CN115730445A - Method, apparatus and storage medium for indicating engine nozzle ablation - Google Patents
Method, apparatus and storage medium for indicating engine nozzle ablation Download PDFInfo
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Abstract
The application provides a method, a device and a storage medium for indicating ablation of an engine nozzle, wherein the method comprises the following steps: calculating an initial inner ballistic performance parameter and an initial outer ballistic performance parameter of the spacecraft according to the initial nozzle ablation rate prediction model; determining a parameter variation range of 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; 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 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. According to the method, the device and the storage medium for predicting the ablation of the engine nozzle, the inner ballistic trajectory parameter and the outer ballistic parameter are calculated through a given initial nozzle ablation rate prediction model, simulation is carried out based on the inner ballistic parameter and the outer ballistic parameter, and the finally obtained target nozzle ablation prediction model is closer to the ablation rule of the nozzle in the actual flight process of the spacecraft.
Description
Technical Field
The application relates to the technical field of engines, in particular to a method and a device for indicating ablation of an engine nozzle and a storage medium.
Background
In the field of engines, especially in the field of aircraft engines, with the continuous development of aerospace science and technology and the continuous improvement of requirements of people for entering and exiting space, the design and development of advanced aerospace vehicles become the current research focus and hot spot. The solid rocket adopts a solid rocket engine as a power system, has the advantages of simple structure, high reliability, short response time, low development difficulty and the like, and is an important component in a carrier rocket system. The method has the advantages that the actual flight performance of the solid rocket engine is accurately evaluated, the design potential of the solid rocket engine is fully excavated, and the method has very important significance for improving the design level of the solid rocket.
A traditional solid rocket engine design model usually comprises components such as a drug shape design, internal ballistic performance calculation, component design and the like, and the change rule of internal ballistic parameters such as pressure intensity and thrust of an engine combustion chamber along with time is calculated by adopting a zero-dimensional internal ballistic equation, wherein key system parameters such as propellant burning speed and nozzle ablation rate give experience values according to ground test results. The design process does not take into account the influence of the out-of-rocket ballistic flight process on the in-engine ballistic operation process. Because the actual flight state of the rocket deviates from the ground test state of the engine, the performance parameters of the engine obtained by flight test calculation also deviate from the calculation parameters of the engine design model based on the ground test to a certain extent, and the deviation can influence the accurate evaluation of the engine performance and the overall flight performance of the solid rocket in the design and development stage to a certain extent.
Specifically, for predicting the ablation rule of the throat part of the jet pipe, the ablation rate of the throat part influences the change rule of the diameter of the throat part of the engine, so that parameters such as the pressure of a combustion chamber of the engine, the expansion ratio of the jet pipe, the specific impulse efficiency and the like are changed, and the prediction accuracy of the ballistic performance in the engine is influenced. Relevant research shows that the ablation rate of the nozzle throat is obviously influenced by the overload change of the outer ballistic flight in the actual flight process. However, in the prior art, the predicted result of the ablation rule of the engine nozzle has certain deviation from the actual flight test result.
Disclosure of Invention
The embodiment of the application aims to provide a method, a device and a storage medium for predicting the ablation of an engine nozzle.
In a first aspect, embodiments of the present application provide a method of indicating nozzle erosion in an engine, comprising: calculating initial inner ballistic performance parameters and initial outer ballistic performance parameters of the spacecraft according to an initial nozzle ablation rate prediction model; 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; 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 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.
According to the prediction method for the engine nozzle ablation, the change range of the working condition parameters for simulating the engine nozzle ablation is determined by calculating the initial inner ballistic performance parameters and the initial outer ballistic performance parameters, the nozzle ablation rate in the parameter change range is obtained by simulation, the initial inner ballistic performance parameters and the initial outer ballistic performance parameters are introduced, and the target nozzle ablation prediction model is obtained. The problem of among the prior art, through simple ground experiment, lack the consideration of external ballistic trajectory flight in-process influence factor, there is the deviation in obtaining the ablation law of spray tube and the ablation law of spray tube in the actual flight process is solved. Further, more valuable guidance is provided for the design of the engine.
With reference to the first aspect, optionally, wherein the operating condition parameters comprise an engine parameter and a flight overload parameter; the determining the parameter variation range of the 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 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.
According to the prediction method for the ablation of the engine nozzle, other parameters which have small influence on the ablation rule of the engine nozzle in the flight process of the spacecraft are eliminated by determining the variation range of the engine parameters and the variation range of the flight overload parameters, so that the complexity of simulating the ablation rate of the nozzle is reduced, and the calculation process of the ablation prediction rule of the nozzle is simplified.
With reference to the first aspect, optionally, wherein the performing nozzle ablation rate simulation on the parameter variation range to obtain a nozzle ablation rate approximation model includes: 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 constructing the approximate model of the nozzle ablation rate according to the first nozzle ablation rate corresponding to each working condition parameter.
