CN109800488A - Numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment - Google Patents
Numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment Download PDFInfo
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Abstract
The invention discloses a kind of numerical computation methods about liquid rocket high altitude environment lower bottom part thermal environment, initially set up liquid rocket 3-D geometric model；Then grid dividing is carried out to threedimensional model using multiple blockstructured grid method and encrypted；It resettles convection current/heat radiation coupling model of the carrier rocket combustion gas plume of high-altitude free flow containing supersonic speed: carrying out again, to the convective term in N-S equation discrete: discrete using Second-order Up-wind TVD format；Radiation patterns are finally established using discrete ordinates method, large-scale parallel computation, output Mach number, temperature, pressure flow field and convection current/coupling hot-fluid cloud atlas are carried out to rocket whole flow field and its Base Heat stream.The present invention provides a kind of high-precision, lower calculating cost and meet the actual numerical value emulation method of engineering.
Description
Technical field
The invention belongs to supersonic aircraft Simulation of Thermal Environment fields, and especially one kind is about the high idle loop of liquid rocket
The numerical computation method of border lower bottom part thermal environment.
Background technique
In recent years, as a series of great solar-system operations such as manned moon landing, space station establishment, deep space exploration are carried out in succession,
China increases the research work to heavy launcher and its dynamical system, it is determined that oxygen kerosene engine and liquid hydrogen liquid oxygen
Engine is the optimal selection of Launch Vehicle Force system.Liquid launch vehicle engine at work, due to external environment pressure
Power is too low, and combustion gas stream sharply expands after entering external environment, is formed and is flowed back in rocket bottom, and forms convection current to rocket body bottom and add
Fuel factor, meanwhile, red-hot CO is sprayed in high-temperature fuel gas2、H2The mixed airflows such as O form radiant heating effect to rocket bottom.Fire
Arrow bottom is equivalent to leeward, is easy Convective Heating and radiant heating coupling caused by being flowed back by jet cutting car flow, causes
Temperature increases rapidly.Great threat can be brought to the safety of rocket entirety to rocket bottom thermal environment underrating, or even lured
The major accidents such as detonation is fried cause flight to fail, and overestimate will lead to heat protective provisions conservative design, increase emit at
This, therefore, carrying out analysis of Thermal Environment research to liquid launch vehicle bottom becomes the current most important thing.
Compared to foreign countries to rocket bottom thermal environment research degree, the country is at present also in early stage, especially China
Liquid launch vehicle test flight data is few, domestic scholars be mostly numerical simulation to rocket bottom thermal environment in terms of research,
Lacking comparative analysis causes the validity of emulation mode to be difficult to be verified.And since computing resource is limited, so most of
Scholar's logarithm model has carried out different degrees of simplification, as geometrical model only considers the boundless interlayer of a quarter, grid, heat radiation
Lower P-1 model of precision etc. is used in calculating, these all largely affect computational accuracy.
With the continuous development of Fluid Mechanics Computation and the continuous improvement of computer performance, numerical simulation has become
Study a kind of effective means in flow field.And with the development of Space Science and Technology, rocket load increases, and volume increases, and leads to rocket
Whole Flow Field Calculation region is big, and grid scale is big, and calculation amount increases, and needs to solve by Large-scale parallel computing, at this moment just compels
It is essential and wants a kind of high-precision, the low numerical value emulation method for calculating cost simulates liquid rocket high altitude environment lower bottom part thermal environment.
Summary of the invention
The purpose of the present invention is to provide a kind of numerical value calculating sides about liquid rocket high altitude environment lower bottom part thermal environment
Method, to realize the forecasting problem of thermal environment under liquid rocket high altitude environment.
Realize the technical solution of the object of the invention are as follows:
A kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment, which is characterized in that including with
Lower step:
Step 1 establishes liquid rocket 3-D geometric model；
Step 2 carries out grid dividing to threedimensional model using multiple blockstructured grid method；
Step 3, establish high-altitude free flow containing supersonic speed carrier rocket combustion gas plume convection current/heat radiation coupling model:
Navier-Stokes equation, Realizable k- ε two equation turbulence model are transported based on combustion gas multicomponent；
Step 4, in N-S equation convective term carry out it is discrete: using Second-order Up-wind TVD format it is discrete；
Step 5, radiation patterns use discrete ordinates method (Discrete-Ordinates Methods, DOM), complete to rocket
Flow field and its Base Heat stream carry out large-scale parallel computation, output Mach number, temperature, pressure flow field and convection current/coupling hot-fluid cloud atlas.
