; Variation range of overload parameters.

The prediction method of engine nozzle ablation mentioned above, by determining the engine parameter variation range and the flight overload parameter variation range, excludes other parameters that have little influence on the engine nozzle ablation law during the flight of the spacecraft, and reduces the impact on the nozzle burnout. The complexity of simulating the erosion rate and simplifying the calculation process of nozzle ablation predictive law.

In combination with the first aspect, optionally, wherein said performing nozzle ablation rate simulation within the parameter variation range to obtain an approximate nozzle ablation rate model includes: performing the spacecraft ablation rate simulation within the parameter variation range The two-phase flow simulation of the engine obtains the particle distribution law of the condensed phase particles of the engine corresponding to each of the working condition parameters within the parameter variation range; the particle distribution law is input into the nozzle throat lining material ablation model to obtain The first nozzle ablation rate corresponding to each working condition parameter; and constructing the nozzle ablation rate approximate model according to the first nozzle ablation rate corresponding to each working condition parameter.

The prediction method of engine nozzle ablation mentioned above obtains the particle distribution law of condensed phase particles through the two-phase flow simulation of the spacecraft engine, and further inputs the particle distribution law into the nozzle throat lining material ablation model, and finally obtains the The ablation rate of the first nozzle corresponding to the operating condition parameters is used to construct an approximate model of the ablation rate of the nozzle. This makes the final target nozzle ablation prediction model more accurate. With reference to the first aspect, optionally, wherein the performing the two-phase flow simulation of the spacecraft engine within the parameter variation range includes: constructing a three-dimensional flow field physical model of the engine nozzle; wherein, the The three-dimensional flow field physical model includes the free volume of the engine nozzle and the flow field area; based on the three-dimensional flow field physical model, the two-phase flow simulation of the spacecraft engine is performed within the parameter variation range.

The prediction method of the engine nozzle ablation mentioned above, by constructing a three-dimensional flow field physical model to simulate the two-phase flow of the spacecraft engine, makes the simulation calculation results more accurate, and finally improves the accuracy of the target nozzle ablation prediction model. Moreover, the three-dimensional flow field physical model can describe the characteristics of the real flow field, and is suitable for the two-phase flow simulation of spacecraft engines of different scales under various overload conditions.

With reference to the first aspect, optionally, wherein the performing the two-phase flow simulation of the spacecraft engine within the parameter variation range includes: using the kinetic equation of condensed-phase particles as a simulation model to perform the simulation of the spacecraft engine Two-phase flow simulation of the engine; the kinetic equation of the condensed phase particles is:

为颗粒速度矢量，t为时间，m_p为颗粒质量，

为曳力，

为压差力，

为过载引起的惯性力。in,

is the particle velocity vector, t is the time, m _p is the particle mass,

is the drag force,

is the differential pressure,

Inertial force caused by overload.

更准确直观地表达出了航天器发动机凝相粒子的粒子分布规律，最终提高了目标喷管烧蚀预示模型的准确性。The prediction method of engine nozzle ablation mentioned above, by combining the dynamics primitive equation of condensed phase particles and the influence of flight overload, introduces the inertial force parameter caused by overload

The particle distribution law of the condensed phase particles of the spacecraft engine is expressed more accurately and intuitively, and finally the accuracy of the target nozzle ablation prediction model is improved.

In conjunction with the first aspect, optionally, wherein the nozzle throat lining material ablation model includes an Oka erosion ratio model; the calculation formula of the Oka erosion ratio model is:

Among them, e _r is the erosion ratio of particles to the nozzle throat lining material, which indicates the relationship between the erosion ratio and the incident angle, e _ref represents the erosion ratio under the reference condition of normal incidence, U is the particle velocity, D is the particle diameter, U _ref and D _ref represent the velocity and diameter of particles under reference conditions, and k ₁ and k ₂ are empirical coefficients in the model;

The step of inputting the particle distribution law into the nozzle throat lining material ablation model to obtain the first nozzle ablation rate corresponding to the parameters of each working condition includes: inputting the diameter of the condensed phase particles and the particle distribution law Oka erosion ratio model to get the ablation rate of the first nozzle.

The prediction method of engine nozzle ablation mentioned above is simulated by the Oka erosion ratio model, which also more accurately and intuitively expresses the distribution law of the mechanical erosion of the spacecraft engine nozzle by condensed phase particles, and finally further improves the target nozzle ablation rate. The accuracy of the eclipse prediction model.

With reference to the first aspect, optionally, wherein the constructing the approximate model of the nozzle ablation rate according to the first nozzle ablation rate corresponding to the various working condition parameters includes: according to the corresponding first nozzle ablation rate of the various working condition parameters The first nozzle ablation rate, adopt polynomial fitting method to obtain the approximate model of the nozzle ablation rate corresponding to the working condition parameter range; the nozzle ablation rate approximate model is:

Among them, r is the ablation rate of the nozzle, the unit is mm/s; p _c is the pressure of the combustion chamber; M _W is the mole percentage of water in the gas phase product when the engine charge is burned; a _y is the lateral overload in flight; r ₀ Represents the nozzle ablation rate when there is no overload, the combustion chamber pressure is the reference value p _av , and the propellant formulation is the reference formulation.

The prediction method of the engine nozzle ablation mentioned above uses the polynomial fitting method to fuse the first nozzle ablation rate corresponding to each working condition parameter into a nozzle ablation rate approximation model, which facilitates subsequent input of the initial internal The ballistic performance parameters and the initial outer ballistic performance parameters are used to obtain the target nozzle ablation prediction model of the spacecraft.

In combination with the first aspect, optionally, wherein, after inputting the actual operating condition parameters of the spacecraft into the nozzle ablation rate approximation model and obtaining the target nozzle ablation prediction model of the spacecraft, the The method further includes: calculating the target inner ballistic performance parameters and target outer ballistic performance parameters of the spacecraft according to the target nozzle ablation prediction model; judging whether the target outer ballistic performance parameters converge; If the ballistic performance parameters do not converge, the target inner ballistic performance parameters and target outer ballistic performance parameters are input into the nozzle ablation rate approximation model to obtain the iterated nozzle ablation prediction model of the spacecraft.