According to the prediction method for the engine nozzle ablation, the particle distribution rule of condensed-phase particles is obtained by performing two-phase flow simulation on the spacecraft engine, the particle distribution rule is further input into a nozzle throat lining material ablation model, and finally a nozzle ablation rate approximation model is constructed by obtaining the first nozzle ablation rate corresponding to each working condition parameter. The finally obtained ablation prediction model of the target nozzle is more accurate. In combination with the first aspect, optionally, wherein the performing two-phase flow simulation of the spacecraft engine over the parameter variation range 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.
According to the prediction method for the ablation of the engine nozzle, the two-phase flow simulation is carried out on the spacecraft engine by constructing the three-dimensional flow field physical model, so that the simulation calculation result is more accurate, and the accuracy of the target nozzle ablation prediction model is finally improved. And the three-dimensional flow field physical model can describe the characteristics of a real flow field and is suitable for two-phase flow simulation of spacecraft engines with different scales under various overload conditions.
With reference to the first aspect, optionally, wherein the performing a two-phase flow simulation of the spacecraft engine over the parameter variation range 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,
as the result of the drag force,
in order to obtain the pressure difference force,
is the inertial force caused by overload.
The method for predicting the ablation of the engine nozzleIntroducing an inertia force parameter caused by overload by combining a dynamic primitive equation of condensed-phase particles and the influence of flight overload
The particle distribution rule of condensed phase particles of the spacecraft engine is expressed more accurately and visually, and the accuracy of the ablation prediction model of the target nozzle is finally improved.
With reference to the first aspect, optionally, wherein the 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 conditions 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.
According to the prediction method for the ablation of the engine nozzle, simulation is carried out through the Oka erosion ratio model, the distribution rule of the mechanical erosion of condensed phase particles of the spacecraft engine nozzle is expressed more accurately and intuitively, and finally the accuracy of the target nozzle ablation prediction model is further improved.
With reference to the first aspect, optionally, wherein the constructing the nozzle ablation rate approximation model according to the first nozzle ablation rate corresponding to each operating condition parameter includes: 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 percentage of water in the gas phase product upon combustion of the engine charge; a is a y Is a lateral overload in flight; r is 0 Representing no overload, combustion chamber pressure being a reference value p av Nozzle ablation rate with propellant formulation as baseline formulation.
According to the method for predicting the ablation of the engine nozzle, the first nozzle ablation rates corresponding to the working condition parameters are fused into a nozzle ablation rate approximate model by a polynomial fitting method, and the target nozzle ablation prediction model of the spacecraft can be obtained conveniently by inputting the initial inner ballistic performance parameters and the initial outer ballistic performance parameters in the follow-up process.
With reference to the first aspect, optionally, after the actual operating condition parameters of the spacecraft are input into the nozzle ablation rate approximation model to obtain a target nozzle ablation predictive model of the spacecraft, the method further includes: 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 parameter is 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 spacecraft is iterated.
According to the prediction method for the engine nozzle ablation, repeated iteration of the inner ballistic performance parameter and the outer ballistic performance parameter calculated based on the currently obtained nozzle ablation prediction model and the inner ballistic performance parameter and the outer ballistic performance parameter calculated based on the nozzle ablation prediction model is carried out until the outer ballistic performance parameter calculated by the nozzle ablation prediction model converges. The finally obtained post-iteration spray pipe ablation prediction model more accurately reflects the ablation rule of the spray pipe in the actual flight process, and further provides more valuable guidance for the design of the engine.
In a second aspect, embodiments of the present application further provide a device for indicating nozzle ablation in an engine, the engine including a spacecraft engine, the device 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 also 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.
The foregoing embodiment provides a device for indicating nozzle ablation having the same advantages as those of the foregoing first aspect, or a method for indicating nozzle ablation provided in any one of the alternative embodiments of the first aspect, and details are not described herein.
In a third aspect, the present application also provides a computer-readable storage medium, on which a computer program is stored, where the computer program is executed by a processor to perform the above-described method.
The foregoing embodiment provides a computer-readable storage medium having the same advantages as the foregoing first aspect, or any alternative embodiment of the first aspect, and will not be described herein again.
In a third aspect, an embodiment of the present application further provides an electronic device, including: a processor and a memory, the memory storing processor-executable machine-readable instructions which, when executed by the processor, perform a method as described above.
The foregoing embodiment provides an electronic device having the same advantages as the foregoing first aspect, or the method for predicting engine nozzle ablation provided in any one of the alternative embodiments of the first aspect, and details are not repeated here.
In summary, according to the method, the device and the storage medium for predicting the ablation of the engine nozzle, the inner ballistic performance parameter and the initial outer ballistic performance parameter of the spacecraft are calculated through the initial nozzle ablation rate prediction model obtained based on the prior art, and then simulation is performed based on the initial inner ballistic performance parameter and the initial outer ballistic performance parameter, so that the obtained target nozzle ablation prediction model is closer to the ablation rule of the nozzle in the actual flight process of the spacecraft. Meanwhile, a two-phase flow simulation method is adopted during simulation, the particle distribution rule of condensed-phase particles of the spacecraft engine can be more accurately and visually reflected, and the accuracy of the ablation prediction model of the target nozzle is improved. And finally, optimizing the target nozzle ablation prediction model through repeated iteration, and reflecting the ablation rule of the nozzle in the actual flight process more accurately by the finally obtained post-iteration nozzle ablation prediction model, thereby further providing more valuable guidance for the design of the engine.
