CN115034157A - Method for predicting working characteristics of thixotropic propellant rocket engine - Google Patents
Method for predicting working characteristics of thixotropic propellant rocket engine Download PDFInfo
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- GDDNTTHUKVNJRA-UHFFFAOYSA-N 3-bromo-3,3-difluoroprop-1-ene Chemical compound FC(F)(Br)C=C GDDNTTHUKVNJRA-UHFFFAOYSA-N 0.000 claims description 3
- 229920002121 Hydroxyl-terminated polybutadiene Polymers 0.000 claims description 3
- 239000000654 additive Substances 0.000 claims description 3
- 239000002923 metal particle Substances 0.000 claims description 3
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- KAKZBPTYRLMSJV-UHFFFAOYSA-N Butadiene Chemical compound C=CC=C KAKZBPTYRLMSJV-UHFFFAOYSA-N 0.000 description 2
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Abstract
The invention provides a prediction method for the operating characteristics of a thixotropic propellant rocket engine, which is characterized by acquiring three-dimensional characteristics of a flow field in the engine based on various design parameters and working condition parameters of the rocket engine and performing geometric dispersion on a calculation domain through the characteristics. And then establishing a gas phase control equation, a chemical reaction kinetic system, boundary conditions and the like for the calculation domain. And finally, obtaining various indexes of the three-dimensional combustion flow of the flow field in the engine through steady-state iterative calculation. The prediction method provided by the invention can realize accurate prediction of the three-dimensional characteristics of the thixotropic propellant rocket engine, overcomes the defects that the existing prediction method has single index and cannot accurately simulate three-dimensional combustion flow behaviors, has the advantages that the calculation simulation is completely matched with an actual physical model, the calculation speed is high, the calculation precision is higher, the development time is effectively shortened, and the development cost is reduced.
Description
Technical Field
The invention relates to the field of combustion flow multi-physical field calculation, in particular to a prediction method for the working characteristics of a thixotropic propellant rocket engine.
Background
The thixotropic propellant rocket engine is considered to be a promising advanced chemical propulsion technology, can realize wide thrust regulation, is easy to control, and has high safety. A propellant in the form of a paste is a non-newtonian fluid that behaves as a solid at rest and is broken by sufficiently high shear stress that the propellant will flow like a liquid. During flow, thixotropic propellants perform very similarly to conventional liquid propellants at high shear rates. Therefore, the thixotropic propellant rocket engine organically combines the advantages of liquid and solid propelling systems, and has the advantages of high specific impact, simple structure, safety, reliability, capability of pulse working, safety in storage and transportation, wide performance adjustment range and the like.
During the operation of a thixotropic rocket engine, the pyrolysis of the propellant, the flow of the combustion gases and the combustion behavior exhibit completely different characteristics from those of a conventional rocket engine. In terms of design optimization of rocket engines, obtaining sufficient data through multiple trials consumes a great deal of time and economic cost. With the rise of computer software and hardware technologies, researchers calculate and obtain a plurality of performance parameters through simplified numerical simulation based on reasonable assumptions, so that a large amount of test work is avoided, and a basis is provided for design iteration. Therefore, constructing an accurate numerical simulation model is the key to the optimal design.
At present, the research on the internal flow field characteristics of a thixotropic propellant rocket engine is few, the combustion and flow characteristics of the internal flow field are not clear, and the internal flow field is mainly researched in two ways in the prior art. One is to use a zero-dimensional internal ballistic computational model. The method provides a conical combustion model based on combustion speed and supply flow for the unique flow and combustion mechanism of the thixotropic rocket engine, and has the problems that the characterization parameters are single, the complex heat and mass transfer mechanism under the condition of three-dimensional multi-physical fields in a combustion chamber of the rocket engine cannot be revealed, the factors such as turbulent flow, combustion, fluid-solid heat transfer and the like cannot be considered, and meanwhile, the accurate prediction on the flow and combustion behaviors of a grain pyrolysis and a flow field in a combustion chamber-spray pipe cannot be carried out. The other method is a numerical simulation method combined with thermodynamic calculation, omits the heat transfer process of the pyrolysis of the explosive column and the violent reaction process of pyrolysis products, directly solves the components of the combusted fuel gas through the thermodynamic calculation, adds the components into a fluid domain in the form of a flow inlet, and calculates the flow and heat transfer properties of the components. Although the method can obtain the three-dimensional distribution of the working characteristics, the method neglects the processes of pyrolysis, chemical reaction and the like of the propellant, so that the solving precision of the surface temperature of the grain is poor, and the combustion process and the distribution condition of the propellant components cannot be obtained, which is unfavorable for the prediction of the overall performance of the engine.