The prediction method of the engine nozzle ablation mentioned above, through the current nozzle ablation prediction model, and repeated iterations of the inner ballistic performance parameters and outer elastic energy parameters calculated based on the nozzle ablation prediction model, until Nozzle ablation predicts convergence of external ballistic performance parameters calculated by the model. The final iterated nozzle ablation predictive model more accurately reflects the nozzle ablation law in the actual flight process, and further provides more valuable guidance for the design of the engine.

In the second aspect, the embodiment of the present application also provides a predictive device for engine nozzle ablation, the engine includes a spacecraft engine, and the device includes: a calculation module, a determination module, and a simulation module; wherein the calculation module It is used to calculate the initial inner ballistic performance parameters and the initial outer ballistic performance parameters of the spacecraft according to the initial nozzle ablation rate prediction model; the determination module is used to determine the The parameter change range of the operating condition parameter during the flight of the spacecraft; the simulation module is used to simulate the nozzle ablation rate within the parameter change range to obtain an approximate model of the nozzle ablation rate; and the simulation module is also used Inputting the initial inner ballistic performance parameter and the initial outer ballistic performance parameter into the nozzle ablation rate approximation model to obtain the target nozzle ablation prediction model of the spacecraft.

In the above embodiment, the prediction device of engine nozzle ablation provided has the same benefits as the prediction method of engine nozzle ablation provided in the above first aspect, or any optional implementation of the first aspect. effect, which will not be described here.

In a third aspect, the embodiment of the present application further provides a computer-readable storage medium, where a computer program is stored on the computer-readable storage medium, and when the computer program is run by a processor, the method described above is executed.

The computer-readable storage medium provided by the above-mentioned embodiment has the same beneficial effect as the prediction method of engine nozzle ablation provided by the above-mentioned first aspect, or any optional implementation of the first aspect. I won't go into details here.

In the third aspect, the embodiment of the present application also provides an electronic device, including: a processor and a memory, the memory stores machine-readable instructions executable by the processor, and when the machine-readable instructions are executed by the processor, they are executed as described above method.

The electronic device provided by the above embodiment has the same beneficial effect as the prediction method of engine nozzle ablation provided by the above first aspect, or any optional implementation of the first aspect, and will not be repeated here. .

To sum up, the prediction method, device and storage medium of engine nozzle ablation provided by this application calculate the internal ballistic performance parameters and initial The outer ballistic performance parameters are then simulated based on the initial inner ballistic performance parameters and the initial outer ballistic performance parameters, and the target nozzle ablation prediction model obtained is closer to the nozzle ablation law during the actual flight of the spacecraft. At the same time, the two-phase flow simulation method is used in the simulation, which can more accurately and intuitively reflect the particle distribution of the condensed phase particles of the spacecraft engine, so that the accuracy of the target nozzle ablation prediction model is improved. Finally, through repeated iterations, the target nozzle ablation prediction model is optimized, and the final iterated nozzle ablation prediction model more accurately reflects the nozzle ablation law in the actual flight process. Engine design provides even more valuable guidance.

Description of drawings

In order to more clearly illustrate the technical solutions of the embodiments of the present application, the accompanying drawings that need to be used in the embodiments of the present application will be briefly introduced below. It should be understood that the following drawings only show some embodiments of the present application, so It should not be regarded as a limitation on the scope, and those skilled in the art can also obtain other related drawings according to these drawings without creative work.

Fig. 1 is the first flow chart of the prediction method of engine nozzle ablation provided by the embodiment of the present application;

Fig. 2 is a detailed flow chart of step S120 in the prediction method of engine nozzle ablation provided by the embodiment of the present application;

Fig. 3 is a detailed flow chart of step S130 in the prediction method of engine nozzle ablation provided by the embodiment of the present application;

Fig. 4 is a detailed flow chart of step S131 in the predictive method of engine nozzle ablation provided by the embodiment of the present application;

Fig. 5 is the second flow chart of the prediction method of engine nozzle ablation provided by the embodiment of the present application;

Fig. 6 is a functional block diagram of a predictive device for engine nozzle ablation provided by an embodiment of the present application;

FIG. 7 is a schematic structural diagram of an electronic device provided by an embodiment of the present application;

Fig. 8 is the distribution diagram of condensed phase particles simulated when the transverse overload is 0 provided by the embodiment of the present application;

Fig. 9 is the distribution diagram of condensed phase particles simulated when the transverse overload provided by the embodiment of the present application is 5g;

Fig. 10 is the distribution diagram of condensed phase particles simulated when the transverse overload is 10g provided by the embodiment of the present application;

Fig. 11 is the distribution diagram of the mechanical erosion rate of the nozzle simulated when the lateral overload provided by the embodiment of the present application is 0;

Fig. 12 is the distribution diagram of the mechanical erosion rate of the nozzle simulated when the transverse overload provided by the embodiment of the present application is 5g;

Fig. 13 is a distribution diagram of the simulated mechanical erosion rate of the nozzle provided by the embodiment of the present application when the lateral overload is 10g.

Detailed ways

Embodiments of the technical solutions of the present application will be described in detail below in conjunction with the accompanying drawings. The following examples are only used to illustrate the technical solution of the present application more clearly, and therefore are only examples, rather than limiting the protection scope of the present application.

Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by those skilled in the technical field of the application; the terms used herein are only for the purpose of describing specific embodiments, and are not intended to Limit this application.

In the description of the embodiments of the present application, the technical terms "first", "second" and so on are only used to distinguish different objects, and should not be understood as indicating or implying relative importance or implicitly indicating the number of indicated technical features, A specific order or primary-secondary relationship. In the description of the embodiments of the present application, "plurality" means two or more, unless otherwise specifically defined.

Based on the description of the prior art in the aforementioned background technology section, the applicant found that by considering the coupling mechanism of internal and external ballistics, establishing a solid rocket motor nozzle ablation prediction model with good consistency between the sky and the ground will help to make the design and development of the solid rocket more efficient. Accurately calculates ballistic performance within the engine.