Drawings
To more clearly illustrate the technical solutions of the embodiments of the present application, the drawings that are required to be used in the embodiments of the present application will be briefly described below, it should be understood that the following drawings only illustrate some embodiments of the present application and therefore should not be considered as limiting the scope, and that those skilled in the art can also obtain other related drawings based on the drawings without inventive efforts.
FIG. 1 is a first flowchart of a method for indicating nozzle erosion in an engine provided in accordance with an embodiment of the present disclosure;
FIG. 2 is a detailed flowchart of step S120 of a method for indicating nozzle ablation in an engine provided in accordance with an embodiment of the present application;
FIG. 3 is a detailed flowchart of step S130 of a method for indicating nozzle ablation according to an embodiment of the present application;
FIG. 4 is a detailed flowchart of step S131 in a method for predicting nozzle erosion according to an embodiment of the present application;
FIG. 5 is a second flowchart of a method for indicating nozzle erosion according to an embodiment of the present application;
FIG. 6 is a functional block diagram of a predictive device for nozzle ablation in an engine provided in accordance with an embodiment of the present application;
fig. 7 is a schematic structural diagram of an electronic device according to an embodiment of the present application;
FIG. 8 is a diagram illustrating a condensed-phase particle distribution simulated when the lateral overload is 0 according to an embodiment of the present application;
FIG. 9 is a diagram of a condensed-phase particle distribution simulated at a transverse overload of 5g according to an embodiment of the present application;
FIG. 10 is a diagram illustrating a condensed-phase particle distribution simulated at a lateral overload of 10g according to an embodiment of the present application;
FIG. 11 is a graph illustrating a simulated nozzle mechanical erosion rate distribution when a lateral overload is 0 according to an embodiment of the present application;
FIG. 12 is a graph illustrating a simulated nozzle mechanical erosion rate distribution when a lateral overload of 5g is applied according to an embodiment of the present disclosure;
FIG. 13 is a graph illustrating the mechanical erosion rate of the nozzle simulated at a transverse overload of 10g according to an embodiment of the present application.
Detailed Description
Embodiments of the present invention will be described in detail below with reference to the accompanying drawings. The following examples are merely used to more clearly illustrate the technical solutions of the present application, and therefore are only examples, and the protection scope of the present application is not limited thereby.
Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this application belongs; the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the application.
In the description of the embodiments of the present application, the technical terms "first", "second", and the like are used only for distinguishing different objects, and are not to be construed as indicating or implying relative importance or to implicitly indicate the number, specific order, or primary-secondary relationship of the technical features indicated. In the description of the embodiments of the present application, "a plurality" means two or more unless specifically defined otherwise.
Based on the description of the prior art in the background section, the applicant finds that by considering the internal and external ballistic coupling mechanism, a solid rocket engine nozzle ablation prediction model with good celestial and terrestrial consistency is established, and the internal ballistic performance of the engine can be calculated more accurately in the design and development stage of the solid rocket.
The main action mechanism that the ablation rate of the throat part of the spray pipe is influenced by external flight overload is that the external ballistic flight overload acts on a two-phase flow field in the solid rocket engine to cause the movement locus of condensed phase particles in the engine to deviate, the condensed phase particles are gathered in a combustion chamber and the throat part of the spray pipe, and the concentration of the condensed phase particles is increased. Namely, a two-phase flow simulation method of a solid rocket engine under the condition of flight overload is provided. The change of the motion law of the condensed phase components increases the mechanical erosion to the surface of the throat lining material, so that the ablation rate of the throat lining of the jet pipe is increased, and the phenomena of burning deviation along the overload direction occur.
In the prior art, although a solid rocket engine design model based on ground tests is established, wherein the solid rocket engine design model comprises an empirical model for predicting nozzle ablation rate, and theoretical analysis and experimental research are carried out on the nozzle ablation rule of the solid rocket engine under the influence of external ballistic flight overload, the solid rocket engine design model still has defects and is mainly embodied in the following two aspects:
on one hand, the nozzle ablation rate in the existing solid rocket engine design model is generally given based on engineering experience of ground tests, for example, the average linear ablation rate is obtained by calculating the throat diameter change before and after the tests, the influence factors (mainly flight overload) of the external ballistic flight process are not considered, and the predicted result of the engine nozzle ablation rule has certain deviation from the actual flight test result;
on the other hand, the conventional research only generally carries out theoretical analysis and experimental exploration on a mechanism level on a nozzle ablation rule under the overload influence, a relevant influence rule formed by simulation and experiment is not closely combined with an engine design and development process, and a nozzle ablation prediction model considering the coupling influence of internal and external ballistic trajectories can be used for guiding the engine design process.
In view of the above two drawbacks, the present application provides a method, an apparatus and a storage medium for indicating nozzle ablation in an engine, so as to solve the above drawbacks in the prior art. Specifically, please refer to the embodiments and drawings provided in the present application.