Disclosure of Invention
Aiming at the requirements in the prior art, the invention provides a prediction method of the operating characteristics of a thixotropic propellant rocket engine. And then establishing a gas phase control equation, a chemical reaction kinetic system, boundary conditions and the like for the calculation domain. And finally, obtaining various indexes of the three-dimensional combustion flow of the flow field in the engine through steady iterative calculation.
The specific technical scheme of the invention is as follows:
a method for predicting the operating characteristics of a thixotropic propellant rocket engine comprises the following steps:
s1: acquiring geometric parameters and working condition parameters of a thixotropic propellant rocket engine and formula parameters of the thixotropic propellant;
s2: carrying out burning rate measurement on the thixotropic propellant to obtain a burning rate coefficient and a pressure index;
s3: under the steady-state work, calculating the trajectory in the internal flow field of the engine combustion chamber to obtain the three-dimensional geometric characteristics of the internal flow field;
s4: simplifying the internal flow according to the three-dimensional geometrical characteristics of the internal flow field, determining a calculation domain, performing spatial dispersion, and deriving a calculation grid;
s5: importing the computational grid into CFD software, and defining the attributes of the computational domain;
s6: describing the flow state of the fluid by adopting a three-dimensional multi-component N-S control equation, wherein a turbulent flow is described by adopting an SST k-omega model; adding pyrolysis gas of the propellant grain into a calculation domain as a source item in an N-S control equation;
s7: defining a boundary condition;
s8: defining a chemical reaction kinetic system in the combustion chamber by adopting an Eddy Dissipation Concept (EDC) model;
s9: performing space dispersion on the N-S control equation by adopting a second-order windward format, and performing coupling solution of pressure and density by adopting a coupling algorithm;
s10: starting three-dimensional combustion flow steady-state simulation calculation, and monitoring residual errors of all parameters until calculation results are converged;
s11: and extracting the working characteristic parameters in the thixotropic propellant engine from the calculation result, and performing statistical analysis to finish prediction.
Preferably, the geometric parameters in step S1 include the geometric size of the combustion chamber, the diameter of the nozzle throat, the ratio of the outlet area of the delivery pipe to the area of the nozzle throat, and the diameter of the delivery pipe.
Preferably, the formulation parameters of the thixotropic propellant are as follows: 70% of ammonium perchlorate, 10% of hydroxyl-terminated polybutadiene, 5% of metal particles and 15% of other additives.
Preferably, the step S3 specifically includes:
s301: determining the pressure p in the combustion chamber eq ：
Where ρ is p Is the density of the thixotropic propellant, v is the feed rate of the thixotropic propellant, A is the total cross-sectional area of the conveying pipe, C * Characteristic speed of thixotropic propellants, A t Is the nozzle throat area;
s302: determining the cone height h:
wherein d is the diameter of the delivery conduit,
the burning speed is set;
s303: based on the cone height h and the diameter d of the conveying pipe, all three-dimensional features of the inner flow field area are determined.