The main mechanism for the ablation rate of the throat of the nozzle to be affected by the external flight overload is that the external ballistic flight overload acts on the two-phase flow field in the solid rocket motor, causing the movement trajectory of the condensed phase particles in the engine to deviate, and in the combustion chamber and nozzle Part of the throat area gathers, and the concentration of condensed phase particles increases. That is to say, a two-phase flow simulation method of solid rocket motor under flight overload is given. The above-mentioned changes in the movement law of the condensed phase components increase the mechanical erosion of the surface of the throat lining material, resulting in an increase in the ablation rate of the nozzle throat lining, and the phenomenon of burning deviation with the direction of overload.

In the prior art scheme, although a solid rocket motor design model based on ground tests has been established, including an empirical model predicting the nozzle ablation rate, and the theory of the solid rocket motor nozzle ablation law under the influence of external ballistic flight overload is carried out. Analysis and experimental research, but it still has shortcomings and is mainly reflected in the following two aspects:

On the one hand, the nozzle ablation rate in the existing solid rocket motor design model is generally given based on the engineering experience of the ground test. For example, the average linear ablation rate is calculated through the throat diameter change before and after the test, and there are no factors affecting the external ballistic flight process. (mainly flight overload) considerations, there is a certain deviation between the prediction results of the engine nozzle ablation law and the actual flight test results;

On the other hand, the existing research on the nozzle ablation law under the influence of overload usually only conducts theoretical analysis and experimental exploration at the mechanism level, and the relevant influence laws formed through simulation and testing are not closely integrated with the engine design and development process, and lack consideration. A predictive model of nozzle ablation for the coupled effects of internal and external ballistics can be used to guide the engine design process.

Aiming at the defects in the above two aspects, the present application provides a method, device and storage medium for predicting engine nozzle ablation, so as to solve the above defects in the prior art. Specifically, please refer to the embodiments and drawings provided in this application.

Please refer to FIG. 1 . FIG. 1 is a first flow chart of a method for predicting engine nozzle ablation provided by an embodiment of the present application. The prediction method of engine nozzle ablation provided in the embodiment of the present application includes:

Step S110: Calculate initial inner ballistic performance parameters and initial outer ballistic performance parameters of the spacecraft according to the initial nozzle ablation rate prediction model.

In the above step S110, the initial nozzle ablation prediction model may be given based on the engineering experience of the engine ground test. According to the initial nozzle ablation prediction model, the internal ballistic performance of the engine is calculated, and the initial internal ballistic performance parameters of the spacecraft are obtained. Initial internal ballistic performance parameters include, but are not limited to: engine thrust-time curves and mass change-time curves. The initial outer ballistic performance parameter can be calculated according to the initial inner ballistic performance parameter. Specifically, it can be implemented as inputting the initial internal ballistic performance parameters into the solid rocket external ballistic design model in the prior art to calculate the external ballistic performance of the spacecraft to obtain the initial external ballistic performance parameters. The initial outer ballistic performance parameters include, but are not limited to: flight speed, flight overload parameters, shutdown point height parameters, and ballistic inclination parameters.

Step S120: According to the initial inner ballistic performance parameter and the initial outer ballistic performance parameter, determine the parameter variation range of the operating condition parameter during the flight of the spacecraft.

In the above step S120, the variation range of the operating condition parameters includes but not limited to: the variation range of the engine thrust of the spacecraft during flight, the variation range of the spacecraft mass parameter, the variation range of the flight speed, the variation range of the flight overload parameter and the ballistic inclination angle parameter range of change.

Step S130: Perform nozzle ablation rate simulation within the parameter variation range to obtain an approximate model of nozzle ablation rate.

In the above step S130, the ablation rate of the nozzle refers to the law of ablation of the engine nozzle of the spacecraft during flight. According to the variation range of the operating condition parameters determined in the aforementioned step S120, the nozzle ablation rate of the engine during the flight of the spacecraft can be simulated to obtain an approximate model of the nozzle ablation rate.

Step S140: Input the initial inner ballistic performance parameters and the initial outer ballistic performance parameters into the nozzle ablation rate approximation model to obtain the target nozzle ablation prediction model of the spacecraft.

In the above step S140, the initial inner ballistic performance parameter and the initial outer ballistic performance parameter are input into the above-mentioned nozzle ablation rate approximation model as input parameters, and the output of the model is the nozzle ablation predictive law of the spacecraft, namely: the target nozzle ablation rate Tube ablation predictor model.

In the above implementation process, by calculating the initial internal ballistic performance parameters and initial external ballistic performance parameters, the range of variation of the working condition parameters for the simulation of the spacecraft flight process is determined, and then the nozzle within the range of the parameter variation is obtained through simulation. Ablation rate, the initial inner ballistic performance parameters and the initial outer ballistic performance parameters are introduced to obtain the target nozzle ablation prediction model. It solves the problem that in the prior art, through simple ground experiments, there is a deviation between the ablation law of the nozzle and the ablation law of the nozzle in the actual flight process due to the lack of consideration of the influencing factors in the external ballistic flight process. Furthermore, it provides more valuable guidance for the design of the engine.

Please refer to Fig. 2, Fig. 2 is a detailed flow chart of step S120 in the prediction method of engine nozzle ablation provided by the embodiment of the present application. In an optional embodiment, the working condition parameters include engine parameters and flight overload parameters .

Above-mentioned step S120 comprises:

Step S121: Determine the variation range of the engine parameters during the flight of the spacecraft according to the initial internal ballistic performance parameters.

Step S122: According to the initial outer ballistic performance parameters, the variation range of the flight overload parameters during the flight of the spacecraft can be determined.

In the above steps, the engine parameters in the initial inner ballistic performance parameters and the flight overload parameters in the initial outer ballistic performance parameters are the most important parameters affecting the engine nozzle ablation law during the flight of the spacecraft.

In the above implementation process, by determining the variation range of the engine parameters and the flight overload parameters, other parameters that have little influence on the ablation law of the engine nozzle during the flight of the spacecraft are excluded, and the simulation of the nozzle ablation rate is reduced. complexity, and simplify the calculation process of nozzle ablation predictive law.