Referring to FIG. 1, FIG. 1 is a first flowchart illustrating a method for predicting nozzle erosion in an engine according to an embodiment of the present disclosure. The method for indicating the ablation of the engine nozzle provided by the embodiment of the application comprises the following steps:
step S110: and calculating initial inner ballistic performance parameters and initial outer ballistic performance parameters of the spacecraft according to the initial nozzle ablation rate prediction model.
In step S110, the initial nozzle ablation prediction model may be given based on engineering experience of the ground experiment of the engine. And calculating the internal ballistic performance of the engine according to the initial nozzle ablation prediction model to obtain the initial internal ballistic performance parameters of the spacecraft. 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 may be calculated from the initial inner ballistic performance parameter. Specifically, the initial inner ballistic performance parameters are input into a solid rocket outer ballistic design model in the prior art, and outer ballistic performance of the spacecraft is calculated to obtain initial outer ballistic performance parameters. Initial outer ballistic performance parameters include, but are not limited to: flight speed, flight overload parameters, shutdown point height parameters, and ballistic dip parameters.
Step S120: and 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.
In step S120, the variation range of the operating condition parameter includes, but is not limited to: the variation range of the engine thrust of the spacecraft in the flying process, the variation range of the spacecraft mass parameters, the variation range of the flying speed, the variation range of the flying overload parameters and the variation range of the ballistic inclination parameters.
Step S130: and (5) 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.
In step S130, the ablation rate of the nozzle refers to the ablation rule of the engine nozzle of the spacecraft in the flight process. According to the variation range of the working condition parameters determined in the step S120, the nozzle ablation rate of the engine of the spacecraft in the flying process can be simulated, and an approximate model of the nozzle ablation rate is obtained.
Step S140: 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 prediction model of the spacecraft.
In the step S140, the initial inner ballistic performance parameter and the initial outer ballistic performance parameter are input into the nozzle ablation rate approximation model as input parameters, and the output of the model is the nozzle ablation prediction law of the spacecraft, that is: and (4) a target nozzle ablation prediction model.
In the implementation process, the variation range of the working condition parameters for simulating the flight process of the spacecraft is determined by calculating the initial inner ballistic performance parameters and the initial outer ballistic performance parameters, the nozzle ablation rate in the parameter variation range is obtained by simulation, the initial inner ballistic performance parameters and the initial outer ballistic performance parameters are introduced, and the target nozzle ablation prediction model is obtained. The problem of among the prior art, through simple ground experiment, lack the consideration of external ballistic trajectory flight in-process influence factor, there is the deviation in obtaining the ablation law of spray tube and the ablation law of spray tube in the actual flight process is solved. Further, more valuable guidance is provided for the design of the engine.
Referring to FIG. 2, FIG. 2 is a detailed flowchart of step S120 of a method for indicating nozzle erosion in an engine according to an embodiment of the present disclosure, wherein in an alternative embodiment, the operating parameters include engine parameters and flight overload parameters.
The step S120 includes:
step S121: and determining the variation range of the engine parameters in the flight process of the spacecraft according to the initial internal ballistic performance parameters.
Step S122: and determining the flight overload parameter variation range of the spacecraft in the flight process according to the initial outer ballistic performance parameters.
In the above steps, the engine parameter in the initial inner ballistic performance parameter and the flight overload parameter in the initial outer ballistic performance parameter are the most important parameters influencing the ablation rule of the engine nozzle in the flight process of the spacecraft.
In the implementation process, other parameters which have small influence on the ablation rule of the engine nozzle in the flight process of the spacecraft are eliminated by determining the variation range of the engine parameters and the variation range of the flight overload parameters, so that the complexity of simulating the ablation rate of the nozzle is reduced, and the calculation process of the ablation prediction rule of the nozzle is simplified.
Referring to fig. 3, fig. 3 is a detailed flowchart of step S130 of a method for predicting nozzle erosion in an engine according to an embodiment of the present application, where in an alternative embodiment, the step S130 includes:
step S131: and (3) performing two-phase flow simulation on the spacecraft engine within the parameter variation range to obtain the particle distribution rule of the condensed-phase particles of the engine corresponding to each working condition parameter within the parameter variation range.
In step S131, related studies in the prior art in the field indicate that the longitudinal overload has little influence on the ablation rate of the nozzle throat, the lateral overload has a more significant influence on the ablation rate of the nozzle throat, and the ablation rate changes more significantly as the lateral overload increases. Therefore, in the embodiment of the application, in order to clearly show the change of the nozzle throat ablation rule when the coupling of the inner and outer trajectories is considered and not considered, the change range of the transverse flight overload of the outer trajectory can be considered and only given as the working condition range for carrying out numerical simulation calculation of the flow field in the solid rocket engine. And the particle distribution rule of the condensed-phase particles of the engine corresponding to each working condition parameter can be obtained through simulation.
Step S132: and 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.
In step S132, the nozzle throat lining material ablation model includes, but is not limited to, an Oka erosion ratio model, a Neilson model, and a Gilchrist model, and the particle distribution rule corresponding to each parameter is input into the nozzle throat lining material ablation model, so that the model outputs the first nozzle ablation rate corresponding to each operating condition parameter.