Preferably, the step S4 specifically includes:
s401: neglecting components in front of a conveying pipeline of the thixotropic propellant rocket engine, and simplifying the geometric structure of the components into a combustion chamber and a spray pipe structure;
s402: the method comprises the steps that injection holes of an injector are periodically and circularly distributed at 60 degrees, one sixth area of the injection holes is set as a calculation domain, and a circular boundary is set as a periodic surface;
s403: according to the estimated flow rate and y + Determining the height of a first layer of computational grid according to the standard of 1, and applying boundary layer computational grids on all solid wall surfaces;
s404: and dividing the whole calculation domain by adopting a Poly-Hexcore technology to capture fine geometric features.
Preferably, the step S5 specifically includes: and importing the computational grid into CFD software, defining the whole computational domain as a fluid domain, and separating the first-layer computational grid on the surface of the grain for directionally adding source items to realize the introduction of pyrolysis gas of the grain.
Preferably, the source term in step S6 is determined according to the following formula:
wherein S is Quality of 、S Momentum 、S (Energy) 、S Components Respectively represents the mass, momentum, energy and component source terms added in the N-S control equation, V is the flow velocity,
representing the basis vector of the three-dimensional coordinate axis, Δ y being the height of the first layer of computational grid on the boundary, h g Is the enthalpy value, Y, of the column pyrolysis gas i Is the concentration of each component of the pyrolysis gas of the grain.
Preferably, the parameters in step S10 include: pressure, temperature, density, turbulence intensity, energy, component concentration.
Preferably, in step S7, the boundary conditions include an open boundary, a wall boundary, and a grain pyrolysis surface boundary.
Preferably, the operating characteristic parameters in step S11 include: three-dimensional flow line, vortex distribution, temperature distribution, density distribution, pressure distribution, gas component distribution and grain temperature information.
Compared with the prior art, the invention has the beneficial effects that:
the invention provides a prediction method for the operating characteristics of a thixotropic propellant rocket engine, which realizes accurate prediction of the three-dimensional characteristics of the thixotropic propellant rocket engine, overcomes the defects that the existing prediction method has single index and cannot accurately simulate three-dimensional combustion flow behaviors, has the advantages of completely matching calculation simulation with an actual physical model, high calculation speed and higher calculation precision, and effectively shortens development time, thereby reducing development cost.
Drawings
In order to illustrate embodiments of the invention or solutions in the prior art more clearly, the drawings that are needed in the embodiments will be briefly described below, so that the features and advantages of the invention will be more clearly understood by referring to the drawings that are schematic and should not be understood as limiting the invention in any way, and other drawings may be obtained by those skilled in the art without inventive effort.
FIG. 1 is a flow chart of a predictive method of operating characteristics of a thixotropic propellant rocket engine of the present invention;
FIG. 2 is a sub-flowchart of step S4 in FIG. 1;
FIG. 3 is a diagram illustrating boundary condition types and source item addition areas according to the present invention;
FIG. 4 is a three-dimensional flow chart calculated in example 1 of the present invention;
FIG. 5 is a two-dimensional flow chart calculated in example 1 of the present invention on a cross section at 25mm, 50mm from the injector surface;
fig. 6 is a graph showing the relationship between the pressure in the combustion chamber and the thrust force as a function of the flow rate in embodiment 1 of the present invention.
Detailed Description
In order that the above objects, features and advantages of the present invention can be more clearly understood, a more particular description of the invention will be rendered by reference to the appended drawings. It should be noted that the embodiments of the present invention and features of the embodiments may be combined with each other without conflict.
In the following description, numerous specific details are set forth in order to provide a thorough understanding of the present invention, however, the present invention may be practiced in other ways than those specifically described herein, and therefore the scope of the present invention is not limited by the specific embodiments disclosed below.