Compared with prior art, the present invention its remarkable advantage is:
(1) a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment of the invention, to grid
It divides and uses multiplet grid method, efficiently solve that liquid rocket entirety flow field regions are big, grid scale is big, are difficult to
The problem of calculating, while solving the problems, such as that convection current/radiation Coupled Heat Transfer is difficult to calculate；
(2) a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment of the invention, use
Second-order Up-wind TVD format is most suitable for solving height compressible flow problem, and automatic capturing shock ability is strong, due to devising robustness
Strong flux limiter can not only acquire high-resolution complex wave architecture, and can inhibit the non-physical shake of strong discontinuity solution
It swings；
(3) a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment of the invention, turbulent flow mould
Type uses Realizable k- ε two-equation model, compared with standard k- ε two-equation model, in turbulence dissipative shock wave ε equation
Generating item no longer includes the generation item G in tubulence energy k equationk, and turbulence viscosity, mutIn coefficient CμNot instead of constant,
Related to strain rate, such form preferably indicates that energy is converted；
(4) a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment of the invention, to radiation
Hot-fluid is solved using discrete ordinates method (DOM), and this method, which has, is easily handled scattering problems, is easy to and flow equation simultaneous
Solution and the higher feature of computational accuracy.
Detailed description of the invention
Fig. 1 is a kind of numerical computation method structure flow chart about liquid rocket high altitude environment lower bottom part thermal environment.
Fig. 2 is liquid rocket three-dimensional model diagram.
Fig. 3 is the multiple blockstructured grid chart of multi nozzle liquid rocket.
Fig. 4 is multi nozzle liquid rocket Mach number field cloud atlas.
Fig. 5 is multi nozzle liquid rocket temperature field cloud atlas.
Fig. 6 is multi nozzle liquid rocket pressure field cloud atlas.
Fig. 7 is multi nozzle liquid rocket bottom convection current/radiant heat flux coupling cloud atlas.
Fig. 8 is multi nozzle liquid rocket bottom monitoring point value analog result and measured result contrast schematic diagram.
Specific embodiment
In order to illustrate technical solution of the present invention and technical purpose, with reference to the accompanying drawing and specific embodiment is the present invention
It is further to introduce.
In conjunction with Fig. 1, a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment of the invention, packet
Include following steps:
Step 1 establishes liquid rocket 3-D geometric model；
1.1, modeling rendering is carried out according to Real Rocket 1:1；
1.2, in conjunction with Fig. 2,3-D geometric model needs following parameter: -1 height of core level segment and radius, core level segment bullet -2
Curvature, -3 length of jet pipe, -3 entrance of jet pipe and exit radius, -4 radius of nozzle throat.
Step 2 carries out grid dividing to threedimensional model using multiple blockstructured grid method and encrypts；
2.1, in conjunction with Fig. 3, piecemeal processing is carried out to multi nozzle rocket threedimensional model, overall computational domain is divided into two sons
Domain: core level segment and the foreign lands -1 around core level segment；Combustion gas plume region -2 below jet pipe and jet pipe；
2.2, jet pipe and plume area grid are encrypted, entire computational domain is all made of structured grid to guarantee net
The orthogonality and fairness of lattice.In numerical value calculating simultaneously as the physical quantity variations such as temperature, pressure of boundary are violent, wall surface
Function cannot accurately calculate the flowing of near wall, therefore further encrypt to the grid of boundary layer computational domain, it is ensured that wall surface
Neighbouring grid Y+ value is less than 3.
Step 3, establish high-altitude free flow containing supersonic speed carrier rocket combustion gas plume convection current/heat radiation coupling model:
Navier-Stokes equation, Realizable k- ε two equation turbulence model are transported based on combustion gas multicomponent, establishes high-altitude containing super
Convection current/heat radiation coupling model of the carrier rocket combustion gas plume of velocity of sound free flow；
Since multi nozzle rocket mostly uses gaseous fuel, gas component includes H2O, CO2, CO, H2, N2Deng, therefore using combustion
Gas and air multi-component stream movable model.