Step S133: and constructing a nozzle ablation rate approximate model according to the first nozzle ablation rate corresponding to each working condition parameter.
In step S133, since the first nozzle ablation rates corresponding to the operating condition parameters are obtained, the nozzle ablation rates of the spacecraft in the states corresponding to the different operating condition parameters during the flight process can be obtained accordingly. Therefore, a nozzle ablation rate approximate model is constructed by the first nozzle ablation rates corresponding to the working condition parameters, and a target nozzle ablation prediction model of the spacecraft in the corresponding flight state can be finally obtained by inputting the corresponding initial inner ballistic performance parameters and initial outer ballistic performance parameters.
In the implementation process, a particle distribution rule of condensed-phase particles is obtained by performing two-phase flow simulation on the spacecraft engine, the particle distribution rule is further input into a nozzle throat lining material ablation model, and finally a nozzle ablation rate approximation model is constructed by obtaining a first nozzle ablation rate corresponding to each working condition parameter. The finally obtained ablation prediction model of the target nozzle is more accurate.
Referring to FIG. 4, FIG. 4 is a flowchart illustrating step S131 of a method for predicting nozzle erosion in an engine according to an embodiment of the present disclosure. In an alternative embodiment, the step S131 includes:
step S1311: and constructing a three-dimensional flow field physical model of the engine spray pipe. The three-dimensional flow field physical model comprises a free volume and a flow field area of the engine nozzle.
In step S1311, the three-dimensional flow field physical model includes a free volume in the combustion chamber and a flow field region in the nozzle. In practical application, the three-dimensional flow field physical model can be properly simplified according to conditions, the calculation efficiency is improved on the premise of not influencing the calculation result, for example, when the three-dimensional flow field physical characteristic is symmetrical in a geometric plane and is overloaded without a lateral component, half of the three-dimensional flow field physical model can be adopted; if the nozzle expansion is relatively large, it may be cut back appropriately.
Step S1312: and performing two-phase flow simulation of the spacecraft engine in a parameter change range based on the three-dimensional flow field physical model.
In the step S1312, a hexahedral structured grid discrete simulation region may be adopted, the boundary conditions of the simulation region are determined according to the working condition parameters of the spacecraft engine during actual operation, the combustion products of the propellant are regarded as two parts, namely, a gas phase part and a condensed phase particle part, and the simulated boundary conditions are set for the two parts respectively. The two-phase flow field is calculated and solved by adopting an Euler-Lagrange method, namely, the gas phase is regarded as a continuous phase, a control equation of the gas phase is expressed in an Euler form, and condensed phase particles are regarded as a discrete phase and are solved under a Lagrange coordinate. There is a transfer of momentum and energy between the gas phase and condensed phase particles, affecting the gas phase flow and particle trajectories. And (3) carrying out two-phase flow field simulation calculation on the target solid rocket engine by adopting a certain condensed phase particle size distribution rule and certain gas phase parameters, and obtaining the motion and distribution rule of the particles in the flow field.
In the implementation process, the two-phase flow simulation is carried out on the spacecraft engine by constructing the three-dimensional flow field physical model, so that the simulation calculation result is more accurate, and the accuracy of the ablation prediction model of the target nozzle is finally improved. And the three-dimensional flow field physical model can describe the characteristics of a real flow field and is suitable for two-phase flow simulation of spacecraft engines with different scales under various overload conditions.
In an optional implementation manner, the S131 further includes:
step S1313: and (3) 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 for the condensed phase particles is:
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 an inertial force caused by overload.
In the step S1313, the change range of the transverse flight overload may be set to 0 to 10g, and the numerical simulation of the two-phase flow of the solid rocket engine considering the overload influence is performed. And (3) simulating to obtain the distribution rule of condensed Phase particles in the engine in the flow field based on a high-precision drag model and a DPM (Discrete Phase Method) model established by particle motion numerical simulation.
In the simulation process, the condensed-phase particles are assumed to be spherical, the splitting and merging among the particles are neglected, and the actual continuous particle size distribution is replaced by a plurality of discrete values with different diameters so as to simplify the calculation. The DPM model is adopted, and a drag coefficient model is input in the form of UDF (user defined function) and is used for calculating the drag force of the particles in the supersonic flow of the spray pipe. Using a collision model on the boundary, the particles bounce with a certain coefficient of restitution (set to 0.8 in this example) after colliding with the nozzle wall. The original kinetic equation for 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,
for the particle displacement vector, if affected by flight overload, the kinetic equation for condensed phase particles becomes:
wherein,
is an inertial force caused by overload.
Referring to fig. 8 to 10, fig. 8 is a diagram illustrating a condensed-phase particle distribution simulated when the lateral overload is 0 according to an embodiment of the present application; FIG. 9 is a diagram illustrating a condensed-phase particle distribution simulated at a lateral overload of 5g provided by an embodiment of the present application; fig. 10 is a diagram of a condensed-phase particle distribution simulated at a lateral overload of 10g provided by an embodiment of the present application. The simulation result shows that when no transverse overload exists, the particles are gathered at the central position of the engine and are in a symmetrical and long shape, and the maximum concentration value reaches 11.2kg/m 3 And this high concentration "narrow band" does not hit the nozzle wall. After addition of 5g of transverse overload, the maximum particle concentration had dropped to 3.22kg/m 3 Significant divergence and upward deflection occurs and impinges at a location between the convergent section and the throat of the nozzle. After the transverse overload reaches 10g, the particles are distributed and deflected upwards obviously, gather on the wall surface flowing in the opposite direction of the external overload and rise to the maximum concentration of 7.33kg/m 3 And the erosion on the wall of the nozzle is wider but less concentrated than with a 5g lateral overload.