The flow of the prediction method of the operating characteristics of the thixotropic propellant rocket engine is shown in figure 1 and roughly comprises the following steps:
s1: acquiring geometric parameters and working condition parameters of a thixotropic propellant rocket engine and formula parameters of a propellant;
s2: carrying out burning rate measurement on the selected propellant to obtain a burning rate coefficient and a pressure index;
s3: calculating the trajectory in the internal flow field under the steady-state work to obtain the geometric characteristics of the internal flow field;
s4: simplifying a calculation domain according to the geometric characteristics of the internal flow field, carrying out space dispersion on the calculation domain, and deriving a calculation grid;
s5: importing the processed computational grid into CFD software, and defining the attribute of a computational domain;
s6: describing the flow state of the fluid by adopting a three-dimensional multi-component N-S control equation, wherein a turbulent flow is described by adopting an SST k-omega model; adding pyrolysis gas of the propellant grain into a calculation domain as a source item in an N-S control equation;
s7: defining boundary conditions including an open boundary, a wall surface boundary, a grain pyrolysis surface boundary and the like;
s8: defining a chemical reaction kinetic system in the combustion chamber by adopting an Eddy Dissipation Concept (EDC) model;
and S9, carrying out numerical calculation, carrying out space dispersion by adopting a second-order windward format, and carrying out coupling solution of pressure and density by adopting a coupling algorithm.
S10: starting three-dimensional combustion flow steady-state simulation calculation, monitoring flow, energy and component parameters, and judging whether the calculation is converged;
s11: and after calculation convergence, checking a result file, extracting information such as three-dimensional flow lines, vortex distribution, temperature distribution, density distribution, pressure distribution, gas component distribution, explosive column temperature and the like in the engine, and performing statistical analysis.
The sub-process of step S4 is shown in fig. 2, and specifically includes:
s401: the geometry is simplified to combustor and nozzle configurations;
s402: periodically simplifying a calculation domain according to an injector structure;
s403: according to y + Determining the height of the first layer grid as 1, and encrypting the solid boundary layer;
s404: and dividing the whole calculation domain by adopting a Poly-hexcore technology to capture fine geometric features.
Example 1
Taking a kilonewton-level thixotropic propellant rocket engine as an example, the prediction method provided by the invention is adopted for calculation, and comprises the following steps:
s1: obtaining geometric parameters and working condition parameters of a thixotropic propellant rocket engine and formula parameters of a propellant: the inner diameter of the combustion chamber adopted in the embodiment is 128mm, the length is 134mm, the diameter of the nozzle throat is 17mm, the ratio of the area of the outlet of the pipeline to the area of the nozzle throat is 5.4, and the conveying pipeline is a straight circular pipe with the diameter of 8 mm. The thixotropic propellant component is 70% of ammonium perchlorate, 10% of hydroxyl-terminated polybutadiene, 5% of metal particles and 15% of other additives.
S2: and (3) carrying out burning rate measurement on the selected propellant to obtain a burning rate coefficient and a pressure index: the burning rate coefficient of the propellant is measured to be 19.798mm s/Mpa, and the pressure index is measured to be 0.4249.
S3: calculating the trajectory in the internal flow field under the steady-state work to obtain the geometrical characteristics of the internal flow field: the pressure in the combustion chamber is determined by the following equation, where ρ p Is the density of the propellant, v is the propellant feed rate, A is the total cross-sectional area of the delivery line, C * For a particular speed of the propellant, A t Is the nozzle throat area.
The cone height h is then determined by the following equation, where d is the diameter of the transfer line,
the burning rate is determined by a Virginia burning rate formula.
And all three-dimensional characteristics of the flow field area can be obtained through two parameters of the cone height and the pipeline diameter.
S4: and simplifying a calculation domain according to the geometrical characteristics of the internal flow field, performing spatial dispersion on the calculation domain, and deriving a calculation grid. The method comprises the following four substeps:
s401: omitting the components before the rocket engine pipeline, and simplifying the geometric structure to a combustion chamber and a nozzle structure;
s402: since the injection holes of the injectors are periodically and cyclically distributed, the cycle period is 60 °. The computational domain is periodically simplified according to the injector structure, and only one sixth of the region is divided for computation. Correspondingly, as shown in fig. 3, the loop boundary is set to a periodic surface in order to implement a loop function in the calculation.