Please refer to FIG. 3. FIG. 3 is a detailed flow chart of step S130 in the method for predicting engine nozzle ablation provided by the embodiment of the present application. In an optional implementation manner, the above step S130 includes:

Step S131: Carry out the two-phase flow simulation of the spacecraft engine within the parameter variation range, and obtain the particle distribution law of the engine condensed phase particles corresponding to each working condition parameter within the parameter variation range.

In the above step S131, relevant research in the prior art in the art shows that longitudinal overload has little influence on the ablation rate of the nozzle throat, and lateral overload has a more obvious influence on the ablation rate of the nozzle throat. As the value increases, the change of ablation rate is more significant. Therefore, in the embodiment of the present application, in order to clearly show the change of nozzle throat ablation law when considering or not considering the coupling of internal and external ballistics, it can be considered that only the variation range of the lateral flight overload of the external ballistic is given as the development of solid rocket motor internal The range of operating conditions calculated by numerical simulation of the flow field. Through the simulation, the particle distribution law of the engine condensed phase particles corresponding to each working condition parameter can be obtained.

Step S132: Input the particle distribution law into the nozzle throat lining material ablation model to obtain the first nozzle ablation rate corresponding to each working condition parameter.

In the above step S132, the nozzle throat lining material ablation model includes but not limited to the Oka erosion ratio model, the Neilson model and the Gilchrist model, and the particle distribution law corresponding to each parameter is input into the nozzle throat lining material ablation model, and the model outputs The ablation rate of the first nozzle corresponding to each working condition parameter.

Step S133: Constructing an approximate model of the nozzle ablation rate according to the first nozzle ablation rate corresponding to each working condition parameter.

In the above-mentioned step S133, since the ablation rate of the first nozzle corresponding to each working condition parameter is obtained, then the nozzle corresponding to the state corresponding to the different working condition parameters during the flight of the spacecraft can also be obtained. Ablation rate. Therefore, by constructing the first nozzle ablation rate corresponding to the above working condition parameters into an approximate model of the nozzle ablation rate, and inputting the corresponding initial inner ballistic performance parameters and initial outer ballistic performance parameters, the aerospace The predictive model of target nozzle ablation in the corresponding flight state.

In the above implementation process, the particle distribution law of the condensed phase particles is obtained through the two-phase flow simulation of the spacecraft engine, and the particle distribution law is further input into the ablation model of the nozzle throat lining material, and finally the parameters corresponding to each working condition are obtained. The ablation rate of the first nozzle constructs an approximate model of the ablation rate of the nozzle. This makes the final target nozzle ablation prediction model more accurate.

Please refer to FIG. 4 , which is a detailed flowchart of step S131 in the method for predicting engine nozzle ablation provided by the embodiment of the present application. In an optional implementation manner, the above step S131 includes:

Step S1311: Construct a three-dimensional flow field physical model of the engine nozzle. Among them, the three-dimensional flow field physical model includes the free volume of the engine nozzle and the flow field area.

In the above step S1311, the three-dimensional flow field physical model includes the free volume in the combustion chamber and the flow field area in the nozzle. In practical applications, the physical model of the 3D flow field can be appropriately simplified depending on the situation, and the calculation efficiency can be improved without affecting the calculation results. For example, when the physical characteristics of the 3D flow field are geometrically symmetrical and the overload has no lateral component , half of the physical model of the three-dimensional flow field can be used; if the expansion of the nozzle is relatively large, it can also be shortened appropriately.

Step S1312: Carry out two-phase flow simulation of the spacecraft engine within the parameter range based on the three-dimensional flow field physical model.

In the above step S1312, the hexahedral structured grid can be used to discrete the simulation area, and the boundary conditions of the simulation area can be determined according to the operating condition parameters of the spacecraft engine in actual operation, and the combustion products of the propellant can be regarded as two parts: gas phase and condensed phase particles , respectively set the boundary conditions of the simulation for these two parts. The two-phase flow field is calculated and solved by the Euler-Lagrangian method, that is, the gas phase is regarded as a continuous phase, and its governing equation is expressed in Euler form, while the condensed phase particles are regarded as a discrete phase, which is solved in Lagrange coordinates . There is a transfer of momentum and energy between gas phase and condensed phase particles, which affects gas phase flow and particle trajectory. Using a certain size distribution law of condensed phase particles and certain gas phase parameters, the simulation calculation of the two-phase flow field of the target solid rocket motor can be carried out to obtain the movement and distribution law of the particles in the flow field.

In the above implementation process, the two-phase flow simulation of the spacecraft engine is carried out by building a three-dimensional flow field physical model, which makes the simulation calculation results more accurate, and finally improves the accuracy of the target nozzle ablation prediction model. Moreover, the three-dimensional flow field physical model can describe the characteristics of the real flow field, and is suitable for the two-phase flow simulation of spacecraft engines of different scales under various overload conditions.

In an optional implementation manner, the above S131 also includes:

Step S1313: Using the kinetic equation of condensed phase particles as a simulation model to perform two-phase flow simulation of the spacecraft engine. The kinetic equation of the condensed phase particle is:

为颗粒速度矢量，t为时间，m_p为颗粒质量，

为曳力，

为压差力，

为过载引起的惯性力。in,

is the particle velocity vector, t is the time, m _p is the particle mass,

is the drag force,

is the differential pressure,

Inertial force caused by overload.

In the above step S1313, the variation range of the lateral flight overload can be given as 0-10g, and the numerical simulation of the two-phase flow of the solid rocket motor considering the influence of the overload can be carried out. Based on the high-precision drag force model and the DPM (Discrete Phase Method) model established by numerical simulation of particle motion, the distribution of condensed phase particles in the engine in the flow field is obtained through simulation.