3.1, jet-flow satisfaction is set: continuous perfect gas and component is directly without chemical reaction；That is:
(1) jet-flow meets continuous media；(2) jet-flow is compressible pure gas phase media；(3) inside jet-flow
Occur without chemical reaction；(4) ideal gas behavior is all satisfied using combustion gas and air multiple groups part mixed flow model, Multicomponent
Equation；
3.2, establishing combustion gas multiple groups part transport equation is
Wherein, YlFor the mass fraction of liquid rocket Gas Components l, RlAfter chemical reaction for liquid rocket Gas Components l
Net production rate, SlProduction rate caused by discrete phase for customized source item.JlFor liquid rocket Gas Components diffusion flux, t is
Rocket flight time, ρ are the fluid density of rocket combustion gas, and it is the divergence of combustion gas stream micro unit that v, which is rocket velocity vector,；
Wherein rocket combustion gas component diffusion flux is
In formula, DL, mFor the mass loss coefficient of liquid rocket combustion gas component l, DT, lFor the heat of liquid rocket combustion gas component l
Diffusion coefficient；
3.3, under rectangular coordinate system, the compressible Navier-Stokes equations model of one-component l is established:
In formula (3)-(6), U is liquid rocket fuel gas flow variable；F, G, H are liquid rocket fuel gas flow momentum flow vector,
Fv、Gv、HvFor liquid rocket combustion gas stickiness momentum flow vector, K is the liquid rocket combustion gas coefficient of heat conduction；T is environment temperature；p,ρ,
E, τ, μ are respectively rocket gaseous-pressure, density, than kinetic energy, stress, viscosity coefficient, and u, v, w are respectively liquid rocket combustion gas speed
Component in the x, y, z-directions, rocket bottom centre are origin, and rocket flight direction is the direction z, non-conterminous nozzle entry center
Line is respectively the direction x and y；
3.4, the turbulence model of plume during rocket flight is established using Realizable k- ε two-equation model:
Compared with standard k- ε two-equation model, which has the flow characteristics such as strong flow curvature and vortex apparent
Description；
Tubulence energy k equation is with turbulence dissipative shock wave ε equation
In formula (7) and (8), GkFor the generation item of tubulence energy k caused by average velocity gradient, μtFor turbulence stickiness, σkAnd σε
It is the Prandtl number of tubulence energy k and turbulence dissipative shock wave ε, C respectively1And C2For equation constant coefficient, C1=1.44, C2=1.9.
Step 4, in N-S equation convective term carry out it is discrete: using Second-order Up-wind TVD format it is discrete；
Carrier rocket uses high thrust motor design scheme, and combustion chamber total temperature, stagnation pressure are high, and jet flow is that the deficient expansion of height is super
Velocity of sound jet stream, structure is complicated for wave system, and Second-order Up-wind TVD format is most suitable for solving height compressible flow problem；
4.1, numerical integration is carried out to each volume mesh unit using finite volume method:
Volume of the flow variables U at body unit center is average are as follows:
In formula, Vi,j,kTo calculate volume mesh unit.
4.2, convective flux is carried out using Second-order Up-wind TVD format discrete.
In formula, numerical value circulationIt is respectively as follows:
In formula, λk, μk, νkFor the characteristic value of linearisation replacement matrix, αk、βk、γkFor the expansion item of linearisation replacement matrix
Coefficient, ek(A)、ek(B)、ekIt (C) is the feature vector of linearisation replacement matrix, i, j, k are respectively the unit on x, y, z direction
Vector, stickiness flux are discrete using Second-Order Central Difference.
Step 5 establishes radiation patterns using discrete ordinates method (Discrete-Ordinates Methods, DOM), to fire
Arrow whole flow field and its Base Heat stream carry out large-scale parallel computation, output Mach number, temperature, pressure flow field and convection current/coupling hot-fluid
Cloud atlas.