S403: according to the estimated flow rate and y + Determining the height of the first layer of grid as 1 standard, wherein the thickness of the first layer of grid is about 0.001mm in the embodiment, so as to encrypt the solid boundary layer, and applying boundary layer grids with the initial thickness of 0.001mm, the transition ratio of 1.2 and the number of layers of 10 on all solid wall surfaces;
s404: and dividing the whole calculation domain by adopting a Poly-Hexcore technology to capture fine geometric features. In the embodiment, a double-layer polyhedral mesh transition mode is adopted for division, and the maximum unit size is 8 mm.
S5: guiding the processed grids into CFD software, defining the whole calculation domain as a fluid domain, and separating the first layer of grids on the surface of the grain from the overall grids so as to directionally add source items in the next step and realize the introduction of pyrolysis gas of the grain;
s6: and describing the flow state of the fluid by adopting a three-dimensional multi-component N-S control equation, wherein an SST k-omega model is adopted to describe turbulent flow, and the charge pyrolysis gas is added into a calculation domain as a source term in the control equation. The mass, momentum, energy source terms are determined according to the following equations. Where Δ is the height of the first layer of mesh on the boundary, h g Is the enthalpy value, Y, of the column pyrolysis gas i Is the concentration of each component of the pyrolysis gas of the grain.
S7: boundary conditions are defined. The boundary conditions selected in this embodiment are shown in fig. 3, and include an open boundary, a wall boundary, a grain pyrolysis surface boundary, and a periodic cycle boundary.
S8: a vortex dissipation concept (EDC) model is used to define the chemical reaction kinetics within the combustion chamber. In the embodiment, the main components of the pyrolysis product of the grain are 1, 3-butadiene and oxygen, so that a simple chemical reaction system of the two is established and introduced into simulation software;
and S9, performing numerical calculation, performing space dispersion by adopting a second-order windward format, and performing coupling solution of pressure and density by adopting a coupling algorithm. In this example, a pressure convergence factor of 0.7, a convergence factor of 0.75 for all components, and a convergence factor of 0.6 for the turbulence-related parameter were used.
S10: starting simulation calculation, monitoring flow, energy and component parameters, and judging whether the calculation is converged; in this embodiment, the residual errors of pressure, temperature, density, turbulence intensity, energy, and component concentration are monitored, wherein the threshold value of the energy residual error is set to 1e -6 And residual thresholds of the rest parameters are set to be 0.001.
S11: after calculation convergence, checking a result file, extracting information such as three-dimensional flow lines, vortex distribution, temperature distribution, density distribution, pressure distribution, gas component distribution, explosive column temperature and the like in the engine, and performing statistical analysis; to show the numerical calculation results, fig. 4 shows a three-dimensional flow chart calculated in the present embodiment, fig. 5 is a two-dimensional flow chart on a cross section at a distance of 25mm and 50mm from the injector surface, and fig. 6 is a relationship between the combustor pressure and the thrust force with the flow rate in the present embodiment.
In the present invention, unless otherwise expressly stated or limited, the terms "mounted," "connected," "secured," and the like are to be construed broadly and can, for example, be fixedly connected, detachably connected, or integrally formed; can be mechanically or electrically connected; either directly or indirectly through intervening media, either internally or in any other relationship. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In the present invention, unless expressly stated or limited otherwise, the recitation of a first feature "on" or "under" a second feature may include the recitation of the first and second features being in direct contact, and may also include the recitation that the first and second features are not in direct contact, but are in contact via another feature between them. Also, the first feature being "on," "above" and "over" the second feature includes the first feature being directly on and obliquely above the second feature, or merely indicating that the first feature is at a higher level than the second feature. A first feature being "under," "below," and "beneath" a second feature includes the first feature being directly under and obliquely below the second feature, or simply meaning that the first feature is at a lesser elevation than the second feature.