During the simulation process, it is assumed that the condensed phase particles are all spherical, and the splitting and merging between particles are ignored, and multiple discrete values of different diameters are used to replace the real particle size continuous distribution to simplify the calculation. Using the DPM model, the drag coefficient model is input in the form of UDF (custom function) and used to calculate the drag force on the particles in the supersonic flow of the nozzle. The collision model is adopted on the boundary, and the particles rebound with a certain restitution coefficient (set to 0.8 in this embodiment) after colliding with the wall of the nozzle. The original kinetic equation of condensed phase particles is as follows:

为颗粒速度矢量，t为时间，m_p为颗粒质量，

为曳力，

为压差力，

为颗粒位移矢量，若受到飞行过载影响，则凝相颗粒的动力学方程变为：in,

is the particle velocity vector, t is the time, m _p is the particle mass,

is the drag force,

is the differential pressure,

is the particle displacement vector, if affected by the flight overload, the dynamic equation of the condensed phase particle becomes:

为过载引起的惯性力。in,

Inertial force caused by overload.

Please refer to Figures 8 to 10 in combination. Figure 8 is the distribution diagram of condensed phase particles simulated when the lateral overload provided by the embodiment of the present application is 0; Phase particle distribution diagram; FIG. 10 is the simulated condensed phase particle distribution diagram provided by the embodiment of the present application when the lateral overload is 10g. It can be seen from the simulation results that when there is no lateral overload, the particles gather at the center of the engine in a symmetrical and long shape, with a maximum concentration of 11.2kg/m ³ , and this high-concentration "strip" will not hit the wall of the nozzle. After adding 5g of lateral overload, the maximum particle concentration dropped to 3.22kg/m ³ , and there was obvious divergence and upward deflection, and hit the position between the convergent section of the nozzle and the throat. After the lateral overload reaches 10g, the upward deflection of the particle distribution is very obvious, and the accumulation flows on the wall in the opposite direction of the external overload, and the maximum concentration rises to 7.33kg/m ³ , and the erosion range of the impact nozzle wall is wider, but not as good as the lateral overload of 5g Time to concentrate.

更准确直观地表达出了航天器发动机凝相粒子的粒子分布规律，最终提高了目标喷管烧蚀预示模型的准确性。In the above implementation process, by combining the original equation of dynamics of condensed phase particles and the influence of flight overload, the inertial force parameter caused by overload is introduced

The particle distribution law of the condensed phase particles of the spacecraft engine is expressed more accurately and intuitively, and finally the accuracy of the target nozzle ablation prediction model is improved.

In an optional embodiment, the nozzle throat lining material ablation model includes an Oka erosion ratio model, and the calculation formula of the Oka erosion ratio model is:

Among them, e _r is the erosion ratio of particles to the nozzle throat lining material, which indicates the relationship between the erosion ratio and the incident angle, e _ref represents the erosion ratio under the reference condition of normal incidence, U is the particle velocity, D is the particle diameter, U _ref and D _ref represent the velocity and diameter of the particle under reference conditions, and k ₁ and k ₂ are empirical coefficients in the model.

Correspondingly, the above step S132 includes:

Step S1321: Input the diameter and particle distribution law of the condensed phase particles into the Oka erosion ratio model to obtain the ablation ratio of the first nozzle.

Please refer to Figures 11 to 13 in conjunction. Figure 11 is a distribution diagram of the mechanical erosion rate of the nozzle provided by the embodiment of the application when the lateral overload is 0; Figure 12 is a simulation of the lateral overload provided by the embodiment of the application. Figure 13 is a distribution diagram of the mechanical erosion rate distribution of the nozzle provided by the embodiment of the present application when the lateral overload is 10g. In the above step S1321, it can be seen from Fig. 11 to Fig. 13 that mechanical erosion does not occur in the upstream of the converging section and the diverging section of the nozzle. Mechanical erosion is mainly concentrated in the converging section and near the throat. The maximum mechanical erosion reached 0.1722kg/(m ² ·s) when the lateral overload was 5g. The local erosion rate increases significantly along the direction opposite to the external overload. The erosion rates in other directions are concentrated in the range of about 0.04-0.07kg/(m ² ·s). When the lateral overload is 10g, the maximum value of mechanical erosion reaches 0.2378kg/(m ² ·s), the local erosion rate increases along the direction opposite to the external overload, and the erosion area of the convergent section of the nozzle increases and is linear to the direction of the combustion chamber. extend. The erosion rates in the other directions of the nozzle are also concentrated in the range of about 0.04-0.07kg/(m ² ·s).

In the above implementation process, the Oka erosion ratio model was used to simulate the distribution law of the mechanical erosion of the nozzle of the spacecraft engine by the condensed phase particles, and finally further improved the accuracy of the target nozzle ablation prediction model. sex.

In an optional implementation manner, the above step S133 includes:

Step S1331: According to the first nozzle ablation rate corresponding to each working condition parameter, the approximate model of the nozzle ablation rate corresponding to the working condition parameter range is obtained by using a polynomial fitting method. The approximate model of the nozzle ablation rate is:

Among them, r is the ablation rate of the nozzle, the unit is mm/s; p _c is the pressure of the combustion chamber; M _W is the mole percentage of water in the gas phase product when the engine charge is burned; a _y is the lateral overload in flight; r ₀ Represents the nozzle ablation rate when there is no overload, the combustion chamber pressure is the reference value p _av , and the propellant formulation is the reference formulation.

In the above implementation process, the ablation rate of the first nozzle corresponding to each working condition parameter is fused into an approximate model of nozzle ablation rate by polynomial fitting method, which facilitates the subsequent input of initial internal ballistic performance parameters and initial The outer ballistic performance parameters are obtained to predict the ablation model of the target nozzle of the spacecraft.

Please refer to FIG. 5 . FIG. 5 is a second flow chart of the prediction method for engine nozzle ablation provided by the embodiment of the present application. In an optional implementation manner, after the above step S140, the method for predicting engine nozzle ablation provided in the embodiment of the present application further includes:

Step S150: Calculate the target inner ballistic performance parameters and target outer ballistic performance parameters of the spacecraft according to the target nozzle ablation predictive model.

Step S160: Judging whether the ballistic performance parameters outside the target converge.

If it is determined that the target outer ballistic performance parameters do not converge, step S170 is performed: input the target inner ballistic performance parameters and target outer ballistic performance parameters into the nozzle ablation rate approximation model to obtain the iterated nozzle ablation prediction model of the spacecraft.

