CN111210503A - Numerical simulation method for liquid rocket afterburning reaction calculation - Google Patents

Numerical simulation method for liquid rocket afterburning reaction calculation Download PDF

Info

Publication number
CN111210503A
CN111210503A CN201911342880.XA CN201911342880A CN111210503A CN 111210503 A CN111210503 A CN 111210503A CN 201911342880 A CN201911342880 A CN 201911342880A CN 111210503 A CN111210503 A CN 111210503A
Authority
CN
China
Prior art keywords
rocket
liquid
model
reaction
numerical simulation
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
CN201911342880.XA
Other languages
Chinese (zh)
Inventor
周志坛
乐贵高
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Nanjing University of Science and Technology
Original Assignee
Nanjing University of Science and Technology
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Nanjing University of Science and Technology filed Critical Nanjing University of Science and Technology
Priority to CN201911342880.XA priority Critical patent/CN111210503A/en
Publication of CN111210503A publication Critical patent/CN111210503A/en
Withdrawn legal-status Critical Current

Links

Images

Classifications

    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06T—IMAGE DATA PROCESSING OR GENERATION, IN GENERAL
    • G06T17/00—Three dimensional [3D] modelling, e.g. data description of 3D objects
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

Landscapes

  • Physics & Mathematics (AREA)
  • Engineering & Computer Science (AREA)
  • Computer Graphics (AREA)
  • Geometry (AREA)
  • Software Systems (AREA)
  • General Physics & Mathematics (AREA)
  • Theoretical Computer Science (AREA)
  • Management, Administration, Business Operations System, And Electronic Commerce (AREA)

Abstract

The invention discloses a numerical simulation method for liquid rocket afterburning reaction calculation, which takes a liquid carrier rocket launching platform as a research object and analyzes the influence of the afterburning reaction on a tail flame impact flow field; the numerical simulation adopts a three-dimensional compressible Reynolds average N-S equation and a realizable k-epsilon turbulence model to obtain a gas jet flow field, and simultaneously adopts a finite rate chemical dynamics model to simulate a chemical reaction process, so that a multi-component tail flame reburning model is established, and compared with literature data, the effectiveness and the correctness of the algorithm are verified. The invention is a numerical simulation method which is high in precision, low in calculation cost and in accordance with the actual engineering and is established after a large number of numerical simulation tests are carried out and numerical results are compared with test results.