In the present invention, the terms "first", "second", "third" and "fourth" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance. The term "plurality" means two or more unless expressly limited otherwise.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (10)
1. A method for predicting the operating characteristics of a thixotropic propellant rocket engine is characterized by comprising the following steps:
s1: acquiring geometric parameters and working condition parameters of a thixotropic propellant rocket engine and formula parameters of the thixotropic propellant;
s2: carrying out burning rate measurement on the thixotropic propellant to obtain a burning rate coefficient and a pressure index;
s3: under the steady-state work, calculating the trajectory in the inner flow field of the engine combustion chamber to obtain the three-dimensional geometric characteristics of the inner flow field;
s4: simplifying the internal flow according to the three-dimensional geometrical characteristics of the internal flow field, determining a calculation domain, performing spatial dispersion, and deriving a calculation grid;
s5: importing the computational grid into CFD software, and defining the attributes of the computational domain;
s6: describing the flow state of the fluid by using a three-dimensional multi-component N-S control equation, wherein a turbulent flow is described by using an SST k-omega model; adding pyrolysis gas of the propellant grain into a calculation domain as a source item in an N-S control equation;
s7: defining a boundary condition;
s8: defining a chemical reaction kinetic system in the combustion chamber by adopting an Eddy Dissipation Concept (EDC) model;
s9: performing space dispersion on the N-S control equation by adopting a second-order windward format, and performing coupling solution of pressure and density by adopting a coupling algorithm;
s10: starting three-dimensional combustion flow steady-state simulation calculation, and monitoring residual errors of all parameters until calculation results are converged;
s11: and extracting the working characteristic parameters in the thixotropic propellant engine from the calculation result, and performing statistical analysis to finish prediction.
2. The predictive method of operating characteristics of a thixotropic propellant rocket engine as recited in claim 1 wherein said geometric parameters of step S1 include the geometric dimensions of the combustion chamber, the nozzle throat diameter, the ratio of the delivery conduit exit area to the nozzle throat area, and the delivery conduit diameter.
3. A method of prognosticating the operational characteristics of a thixotropic propellant rocket engine according to claim 1 wherein the thixotropic propellant has the formulation parameters: 70% of ammonium perchlorate, 10% of hydroxyl-terminated polybutadiene, 5% of metal particles and 15% of other additives.
4. The method for predicting the operating characteristics of a thixotropic propellant rocket engine as recited in claim 1, wherein said step S3 specifically comprises:
s301: determining the pressure p in the combustion chamber eq ：
Where ρ is p Is the density of the thixotropic propellant, v is the feed rate of the thixotropic propellant, A is the total cross-sectional area of the conveying pipe, C * Characteristic speed of thixotropic propellants, A t Is the nozzle throat area;
s302: determining the cone height h:
wherein d is the diameter of the delivery conduit,
the burning rate is set;
s303: based on the cone height h and the diameter d of the conveying pipe, all three-dimensional features of the inner flow field area are determined.
5. The method for predicting the operating characteristics of a thixotropic propellant rocket engine as recited in claim 1, wherein said step S4 specifically comprises:
s401: neglecting components in front of a conveying pipeline of the thixotropic propellant rocket engine, and simplifying the geometric structure of the components into a combustion chamber and a spray pipe structure;
s402: the method comprises the steps that injection holes of an injector are periodically and circularly distributed at 60 degrees, one sixth area of the injection holes is set as a calculation domain, and a circular boundary is set as a periodic surface;
s403: according to the estimated flow rate and y + Determining the height of a first layer of computational grids according to a standard of 1, and applying boundary layer computational grids on all solid wall surfaces;
s404: and dividing the whole calculation domain by adopting a Poly-Hexcore technology to capture fine geometric features.
6. The method for predicting the operating characteristics of a thixotropic propellant rocket engine as recited in claim 1, wherein said step S5 specifically comprises: and introducing the computational grid into CFD software, defining the whole computational domain as a fluid domain, and separating the first layer of computational grid on the surface of the grain for directionally adding source items to realize the introduction of pyrolysis gas of the grain.