In the above steps, the target inner ballistic performance parameters and the target outer ballistic performance parameters are calculated by using the target nozzle ablation prediction model, and it is judged whether the target outer ballistic performance parameters converge. If not, continue based on the target inner ballistic performance parameters and An iterative post-nozzle ablation predictive model for the calculation of off-target ballistic performance parameters for a spacecraft. It should be understood that the iterative nozzle ablation prediction model can be verified again by using substantially the same method as the above steps, and through such an iterative process, until the outer ballistic performance parameters converge.

In the above implementation process, through the currently obtained nozzle ablation prediction model, and repeated iterations of the inner ballistic performance parameters and outer elastic energy parameters calculated based on the nozzle ablation prediction model, until the nozzle ablation prediction The exoelastic energy path parameters calculated by the model converge. The final iterated nozzle ablation predictive model more accurately reflects the nozzle ablation law in the actual flight process, and further provides more valuable guidance for the design of the engine.

Please refer to FIG. 6 , which is a functional block diagram of an engine nozzle ablation predicting device 600 provided in an embodiment of the present application. An engine nozzle ablation prediction device 600 provided in an embodiment of the present application includes: a calculation module 610 , a determination module 620 and a simulation module 630 .

Wherein, the calculation module 610 is used to calculate the initial inner ballistic performance parameters and the initial outer ballistic performance parameters of the spacecraft according to the initial nozzle ablation rate prediction model. The determining module 620 is used to determine the parameter variation range of the operating condition parameter during the flight of the spacecraft according to the inner ballistic performance parameter and the outer ballistic performance parameter. The simulation module 630 is used to simulate the nozzle ablation rate within the parameter variation range to obtain an approximate model of the nozzle ablation rate. The simulation module 630 is also used to input the initial inner ballistic performance parameter and the initial outer ballistic performance parameter into the nozzle ablation rate approximation model to obtain the target nozzle ablation prediction model of the spacecraft.

Please continue to refer to FIG. 6 , in an optional implementation manner, the above operating condition parameters include engine parameters and flight overload parameters.

Correspondingly, the above-mentioned determining module 620 is specifically used for:

According to the initial internal ballistic performance parameters, the variation range of the engine parameters during the flight of the spacecraft is determined; according to the initial external ballistic performance parameters, the range of the flight overload parameters during the flight of the spacecraft can be determined.

Please continue to refer to FIG. 6, an optional implementation manner, the above-mentioned simulation module 630 is specifically used for:

The two-phase flow simulation of the spacecraft engine is carried out within the parameter variation range, and the particle distribution law of the engine condensed phase particles corresponding to each working condition parameter within the parameter variation range is obtained; the particle distribution law is input into the ablation model of the nozzle throat lining material, and the The ablation rate of the first nozzle corresponding to each working condition parameter; and constructing an approximate model of the ablation rate of the nozzle according to the ablation rate of the first nozzle corresponding to each working condition parameter.

Please continue to refer to FIG. 6, an optional implementation manner, the above-mentioned simulation module 630 is also specifically used for:

Construct the three-dimensional flow field physical model of the engine nozzle; the three-dimensional flow field physical model includes the free volume of the engine nozzle and the flow field area; based on the three-dimensional flow field physical model, the two-phase flow simulation of the spacecraft engine is performed within the range of parameter changes .

Please continue to refer to FIG. 6, an optional implementation manner, the above-mentioned simulation module 630 is also specifically used for:

The dynamic equation of condensed phase particles is used as the simulation model to simulate the two-phase flow of the spacecraft engine; the dynamic equation of condensed phase particles is:

为颗粒速度矢量，t为时间，m_p为颗粒质量，

为曳力，

为压差力，

为过载引起的惯性力。in,

is the particle velocity vector, t is the time, m _p is the particle mass,

is the drag force,

is the differential pressure,

Inertial force caused by overload.

Please continue to refer to Fig. 6, in an optional embodiment, the above-mentioned nozzle throat lining material ablation model includes the Oka erosion ratio model; the calculation formula of the Oka erosion ratio model is:

Among them, e _r is the erosion ratio of particles to the nozzle throat lining material, which indicates the relationship between the erosion ratio and the incident angle, e _ref represents the erosion ratio under the reference condition of normal incidence, U is the particle velocity, D is the particle diameter, U _ref and D _ref represent the velocity and diameter of the particle under reference conditions, and k ₁ and k ₂ are empirical coefficients in the model.

Correspondingly, the above-mentioned simulation module 630 is also specifically configured to input the diameter and particle distribution law of the condensed phase particles into the Oka erosion ratio model to obtain the ablation rate of the first nozzle corresponding to each working condition parameter.

Please continue to refer to FIG. 6, an optional implementation manner, the above-mentioned simulation module 630 is also specifically used for:

According to the first nozzle ablation rate corresponding to each working condition parameter, adopt the polynomial fitting method to obtain the nozzle ablation rate approximate model corresponding to the working condition parameter range; the nozzle ablation rate approximate model is:

Among them, r is the ablation rate of the nozzle, the unit is mm/s; p _c is the pressure of the combustion chamber; M _W is the mole percentage of water in the gas phase product when the engine charge is burned; a _y is the lateral overload in flight; r ₀ Represents the nozzle ablation rate when there is no overload, the combustion chamber pressure is the reference value p _av , and the propellant formulation is the reference formulation.

Please continue to refer to FIG. 6, an optional implementation manner, the calculation module 610 is also used to calculate the target internal ballistic performance parameters and target external ballistic performance parameters of the spacecraft according to the target nozzle ablation prediction model; and judge the target external trajectory Whether the performance parameters converge.

If it is determined that the target outer ballistic performance parameters do not converge, correspondingly, the above-mentioned simulation module 630 is also used to input the target inner ballistic performance parameters and target outer ballistic performance parameters into the nozzle ablation rate approximation model to obtain the iterated nozzle of the spacecraft Ablation Prediction Model.

It should be understood that this device corresponds to the above-mentioned embodiment of the prediction method for engine nozzle ablation, and can perform various steps involved in the above-mentioned method embodiment. For the specific functions of the device, please refer to the description above. To avoid repetition, here Detailed description is omitted here. The device includes at least one software function module that can be stored in a memory in the form of software or firmware (firmware) or solidified in an operating system (operating system, OS) of the device.