In the implementation process, the inertia force parameter caused by overload is introduced by combining the original kinetic equation of condensed-phase particles and the influence of flight overload
The particle distribution rule of condensed phase particles of the spacecraft engine is expressed more accurately and visually, and the accuracy of the ablation prediction model of the target nozzle is finally improved.
In an alternative embodiment, the nozzle throat liner material ablation model comprises an Oka erosion ratio model calculated by the equation:
wherein e is r Is the erosion ratio of the particles to the nozzle throat liner material, which indicates the variation of the erosion ratio with incident angle, 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 Are empirical coefficients in the model.
Accordingly, the step S132 includes:
step S1321: 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.
Referring to fig. 11 to 13, fig. 11 is a graph illustrating a mechanical erosion rate distribution of the nozzle simulated when the lateral overload is 0 according to the embodiment of the present application; FIG. 12 is a graph illustrating a simulated nozzle mechanical erosion rate distribution at a lateral overload of 5g provided by an embodiment of the present application; FIG. 13 is a graph illustrating the mechanical erosion rate of the nozzle simulated at a transverse overload of 10g according to an embodiment of the present application. In step S1321, it can be seen from fig. 11 to 13 that the convergent section upstream and the divergent section of the nozzle are not mechanically corroded. Mechanical erosion is mainly concentrated in the convergent section and near the throat. The maximum value of mechanical denudation reaches 0.1722kg when the transverse overload is 5g/(m 2 S). The local ablation rate is obviously increased along the external overload reverse direction. The ablation rate in the remaining direction positions is concentrated to about 0.04-0.07 kg/(m) 2 S). The maximum value of mechanical erosion reaches 0.2378 kg/(m) when the transverse overload is 10g 2 S) along the direction opposite to the external overload, the local erosion rate is increased, the erosion area of the nozzle convergent section is increased, and the nozzle convergent section extends in a linear shape towards the combustion chamber. The erosion rate is also concentrated at about 0.04-0.07 kg/(m) at the rest of the nozzle direction 2 S).
In the implementation process, the Oka erosion ratio model is used for simulation, the distribution rule of the jet pipe of the spacecraft engine, which is mechanically eroded by condensed phase particles, is expressed more accurately and intuitively, and the accuracy of the target jet pipe ablation prediction model is further improved finally.
In an alternative embodiment, the step S133 includes:
step S1331: and solving an approximate model of the ablation rate of the spray pipe corresponding to the range of the working condition parameters by adopting a polynomial fitting method according to the ablation rate of the first spray pipe corresponding to each working condition parameter. 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 a y As a lateral overload in flight; r is 0 Representing no overload, combustion-chamber pressure being a reference value p av And nozzle ablation rate with the propellant formulation as the baseline formulation.
In the implementation process, the first nozzle ablation rates corresponding to the working condition parameters are fused into a nozzle ablation rate approximate model by a polynomial fitting method, so that the target nozzle ablation prediction model of the spacecraft can be obtained conveniently by inputting the initial inner ballistic performance parameters and the initial outer ballistic performance parameters.
Referring to FIG. 5, FIG. 5 is a second flowchart of a method for predicting nozzle erosion in an engine according to an embodiment of the present disclosure. In an alternative embodiment, after step S140, the method for predicting erosion of an engine nozzle provided by the embodiment of the present application further includes:
step S150: and calculating the target inner ballistic performance parameters and the target outer ballistic performance parameters of the spacecraft according to the target nozzle ablation prediction model.
Step S160: and judging whether the target outer trajectory performance parameters are converged.
If it is determined that the target outer trajectory performance parameter does not converge, step S170 is performed: and inputting the target inner trajectory performance parameters and the target outer trajectory performance parameters into the nozzle ablation rate approximate model to obtain a nozzle ablation prediction model after the spacecraft is iterated.
In the steps, the target nozzle ablation prediction model is adopted to calculate target inner trajectory performance parameters and target outer trajectory performance parameters, whether the target outer trajectory performance parameters are converged is judged, and if the target inner trajectory performance parameters and the target outer trajectory performance parameters are not converged, the nozzle ablation prediction model after the spacecraft iteration is calculated on the basis of the target inner trajectory performance parameters and the target outer trajectory performance parameters. It should be appreciated that the post-iteration nozzle ablation prediction model may be verified again using substantially the same method as described above, through such an iterative process, until the outer ballistic performance parameters converge.