7. The predictive method of operating characteristics of a thixotropic propellant rocket engine of claim 4 wherein the source terms in said step S6 are determined according to the following formula:
wherein S is Quality of 、S Momentum 、S (Energy) 、S Components Respectively represents the mass, momentum, energy and component source terms added in the N-S control equation, V is the flow velocity,
representing the basis vector of the three-dimensional coordinate axis, Δ y is the height of the first layer of computational mesh on the boundary, h g Is the enthalpy value, Y, of the column pyrolysis gas i Is the concentration of each component of the pyrolysis gas of the grain.
8. The predictive method of operating characteristics of a thixotropic propellant rocket engine as recited in claim 1, wherein said parameters of step S10 include: pressure, temperature, density, turbulence intensity, energy, constituent concentration.
9. The predictive method for the operating characteristics of a thixotropic propellant rocket engine as recited in claim 1, wherein said boundary conditions include open boundaries, wall boundaries and grain pyrolysis surface boundaries in said step S7.
10. The predictive method of operating characteristics of a thixotropic propellant rocket engine according to claim 1 wherein said operating characteristic parameters of step S11 include: three-dimensional flow line, vortex distribution, temperature distribution, density distribution, pressure distribution, gas component distribution and grain temperature information.
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| CN118133634A (en) * | 2024-05-06 | 2024-06-04 | 西安现代控制技术研究所 | Method for evaluating integrity of charging structure of solid rocket engine |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9977848B1 (en) * | 2014-07-10 | 2018-05-22 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Method and system for predicting rocket nozzle deformation during engine start-up and shut-down transients |
| CN108304684A (en) * | 2018-03-22 | 2018-07-20 | 北京航天动力研究所 | A kind of rocket motor tail injection stream emulation mode and system |
| CN109753716A (en) * | 2018-12-28 | 2019-05-14 | 山东大学 | Core based on flow field simulation/thermal power steam turbine group fluid excitation numerical computation method and system |
| CN109800488A (en) * | 2019-01-02 | 2019-05-24 | 南京理工大学 | Numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment |
| CN110222467A (en) * | 2019-06-20 | 2019-09-10 | 北京理工大学 | A kind of underwater engine jet pipe swings the analysis method of high-speed jet characteristic |
-
2022
- 2022-06-15 CN CN202210685281.3A patent/CN115034157B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9977848B1 (en) * | 2014-07-10 | 2018-05-22 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Method and system for predicting rocket nozzle deformation during engine start-up and shut-down transients |
| CN108304684A (en) * | 2018-03-22 | 2018-07-20 | 北京航天动力研究所 | A kind of rocket motor tail injection stream emulation mode and system |
| CN109753716A (en) * | 2018-12-28 | 2019-05-14 | 山东大学 | Core based on flow field simulation/thermal power steam turbine group fluid excitation numerical computation method and system |
| CN109800488A (en) * | 2019-01-02 | 2019-05-24 | 南京理工大学 | Numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment |
| CN110222467A (en) * | 2019-06-20 | 2019-09-10 | 北京理工大学 | A kind of underwater engine jet pipe swings the analysis method of high-speed jet characteristic |
Non-Patent Citations (3)
| Title |
|---|
| JIAN DAI ET AL.: "Secondary fuel jet strategies on mixing enhancement performance of rocket-based combined cycle engine", 《ACTA ASTRONAUTICA》, 31 December 2021 (2021-12-31) * |
| 张帅: "固液混合火箭发动机工作过程与燃烧组织数值模拟研究", 《中国博士学位论文全文数据库 工程科技II》, 15 February 2020 (2020-02-15) * |
| 程玉峰 等: "电喷雾推力器锥射流形成的电流体力学仿真", 《南昌大学学报（理科版）》, 30 April 2022 (2022-04-30) * |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN118133634A (en) * | 2024-05-06 | 2024-06-04 | 西安现代控制技术研究所 | Method for evaluating integrity of charging structure of solid rocket engine |
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