Description

Numerical simulation method for liquid rocket afterburning reaction calculation
Technical Field
The invention belongs to the technical field of thermal protection design of a launch platform of a carrier rocket, and particularly relates to a numerical simulation method for liquid rocket afterburning reaction calculation.
Background
In order to realize the deep space exploration of moon, wooden star, mars and the like as well as the manned interstellar landing plan, develop and utilize space resources early, a new generation of carrier rocket technology is developed by some aerospace technology universities in the world so as to improve the effective load. The heavy carrier rocket technology becomes an international research hotspot, and the high-thrust rocket is mainly characterized in that: first, the configuration surrounds the bundled booster and the multi-stage combined configuration to improve mission compliance. The energy rocket, warrior V rocket, H-IIA rocket, heavy falcon carrier rocket and the like after the 80 th century in the 20 th century all adopt the design scheme of parallel connection and multistage combination of a plurality of engines, thereby effectively shortening the total length of the rocket and ensuring the flight stability. However, when the core-stage engine and the boosting engine of the binding section work simultaneously, a plurality of jets at the bottom of the rocket interfere with each other, and meanwhile, the air coming from the outside extrudes the jet at the tail part, so that the wave system structure of the whole flow field is very complex. Secondly, adopt liquid engine as power device, liquid fuel has than high, operating time is long, can start advantages such as many times in the air. Meanwhile, the tail flame of the liquid rocket engine has the characteristics of high temperature, high speed and large flow, the ejected high-temperature fuel gas belongs to rich fuel gas, and the incompletely combusted high-temperature combustible gas in the tail flame is mixed with oxygen in the atmosphere to easily generate secondary combustion (afterburning), so that the temperature field of the tail flame and the distribution of fuel gas components are influenced, therefore, the fuel gas afterburning reaction must be considered in predicting the thermal environment of the rocket launching stage, and the establishment of an effective liquid rocket afterburning calculation numerical model has important significance for the thermal protection design of the rocket launching platform.
However, most of the research objects of the related work at present are single-engine solid launch vehicles (Jiang Y, Ma Y, WangW, Shao L., "Inhibition effect of water injection on after burning of rock nozzle exhaust plunger," Chinese Journal and Aeronautics, Vol.23,2010, pp.653-659.), and the research on the influence of the re-combustion reaction on the tail flame flow field of the multi-engine liquid launch vehicle is not complete enough. Meanwhile, due to the limitation of computational resources, most scholars simplify the Numerical modeling to a great extent, for example, only considering a simple two-dimensional model, quarter simplifying a geometric model, calculating a grid without boundary layer (Negishi H, Yamanishi N, Arita M, Namura E, Ohkubo s., "Numerical analysis of complex visualization for H-IIA Launch vertical powered estimate," 43rd AIAA/ASME/SAE/ASEE Joint prediction reference and inhibition, AIAA Paper 2007 and 5505, July2007.), and these processing methods all affect the computational accuracy to different degrees.
Disclosure of Invention
The invention aims to provide a numerical simulation method for liquid rocket afterburning reaction calculation so as to realize reasonable design of thermal protection of a launching platform at the take-off stage of a liquid carrier rocket.
The technical solution for realizing the purpose of the invention is as follows:
the numerical simulation method for the liquid rocket afterburning reaction calculation comprises the following steps:
step 1, establishing a three-dimensional geometric model of a gas impact flow guide device in a liquid carrier rocket launching stage;
step 2, carrying out structural grid division on the liquid rocket impact flow guide device by adopting a grid block scheme;
step 3, establishing a liquid rocket fuel gas impact model: establishing a rocket gas impact model based on a three-dimensional compressible Reynolds average N-S equation and a readable k-epsilon two-stroke turbulence model;
step 4, establishing a liquid rocket fuel gas chemical dynamics model: carrying out numerical simulation on the re-combustion reaction of the rocket gas based on the finite rate chemical dynamics model, and establishing a rocket gas chemical dynamics model;
step 5, carrying out parallel calculation on the gas impact flow field of the liquid carrier rocket, and outputting a Mach number cloud picture, a temperature cloud picture, CO and CO2、H2、H2And O mole fraction cloud.
Compared with the prior art, the method has the following advantages:
(1) the numerical simulation method adopts a Reynolds average N-S equation, has the characteristics of higher precision and low calculation cost, and the calculation result meets the engineering standard.