In the implementation process, the inner ballistic performance parameter and the outer elastic performance parameter calculated based on the current nozzle ablation prediction model are repeatedly iterated until the outer elastic performance parameter calculated by the nozzle ablation prediction model converges. The finally obtained ablation prediction model of the nozzle after iteration reflects the ablation rule of the nozzle in the actual flight process more accurately, and further provides more valuable guidance for the design of the engine.
Referring to FIG. 6, FIG. 6 is a functional block diagram of a predictive device 600 for engine nozzle ablation according to an embodiment of the present disclosure. The embodiment of the application provides a device 600 for indicating engine nozzle ablation, comprising: a calculation module 610, a determination module 620, and a simulation module 630.
The calculation module 610 is used for calculating an initial inner ballistic performance parameter and an initial outer ballistic performance parameter of the spacecraft according to the initial nozzle ablation rate prediction model. The determining module 620 is configured to determine a parameter variation range of a working condition parameter of the spacecraft in the flight process according to the inner ballistic performance parameter and the outer ballistic performance parameter. The simulation module 630 is configured to perform nozzle ablation rate simulation within the parameter variation range to obtain a nozzle ablation rate approximation model. The simulation module 630 is further configured to input the initial inner ballistic performance parameter and the initial outer ballistic performance parameter into the nozzle ablation rate approximation model to obtain a target nozzle ablation prediction model of the spacecraft.
With continued reference to FIG. 6, in an alternative embodiment, the operating parameters include engine parameters and flight overload parameters.
Correspondingly, the determining module 620 is specifically configured to:
determining the variation range of the 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 flight process according to the initial outer ballistic performance parameters.
Referring to fig. 6, in an alternative embodiment, the simulation module 630 is specifically configured to:
two-phase flow simulation of the spacecraft engine is carried out in the parameter variation range, and the particle distribution rule of condensed-phase particles of the engine corresponding to each working condition parameter in the parameter variation range is obtained; 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 constructing a nozzle ablation rate approximate model according to the first nozzle ablation rate corresponding to each working condition parameter.
Referring to fig. 6, in an alternative embodiment, the simulation module 630 is further configured to:
constructing a three-dimensional flow field physical model of the engine spray pipe; the three-dimensional flow field physical model comprises a free volume and a flow field area of the engine nozzle; and performing two-phase flow simulation of the spacecraft engine in a parameter change range based on the three-dimensional flow field physical model.
Referring to fig. 6, in an alternative embodiment, the simulation module 630 is further specifically configured to:
two-phase flow simulation of the spacecraft engine is carried out by taking a kinetic equation of condensed-phase particles as a simulation model; the kinetic equation for condensed phase particles is:
wherein,
is the particle velocity vector, t is time, m p The mass of the particles is the mass of the particles,
as the result of the drag force,
in order to obtain a pressure difference force,
is an inertial force caused by overload.
With continued reference to fig. 6, in an alternative embodiment, the nozzle throat liner material ablation model includes an Oka erosion rate model; the calculation formula of the Oka erosion ratio model is:
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 Denotes the erosion ratio at 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 conditions 1 And k 2 Are empirical coefficients in the model.
Correspondingly, the simulation module 630 is further configured to input the diameter of the condensed-phase particles and the particle distribution rule into the Oka erosion ratio model to obtain the first nozzle ablation rate corresponding to each operating condition parameter.
Referring to fig. 6, in an alternative embodiment, the simulation module 630 is further specifically configured to:
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 formula c Is the combustion chamber pressure; m W Is the mole percent of water in the gas phase product upon combustion of the engine charge; a is a y Is a lateral overload in flight; r is 0 Representing no overload, combustion chamber pressure being a reference value p av And nozzle ablation rate with the propellant formulation as the baseline formulation.
Referring to fig. 6, in an alternative embodiment, the calculating module 610 is further configured to calculate an in-target ballistic performance parameter and an out-target ballistic performance parameter of the spacecraft according to the target nozzle ablation prediction model; and judging whether the target outer trajectory performance parameter is converged.
If the target outer ballistic performance parameter is determined not to be converged, correspondingly, the simulation module 630 is further configured to input the target inner ballistic performance parameter and the target outer ballistic performance parameter into the nozzle ablation rate approximation model to obtain the nozzle ablation prediction model after the iteration of the spacecraft.
It should be understood that the apparatus corresponds to the above-described prophetic method embodiment of nozzle ablation, and is capable of performing the steps involved in the above-described method embodiment, and the specific functions of the apparatus may be referred to the above description, and the detailed description is omitted here where appropriate to avoid repetition. The device includes at least one software functional module that can be stored in memory in the form of software or firmware (firmware) or solidified in the Operating System (OS) of the device.
Based on the same inventive concept, please refer to fig. 7, fig. 7 is a schematic structural diagram of an electronic device 700 provided in the embodiment of the present application. The electronic device 700 may include a memory 711, a memory controller 712, a processor 713, a peripheral interface 714, an input output unit 717, a display unit 716. It will be understood by those skilled in the art that the structure shown in fig. 7 is merely an illustration and is not intended to limit the structure of the electronic device 700. For example, electronic device 700 may also include more or fewer components than shown in FIG. 7, or have a different configuration than shown in FIG. 7.