5.1, the Integrated Derivative fundamental equation of radiant heat transfer is defined:
Component analysis is carried out to engine gas, obtains the molar percentage of rocket Gas Components, and measures and obtains spray more
Pipe rocket environment temperature, flying speed, environmental stress, jet pipe internal pressure, jet pipe wall surface and rocket body bottom initial temperature and
Jet pipe internal temperature substitutes into the Integrated Derivative fundamental equation of radiant heat transfer:
In formula: s is direction vector；For combustion gas diffusion term, I is gas medium radiation intensity；IbFor the spoke of black matrix
Penetrate intensity；ka, ksThe respectively absorption coefficient and scattering coefficient of multi nozzle rocket jet wake flow medium；For combustion gas phase
Function；
5.2, with DOM model discrete irradiation heat transfer basic equation:
Discrete ordinates method has the characteristics that computational accuracy is higher and is easy to and flow equation simultaneous solution；
Discrete obtain is carried out along the direction s by equation of radiative transfer:
Spectral intensity I can be obtained by formula (13)λThe equation of radiative transfer of (r, s) are as follows:
In formula, r is position vector, and s is direction vector, and s ' is scattering direction vector, and s is gas layer thickness, and a is to absorb system
Number, n are refractive index, and σ is black body radiation constant, σsFor scattering coefficient, I is spectral radiance, and T is the thermodynamics temperature of black matrix
Degree, Φ are Scattering Phase Function, and Ω ' is solid angle, and λ is wavelength, IbλThe black matrix intensity given for Planck equation.
5.3, large-scale parallel computation is carried out to rocket whole flow field and its Base Heat stream, exports Mach number, temperature, pressure flow field
And convection current/coupling hot-fluid cloud atlas
Embodiment
A kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment, by above-mentioned specific embodiment
Middle step, comprising the following steps:
Step 1 establishes liquid rocket 3-D geometric model；
In conjunction with Fig. 2,3-D geometric model needs following parameter: -1 height of core level segment and radius, -2 curvature of core level segment bullet,
- 3 length of jet pipe, -3 entrance of jet pipe and exit radius, -4 radius of nozzle throat, 1:1 is modeled after determining parameter.
Step 2 carries out grid dividing to threedimensional model using multiple blockstructured grid method；
In conjunction with Fig. 3, multiple blockstructured grid is carried out to geometrical model and is handled, net is divided to multi nozzle liquid rocket
Lattice, last grid sum is 8,920,000 or so.
Step 3 transports Navier-Stokes equation, two equation turbulent flow mould of Realizable k- ε based on combustion gas multicomponent
Type establishes convection current/heat radiation coupling model of the carrier rocket combustion gas plume of high-altitude free flow containing supersonic speed；
Combustion gas component includes H2O、CO2、CO、H2, molal weight percentage is respectively 38%, 26%, 25%, 9%, 2%,
Substitute into multicomponent equation solution.
Step 4, use Second-order Up-wind TVD format discrete to the convective term in N-S equation；
Step 5, radiation patterns use discrete ordinates method, carry out large-scale parallel meter to rocket whole flow field and its Base Heat stream
It calculates, output Mach number, temperature, pressure flow field and convection current/coupling hot-fluid cloud atlas.
Input following parameter:
Flying height: 20km environment temperature: 216k flying speed: 1.2Ma environmental stress: 6587Pa
Jet pipe internal pressure: 1.8*107Pa jet pipe wall surface and rocket body bottom initial temperature: 300k
Jet pipe internal temperature: 3400k radiation absorption factor: 0.3
A kind of designed numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment is in benefit herein
With time-consuming in the case where 32 CPU parallel computations two days output Mach number fields, temperature field, pressure fields, total hot-fluid cloud charts, such as
Shown in the following figure, Fig. 4 shows that multi nozzle liquid rocket Mach wave is expanded and interacted, and Fig. 5 shows the multi nozzle liquid fire
Arrow temperature maximum crosses position in nozzle hole and shock wave, and Fig. 6 shows multi nozzle liquid rocket bottom plume intersection pressure
Height, caused by plume interferes with each other, Fig. 7 show rocket bottom hot-fluid it is bigger closer to center numerical value, and it is rounded to
It is uniformly distributed outside, Fig. 8 shows that numerical result and the similar engineering test result goodness of fit are high, shows that this emulation mode has
Higher precision.The method of the present invention can reduce while improving computational accuracy calculates cost, and calculated result can be rocket bottom
Portion's thermal protection provides guidance.