Based on the same inventive concept, please refer to FIG. 7 , which is a schematic structural diagram of an electronic device 700 provided in an embodiment of the present application. The electronic device 700 may include a memory 711 , a storage controller 712 , a processor 713 , a peripheral interface 714 , an input and output unit 717 , and a display unit 716 . Those skilled in the art can understand that the structure shown in FIG. 7 is only for illustration, and it does not limit the structure of the electronic device 700 . For example, the electronic device 700 may also include more or fewer components than shown in FIG. 7 , or have a different configuration than that shown in FIG. 7 .

The aforementioned memory 711 , storage controller 712 , processor 713 , peripheral interface 714 , input/output unit 717 and display unit 716 are electrically connected to each other directly or indirectly to realize data transmission or interaction. For example, these components can be electrically connected to each other through one or more communication buses or signal lines. The aforementioned processor 713 is used to execute the executable modules stored in the memory.

Wherein, the memory 711 can be, but not limited to, random access memory (Random Access Memory, referred to as RAM), read-only memory (Read Only Memory, referred to as ROM), programmable read-only memory (Programmable Read-Only Memory, referred to as PROM) , Erasable Programmable Read-Only Memory (EPROM for short), Electric Erasable Programmable Read-Only Memory (EEPROM for short), etc. Wherein, the memory 711 is used to store the program, and the processor 713 executes the program after receiving the execution instruction, and the method performed by the electronic device 700 according to the process definition disclosed in any embodiment of the present application can be applied to processing In the device 713, or implemented by the processor 713.

The above-mentioned processor 713 may be an integrated circuit chip with signal processing capabilities. The above-mentioned processor 713 can be a general-purpose processor, including a central processing unit (Central Processing Unit, referred to as CPU), a network processor (Network Processor, referred to as NP), etc.; it can also be a digital signal processor (digital signal processor, referred to as DSP) , Application Specific Integrated Circuit (ASIC for short), Field Programmable Gate Array (FPGA) or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components. Various methods, steps, and logic block diagrams disclosed in the embodiments of the present application may be implemented or executed. A general-purpose processor may be a microprocessor, or the processor may be any conventional processor, or the like.

The aforementioned peripheral interface 714 couples various input/output devices to the processor 713 and the memory 711 . In some embodiments, peripheral interface 714, processor 713, and memory controller 712 may be implemented in a single chip. In some other instances, they can be implemented by independent chips respectively.

The aforementioned input and output unit 717 is used to provide the user with input data. The input and output unit 717 may be, but not limited to, a mouse and a keyboard.

The above-mentioned display unit 716 provides an interactive interface (such as a user operation interface) between the electronic device 700 and the user or is used to display image data for the user's reference. In this embodiment, the display unit may be a liquid crystal display or a touch display. If it is a touch display, it can be a capacitive touch screen or a resistive touch screen supporting single-point and multi-touch operations. Supporting single-point and multi-touch operations means that the touch display can sense simultaneous touch operations from one or more positions on the touch display, and hand over the sensed touch operations to the processor calculation and processing.

The electronic device 700 in this embodiment may be used to execute each step in each method provided in the embodiment of this application.

The embodiment of the present application also provides a storage medium, on which a computer program is stored, and when the computer program is run by a processor, the above method is executed.

Wherein, the storage medium can be realized by any type of volatile or non-volatile storage device or their combination, such as Static Random Access Memory (Static Random Access Memory, referred to as SRAM), Electrically Erasable Programmable Read-Only Memory (Electrically Erasable Programmable Read-Only Memory, referred to as EEPROM), Erasable Programmable Read-Only Memory (Erasable Programmable Read Only Memory, referred to as EPROM), Programmable Read-Only Memory (Programmable Red-Only Memory, referred to as PROM), read-only Memory (Read-Only Memory, ROM for short), magnetic memory, flash memory, magnetic disk or optical disk.

To sum up, the prediction method, device and storage medium of engine nozzle ablation provided by this application calculate the internal ballistic performance parameters and initial The outer ballistic performance parameters are simulated based on the initial inner ballistic performance parameters and the initial outer ballistic performance parameters, and the obtained target nozzle ablation prediction model is closer to the ablation law of the nozzle during the actual flight of the spacecraft. At the same time, the two-phase flow simulation method is used in the simulation, which can more accurately and intuitively reflect the particle distribution of the condensed phase particles of the spacecraft engine, so that the accuracy of the target nozzle ablation prediction model is improved. Finally, through repeated iterations, the target nozzle ablation prediction model is optimized, and the final iterated nozzle ablation prediction model more accurately reflects the nozzle ablation law in the actual flight process. Engine design provides even more valuable guidance.

In the several embodiments provided in the embodiments of the present application, it should be understood that the disclosed devices and methods may also be implemented in other ways. The device embodiments described above are only illustrative. For example, the flowcharts and block diagrams in the accompanying drawings show possible implementation architectures of devices, methods, and computer program products according to multiple embodiments of the embodiments of the present application. function and operation. In this regard, each block in a flowchart or block diagram may represent a module, program segment, or portion of code that contains one or more executable instruction. It should also be noted that, in some alternative implementations, the functions noted in the block may occur out of the order noted in the figures. For example, two blocks in succession may, in fact, be executed substantially concurrently, or they may sometimes be executed in the reverse order, depending upon the functionality involved. It should also be noted that each block of the block diagrams and/or flowchart illustrations, and combinations of blocks in the block diagrams and/or flowchart illustrations, can be implemented by a dedicated hardware-based system that performs the specified function or action , or may be implemented by a combination of dedicated hardware and computer instructions.

In addition, each functional module in each embodiment of the embodiment of the present application may be integrated to form an independent part, each module may exist independently, or two or more modules may be integrated to form an independent part.

The above description is only an optional implementation of the embodiment of the present application, but the scope of protection of the embodiment of the present application is not limited thereto. Anyone familiar with the technical field can Changes or substitutions that can easily be thought of should fall within the scope of protection of the embodiments of the present application.