The memory 711, the memory controller 712, the processor 713, the peripheral interface 714, the input/output unit 717, and the display unit 716 are electrically connected to each other directly or indirectly, so as to realize data transmission or interaction. For example, the components may be electrically connected to each other via one or more communication buses or signal lines. The processor 713 is used to execute executable modules stored in the memory.
The Memory 711 may be, but is not limited to, a Random Access Memory (RAM), a Read Only Memory (ROM), a Programmable Read-Only Memory (PROM), an Erasable Read-Only Memory (EPROM), an electrically Erasable Read-Only Memory (EEPROM), and the like. The memory 711 is configured to store a program, and the processor 713 executes the program after receiving an execution instruction, and the method performed by the electronic device 700 defined by the process disclosed in any embodiment of the present application may be applied to the processor 713, or implemented by the processor 713.
The processor 713 may be an integrated circuit chip having signal processing capabilities. The Processor 713 may be a general-purpose Processor, and include a Central Processing Unit (CPU), a Network Processor (NP), and the like; the Integrated Circuit may also be a Digital Signal Processor (DSP), an Application Specific Integrated Circuit (ASIC), a Field Programmable Gate Array (FPGA) or other programmable logic device, a discrete gate or transistor logic device, or a discrete hardware component. The various methods, steps, and logic blocks disclosed in the embodiments of the present application may be implemented or performed. A general purpose processor may be a microprocessor or the processor may be any conventional processor or the like.
The peripheral interface 714 couples various input/output devices to the processor 713 and memory 711. In some embodiments, peripheral interface 714, processor 713, and memory controller 712 may be implemented in a single chip. In other examples, they may be implemented separately from the individual chips.
The input/output unit 717 is used to provide input data to the user. The input and output unit 717 may be, but is not limited to, a mouse, a keyboard, and the like.
The display unit 716 provides an interactive interface (e.g., a user operation interface) between the electronic device 700 and a user or is used to display image data for reference by the user. In this embodiment, the display unit may be a liquid crystal display or a touch display. In the case of a touch display, the display can be a capacitive touch screen or a resistive touch screen, which supports single-point and multi-point touch operations. Supporting single-point and multi-point touch operations means that the touch display can sense touch operations simultaneously generated from one or more positions on the touch display, and the sensed touch operations are sent to the processor for calculation and processing.
The electronic device 700 in this embodiment may be configured to perform each step in each method provided in this embodiment.
The embodiment of the application also provides a storage medium, wherein the storage medium is stored with a computer program, and the computer program is executed by a processor to execute the method.
The storage medium may be implemented by any type of volatile or nonvolatile storage device or combination thereof, such as a Static Random Access Memory (SRAM), an Electrically Erasable Programmable Read-Only Memory (EEPROM), an Erasable Programmable Read-Only Memory (EPROM), a Programmable Read-Only Memory (PROM), a Read-Only Memory (ROM), a magnetic Memory, a flash Memory, a magnetic disk, or an optical disk.
In summary, according to the method, the device and the storage medium for predicting the ablation of the engine nozzle, the inner ballistic performance parameter and the initial outer ballistic performance parameter of the spacecraft are calculated through the initial nozzle ablation rate prediction model obtained based on the prior art, and the simulation is performed based on the initial inner ballistic performance parameter and the initial outer ballistic performance parameter, so that the obtained target nozzle ablation prediction model is closer to the ablation rule of the nozzle in the actual flight process of the spacecraft. Meanwhile, a two-phase flow simulation method is adopted during simulation, the particle distribution rule of condensed-phase particles of the spacecraft engine can be more accurately and visually reflected, and the accuracy of the ablation prediction model of the target nozzle is improved. And finally, optimizing the target nozzle ablation prediction model through repeated iteration, and reflecting the ablation rule of the nozzle in the actual flight process more accurately by the finally obtained post-iteration nozzle ablation prediction model, thereby further providing more valuable guidance for the design of the engine.
In the embodiments provided in the present application, it should be understood that the disclosed apparatus and method may be implemented in other ways. The apparatus embodiments described above are merely illustrative, and for example, the flowchart and block diagrams in the figures illustrate the architecture, functionality, and operation of possible implementations of apparatus, methods and computer program products according to various embodiments of the present application. In this regard, each block in the flowchart or block diagrams may represent a module, segment, or portion of code, which comprises one or more executable instructions for implementing the specified logical function(s). 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 shown in succession may, in fact, be executed substantially concurrently, or the blocks may sometimes be executed in the reverse order, depending upon the functionality involved. It will also be noted that each block of the block diagrams and/or flowchart illustration, and combinations of blocks in the block diagrams and/or flowchart illustration, can be implemented by special purpose hardware-based systems which perform the specified functions or acts, or combinations of special purpose hardware and computer instructions.
In addition, functional modules in the embodiments of the present application may be integrated together to form an independent part, or each module may exist alone, or two or more modules may be integrated to form an independent part.
The above description is only an alternative embodiment of the embodiments of the present application, but the scope of the embodiments of the present application is not limited thereto, and any person skilled in the art can easily conceive of changes or substitutions within the technical scope of the embodiments of the present application, and all the changes or substitutions should be covered by the scope 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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