(2) The Realizable k-epsilon two-equation turbulence model adopted by the invention has accurate prediction on the divergence ratio of the flat jet flow and has higher accuracy on the calculation of boundary layer flow, flow separation and secondary flow.
(3) The finite rate chemical dynamics model adopted by the method is based on the Arrhenius equation, so that under the condition that the temperature of the fuel gas is increased sharply, the numerical method can consider the relationship between the temperature and the chemical reaction, and the correct calculation of the chemical reaction rate is facilitated.
Drawings
FIG. 1 is a flow chart of a numerical simulation method for liquid rocket afterburning reaction calculation.
FIG. 2 is a three-dimensional geometric model diagram of an impact diversion device at the launching stage of a liquid carrier rocket.
FIG. 3 is a grid diagram of the impact diversion device at the launching stage of the liquid carrier rocket.
FIG. 4 is a Mach field cloud diagram of a afterburning flow field of the impact flow guiding device at the launching stage of the liquid carrier rocket.
Fig. 5 is a cloud picture of a afterburning flow field temperature field of the impact flow guiding device at the launching stage of the liquid carrier rocket.
FIG. 6 is a CO mole fraction cloud chart of a afterburning flow field of the impact diversion device at the launching stage of the liquid carrier rocket.
FIG. 7 is a afterburning flow field CO of the impact flow guiding device at the launching stage of the liquid carrier rocket2Molar fraction cloud plots.
FIG. 8 is a afterburning flow field H of the impact flow guiding device at the launching stage of the liquid carrier rocket2Molar fraction cloud plots.
FIG. 9 is a afterburning flow field H of the impact flow guiding device at the launching stage of the liquid carrier rocket2And O mole fraction cloud.
FIG. 10 is a graph comparing the post combustion pressure flow field calculated using numerical methods herein with experimental results.
Detailed Description
With reference to fig. 1, the numerical simulation method for liquid rocket afterburning reaction calculation of the present invention includes the following steps:
step 1, establishing a three-dimensional geometric model of a gas impact flow guide device in a liquid carrier rocket launching stage;
1.1, combining with a figure 2, drawing a three-dimensional model according to a liquid carrier rocket 1:1, and needing the following parameters: the curvature and the radius of a carrier rocket warhead-1, the height and the radius of a rocket body-2, the overall length of a Laval nozzle-3, the radius of a nozzle inlet-4, the radius of a nozzle throat-5 and the radius of a nozzle outlet-6;
1.2, drawing a three-dimensional model according to a flow guide device 1:1 by combining with the figure 2, wherein the following parameters are required: length and width of inlet-7 of the flow guide device, length and width of outlet-8 of the flow guide device, and curvature of bottom-9 of the flow guide device.
Step 2, carrying out grid division on the three-dimensional model by using a multiple-block structured grid method and carrying out encryption;
2.1, combining with a figure 3, carrying out grid partitioning on the three-dimensional model of the liquid rocket impact flow guiding device, and dividing the whole calculation area into a carrier rocket sub-area-A, a transition sub-area-B of a rocket engine and the flow guiding device and a flow guiding device sub-area-C;
and 2.2, encrypting the sub-domain of the transition of the rocket engine and the flow guiding device and the sub-domain grids of the flow guiding device, and gradually transitioning the sub-domain grids of the carrier rocket from dense to sparse.
Step 3, establishing a liquid rocket fuel gas impact model: establishing a rocket gas impact model based on a three-dimensional compressible Reynolds average N-S equation and a readable k-epsilon two-stroke turbulence model;
3.1, establishing a three-dimensional compressible Reynolds average N-S equation of the liquid rocket fuel gas:
and establishing a three-dimensional rectangular coordinate system, taking the center of the bottom of the rocket body as an origin, taking the flight direction of the rocket as the positive direction of the z axis, and taking any two mutually perpendicular straight lines in a plane perpendicular to the z axis as the y axis and the z axis. In a three-dimensional rectangular coordinate system, the mass, momentum, and energy equations can be expressed as follows:
Figure BDA0002331902160000031
wherein U is a liquid rocket fuel gas flow variable; F. g, H flux vectors in x, y and z directions of gas flow of liquid rocket, Fv、Gv、HvFlux vectors of the liquid rocket fuel gas sticking in the x, y and z directions are respectively;
3.2, establishing a turbulence model of the liquid rocket gas impact flow guide device by adopting a Realizable k-epsilon two-equation model:
the turbulence model of the Realizable k-epsilon two-equation has accurate prediction on the divergence ratio of the flat jet flow and higher accuracy on the calculation of boundary layer flow, flow separation and secondary flow, and the turbulence energy k equation and the turbulence dissipation rate epsilon equation are
Figure BDA0002331902160000041
Figure BDA0002331902160000042