Claims (6)
1. a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment, which is characterized in that including following
Step:
Step 1 establishes liquid rocket 3-D geometric model；
Step 2 carries out grid dividing to threedimensional model using multiple blockstructured grid method and encrypts；
Step 3, establish high-altitude free flow containing supersonic speed carrier rocket combustion gas plume convection current/heat radiation coupling model: be based on
Combustion gas multicomponent transports Navier-Stokes equation, Realizable k- ε two equation turbulence model, establishes high-altitude containing supersonic speed
Convection current/heat radiation coupling model of the carrier rocket combustion gas plume of free flow；
Step 4, in N-S equation convective term carry out it is discrete: using Second-order Up-wind TVD format it is discrete；
Step 5 establishes radiation patterns using discrete ordinates method, carries out large-scale parallel meter to rocket whole flow field and its Base Heat stream
It calculates, output Mach number, temperature, pressure flow field and convection current/coupling hot-fluid cloud atlas.
2. a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment according to claim 1,
It is characterized in that, step 1 establishes liquid rocket 3-D geometric model, specifically includes the following steps:
1.1, modeling rendering is carried out according to Real Rocket 1:1；
1.2,3-D geometric model needs following parameter: core level segment height and radius, core level segment bullet curvature, jet pipe length, spray
Tube inlet and exit radius, nozzle throat radius.
3. a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment according to claim 1,
It is characterized in that, step 2 carries out grid dividing to threedimensional model using multiple blockstructured grid method, specifically include following
Step:
2.1, piecemeal processing is carried out to multi nozzle rocket threedimensional model, overall computational domain is divided into two subdomains: core level segment and ring
Around the foreign lands of core level segment；Combustion gas plume region below jet pipe and jet pipe；
2.2, jet pipe and plume area grid are encrypted, the grid of boundary layer computational domain is further encrypted.
4. a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment according to claim 1,
It is characterized in that, step 3 establishes convection current/heat radiation coupled mode of the carrier rocket combustion gas plume of high-altitude free flow containing supersonic speed
Type, the specific steps are as follows:
3.1, jet-flow satisfaction is set: continuous perfect gas and component is directly without chemical reaction；
3.2, establishing combustion gas multiple groups part transport equation is
Wherein, YlFor the mass fraction of liquid rocket Gas Components l, RlFor liquid rocket Gas Components l after chemical reaction net
Production rate, SlProduction rate caused by discrete phase for customized source item.JlFor liquid rocket Gas Components diffusion flux, t is rocket
Flight time, ρ are the fluid density of rocket combustion gas, and v is rocket velocity vector；
3.3, under rectangular coordinate system, the compressible Navier-Stokes equations model of one-component l is established:
Wherein U is liquid rocket fuel gas flow variable；F, G, H are liquid rocket fuel gas flow momentum flow vector, Fv、Gv、HvFor liquid
Rocket combustion gas stickiness momentum flow vector；
3.4, the turbulence model of plume during rocket flight is established using Realizable k- ε two-equation model:
Tubulence energy k equation is with turbulence dissipative shock wave ε equation
In formula (7) and (8), GkFor the generation item of tubulence energy k caused by average velocity gradient, μtFor turbulence stickiness, σkAnd σεRespectively
It is the Prandtl number of tubulence energy k and turbulence dissipative shock wave ε, C1And C2For equation constant coefficient.
5. a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment according to claim 4,
It is characterized in that, step 4 is discrete to the convective term in N-S equation, specifically includes the following steps:
4.1, numerical integration is carried out to each volume mesh unit using finite volume method:
Volume of the flow variables U at body unit center is average are as follows:
In formula, Vi,j,kTo calculate volume mesh unit,
4.2, convective flux is carried out using Second-order Up-wind TVD format discrete；
In formula, λk, μk, νkFor the characteristic value of linearisation replacement matrix, αk、βk、γkFor the expansion term system of linearisation replacement matrix
Number, ek(A)、ek(B)、ek(C) be the feature vector of linearisation replacement matrix, i, j, k be respectively unit on x, y, z direction to
Amount,It is discrete using Second-Order Central Difference for the numerical value circulation stickiness flux of equation.