Claims (10)

1. A method of indicating nozzle erosion in an engine, the method comprising:

calculating an initial inner ballistic performance parameter and an initial outer ballistic performance parameter of the spacecraft according to an initial nozzle ablation rate prediction model;

determining a parameter variation range of working condition parameters of the spacecraft in the flying process according to the initial inner ballistic performance parameters and the initial outer ballistic performance parameters;

simulating the ablation rate of the spray pipe in the parameter change range to obtain an approximate model of the ablation rate of the spray pipe; and

and inputting the initial inner ballistic performance parameters and the initial outer ballistic performance parameters into the nozzle ablation rate approximate model to obtain a target nozzle ablation predictive model of the spacecraft.

2. The method of predicting nozzle erosion of claim 1, wherein said operating condition parameters include an engine parameter and a flight overload parameter;

the determining the parameter variation range of the working condition parameters in the flight process of the spacecraft according to the initial inner ballistic performance parameters and the initial outer ballistic performance parameters comprises the following steps:

determining the variation range of engine parameters in the flight process of the spacecraft according to the initial inner ballistic performance parameters;

and determining the flight overload parameter variation range of the spacecraft in the flying process according to the initial outer ballistic performance parameters.

3. The method of indicating nozzle erosion of an engine of claim 1 wherein said simulating nozzle erosion rate over said range of parameter variation to obtain an approximate nozzle erosion rate model comprises:

performing two-phase flow simulation on the spacecraft engine within the parameter variation range to obtain the particle distribution rule of condensed-phase particles of the engine corresponding to each working condition parameter within the parameter variation range;

inputting the particle distribution rule into a nozzle throat lining material ablation model to obtain a first nozzle ablation rate corresponding to each working condition parameter; and

and constructing the approximate model of the nozzle ablation rate according to the first nozzle ablation rate corresponding to each working condition parameter.

4. The method of indicating nozzle ablation of an engine of claim 3, wherein said performing a two-phase flow simulation of said spacecraft engine over a range of said parameter comprises:

constructing a three-dimensional flow field physical model of the engine spray pipe; wherein the three-dimensional flow field physical model comprises a free volume and a flow field region of the engine nozzle;

and performing two-phase flow simulation of the spacecraft engine in the parameter variation range based on the three-dimensional flow field physical model.

5. The method of indicating nozzle ablation of an engine of claim 3, wherein said performing a two-phase flow simulation of said spacecraft engine over a range of variations of said parameter comprises:

performing two-phase flow simulation of the spacecraft engine by using a dynamic equation of condensed-phase particles as a simulation model;

the kinetic equation of the condensed-phase particles is as follows:

wherein,

is the particle velocity vector, t is time, m _p The mass of the particles is the mass of the particles,

in order to obtain the drag force,

in order to obtain a pressure difference force,

is the inertial force caused by overload.

6. The method of indicating nozzle ablation of claim 5, wherein said nozzle throat liner material ablation model comprises an Oka erosion ratio model;

the calculation formula of the Oka erosion ratio model is as follows:

wherein e is _r Is the erosion ratio of the particles to the nozzle throat lining material, which indicates the erosion ratio as a function of the angle of incidence, e _ref Represents the erosion ratio under normal incidence reference conditions, U is the particle velocity, D is the particle diameter, U _ref And D _ref Denotes the velocity and diameter, k, of the particle under the reference condition ₁ And k ₂ Is an empirical coefficient in the model;

the step of inputting the particle distribution rule into a nozzle throat lining material ablation model to obtain the first nozzle ablation rate corresponding to each working condition parameter comprises the following steps:

and inputting the diameter of condensed-phase particles and the particle distribution rule into an Oka erosion ratio model to obtain the ablation rate of the first nozzle.

7. The method of predicting nozzle erosion of an engine of claim 5, wherein said constructing said nozzle erosion rate approximation model based on said first nozzle erosion rate for each operating condition parameter comprises:

according to the first nozzle ablation rate corresponding to each working condition parameter, solving a nozzle ablation rate approximate model corresponding to the working condition parameter range by adopting a polynomial fitting method; the approximate model of the nozzle ablation rate is as follows:

wherein r is the nozzle ablation rate, and the unit is mm/s; p is a radical of _c Is the combustion chamber pressure; m is a group of _W Is the mole percent of water in the gas phase product upon combustion of the engine charge; a is _y Is a lateral overload in flight; r is a radical of hydrogen ₀ Representing no overload, combustion chamber pressure being a reference value p _av Nozzle ablation rate with propellant formulation as baseline formulation.

8. The method of predicting nozzle ablation of an engine as claimed in any one of claims 1 to 7, wherein after inputting the actual operating condition parameters of the spacecraft into the nozzle ablation rate approximation model to obtain the target nozzle ablation prediction model of the spacecraft, the method further comprises:

calculating target inner ballistic performance parameters and target outer ballistic performance parameters of the spacecraft according to the target nozzle ablation prediction model;

judging whether the target outer trajectory performance parameters are converged;

and if the target outer ballistic performance parameters are judged not to be converged, inputting the target inner ballistic performance parameters and the target outer ballistic performance parameters into the nozzle ablation rate approximate model to obtain a nozzle ablation prediction model after the iteration of the spacecraft.

9. An apparatus for indicating nozzle erosion in an engine, comprising: the simulation system comprises a calculation module, a determination module and a simulation module;

the calculation module is used for calculating an initial inner ballistic performance parameter and an initial outer ballistic performance parameter of the spacecraft according to an initial nozzle ablation rate prediction model;

the determining module is used for determining the parameter variation range of the working condition parameters in the flight process of the spacecraft according to the inner ballistic performance parameters and the outer ballistic performance parameters;

the simulation module is used for simulating the ablation rate of the jet pipe in the parameter change range to obtain an approximate model of the ablation rate of the jet pipe; and

the simulation module is further used for inputting the initial inner ballistic performance parameters and the initial outer ballistic performance parameters into the nozzle ablation rate approximate model to obtain a target nozzle ablation prediction model of the spacecraft.

10. A computer-readable storage medium, having stored thereon a computer program which, when executed by a processor, performs the method of any one of claims 1 to 8.

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