6. a kind of numerical computation method about liquid rocket high altitude environment lower bottom part thermal environment according to claim 5,
It is characterized in that, step 5 radiation patterns use discrete ordinates method:
5.1, the Integrated Derivative fundamental equation of radiant heat transfer is defined:
Establish the Integrated Derivative fundamental equation of radiant heat transfer:
In formula: s is direction vector；For combustion gas diffusion term, I is gas medium radiation intensity；IbIt is strong for the radiation of black matrix
Degree；ka, ksThe respectively absorption coefficient and scattering coefficient of multi nozzle rocket jet wake flow medium；For combustion gas phase function；
5.2, with DOM model discrete irradiation heat transfer basic equation:
Spectral intensity IλThe equation of radiative transfer of (r, s) are as follows:
In formula, r is position vector, and s is direction vector, and s ' is scattering direction vector, and s is gas layer thickness, and a is absorption coefficient, n
For refractive index, σ is black body radiation constant, σsFor scattering coefficient, I is spectral radiance, and T is the thermodynamic temperature of black matrix, Φ
For Scattering Phase Function, Ω ' is solid angle, and λ is wavelength, IbλThe black matrix intensity given for Planck equation.
5.3, large-scale parallel computation is carried out to rocket whole flow field and its Base Heat stream, simulation result can be obtained after calculating.
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110457804A (en) * | 2019-07-31 | 2019-11-15 | 南京理工大学 | Predict the numerical method of single spraying pipe carrier rocket jet noise |
| CN111950148A (en) * | 2020-08-11 | 2020-11-17 | 江苏深蓝航天有限公司 | Method and device for calculating inner wall temperature of test run of liquid rocket thrust chamber |
| CN112464358A (en) * | 2020-10-27 | 2021-03-09 | 中国运载火箭技术研究院 | Method, device, terminal and medium for evaluating environmental temperature of carrier rocket equipment |
| CN113550841A (en) * | 2021-09-06 | 2021-10-26 | 中国人民解放军战略支援部队航天工程大学 | Gas rocket engine for unmanned aerial vehicle launching and design method |
| CN114254572A (en) * | 2021-12-16 | 2022-03-29 | 西北工业大学太仓长三角研究院 | Aero-engine compressor flow field performance prediction method and system considering pollutant deposition |
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| CN108304595A (en) * | 2017-05-04 | 2018-07-20 | 北京空天技术研究所 | A kind of structure temperature analysis method for the semiclosed region of hypersonic aircraft |
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- 2019-01-02 CN CN201910001740.XA patent/CN109800488B/en active Active
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN108304595A (en) * | 2017-05-04 | 2018-07-20 | 北京空天技术研究所 | A kind of structure temperature analysis method for the semiclosed region of hypersonic aircraft |
Cited By (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110457804A (en) * | 2019-07-31 | 2019-11-15 | 南京理工大学 | Predict the numerical method of single spraying pipe carrier rocket jet noise |
| CN111950148A (en) * | 2020-08-11 | 2020-11-17 | 江苏深蓝航天有限公司 | Method and device for calculating inner wall temperature of test run of liquid rocket thrust chamber |
| CN111950148B (en) * | 2020-08-11 | 2023-10-03 | 江苏深蓝航天有限公司 | Method and equipment for calculating temperature of inner wall of liquid rocket thrust chamber test run |
| CN112464358A (en) * | 2020-10-27 | 2021-03-09 | 中国运载火箭技术研究院 | Method, device, terminal and medium for evaluating environmental temperature of carrier rocket equipment |
| CN113550841A (en) * | 2021-09-06 | 2021-10-26 | 中国人民解放军战略支援部队航天工程大学 | Gas rocket engine for unmanned aerial vehicle launching and design method |
| CN113550841B (en) * | 2021-09-06 | 2022-04-12 | 中国人民解放军战略支援部队航天工程大学 | Gas rocket engine for unmanned aerial vehicle launching and design method |
| CN114254572A (en) * | 2021-12-16 | 2022-03-29 | 西北工业大学太仓长三角研究院 | Aero-engine compressor flow field performance prediction method and system considering pollutant deposition |
| CN114254572B (en) * | 2021-12-16 | 2024-01-02 | 西北工业大学太仓长三角研究院 | Method and system for predicting flow field performance of aero-compressor by considering pollutant deposition |
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