Wherein G iskFor the generation term of the turbulent kinetic energy k due to the mean velocity gradient, mutFor turbulent viscosity, σkAnd σεRespectively showing Plantt number of turbulence energy k and dissipation rate epsilon of turbulence, α and β are index symbols in tensor, the value ranges are (1, 2 and 3), C1And C2Is a constant coefficient of equation, C1＝1.44，C2＝1.9。
Step 4, establishing a liquid rocket fuel gas chemical dynamics model: carrying out numerical simulation on the re-combustion reaction of the rocket gas based on the finite rate chemical dynamics model, and establishing a rocket gas chemical dynamics model;
4.1, quantification based on mass action, for any one chemical reaction equation:
Figure BDA0002331902160000043
wherein i is a liquid rocket fuel gas component, r is a chemical reaction step, v'irIs the reaction stoichiometry, v ″, of component i in reaction step rirIs the formation of the stoichiometric coefficient, M, of component i in the reaction step riIs the relative molecular mass of component i;
4.2 production ratio ω of component i in reaction step rirComprises the following steps:
Figure BDA0002331902160000044
in the formula (I), the compound is shown in the specification,
Figure BDA0002331902160000045
is the molecular weight of component i, CiIs the molar concentration of component i, Kfr、KbrRespectively a forward reaction constant and a reverse reaction constant of the reaction step r;
the forward reaction constant can be expressed as:
Figure BDA0002331902160000046
in the formula, ArIs a pre-exponential factor of reaction step r, NTAnd NPIndex E of temperature and pressure, respectivelyArIs the activation energy of reaction step R, R0Is a universal gas constant, PatmIs a standard atmospheric pressure;
the inverse reaction constant can be expressed as:
Figure BDA0002331902160000051
wherein the Gibbs free energy change amount of the reaction step r
Figure BDA0002331902160000052
Can be expressed as:
Figure BDA0002331902160000053
in the formula, giGibbs free energy as component i;
step 5, carrying out parallel calculation on the gas impact flow field of the liquid carrier rocket, and outputting a Mach number cloud picture, a temperature cloud picture, CO and CO2、H2、H2And O mole fraction cloud.
5.1 selection of CO-H2An Air chemical reaction model, inputting the pre-exponential factor, the temperature index and the activation energy of each step of reaction into a numerical simulation;
5.2, inputting the total temperature, the total pressure and the mass fraction of each component of the jet pipe of the liquid carrier rocket engine;
5.3 parallel computing output Mach number cloud chart and temperatureCloud picture, CO2、H2、H2And O mole fraction cloud.
Examples
The numerical simulation method for the liquid rocket afterburning reaction calculation comprises the following steps according to the steps in the specific embodiment:
step 1, establishing a safe three-dimensional geometric model of a carrier rocket impact diversion trench;
in connection with fig. 2, the three-dimensional geometric model requires the following parameters: the curvature and the radius of a carrier rocket warhead-1, the height and the radius of a rocket body-2, the overall length of a Laval nozzle-3, the radius of a nozzle inlet-4, the radius of a nozzle throat-5, the radius of a nozzle outlet-6, the length and the width of a flow guide device inlet-7, the length and the width of a flow guide device outlet-8 and the curvature of a flow guide device bottom surface-9, and 1:1 modeling is carried out after all parameters are determined.
Step 2, carrying out structural grid division on the liquid rocket impact flow guide device by adopting a grid block scheme;
combining with the figure 3, carrying out grid partitioning on the three-dimensional model of the liquid rocket impact flow guide device, wherein the total number of flow field grids of the liquid carrier rocket impact flow guide device is 989 ten thousands;
step 3, establishing a liquid rocket fuel gas impact model: establishing a rocket gas impact model based on a three-dimensional compressible Reynolds average N-S equation and a readable k-epsilon two-stroke turbulence model;
step 4, establishing a liquid rocket fuel gas chemical dynamics model: carrying out numerical simulation on the re-combustion reaction of the rocket gas based on the finite rate chemical dynamics model, and establishing a rocket gas chemical dynamics model;
step 5, carrying out parallel calculation on the gas impact flow field of the liquid carrier rocket, and outputting a Mach number cloud picture, a temperature cloud picture, CO and CO2、H2、H2And O mole fraction cloud.
The following parameters were entered:
each item of chemical reaction indicates a pre-factor ArTemperature index NTActivation energy EArAs shown in the following table
Serial number Chemical reaction Ar NT EAr Serial number Chemical reaction Ar NT EAr
R1 CO+OH＝CO2+H 2.25E+10 1.55 -3.34 R10 N2+O＝N+NO 1.81E+14 0 318
R2 OH+H2＝H2O+H 1.8E+12 1.3 15.3 R11 N+O2＝O+NO 2.69E-10 1 27.19
R3 H+O2＝OH+O 9.75E+13 0 62.11 R12 N+OH＝H+NO 2.83E-10 0 0
R4 H2+O＝OH+H 1.84E+11 2.7 26.19 R13 H+O+M＝HO+M 2.62E-08 -1 0
R5 OH+OH＝H2O+O 6.14E+11 1.4 -1.66 R14 N+O+M＝NO+M 3.28E-09 0 -1.29
R6 CO+O+M＝CO2+M 8.79E-11 0 -18.96 R15 NO+NO＝O2+N2 3.01E+12 0.5 254
R7 H+H+M＝H2+M 5.32E-09 -0.6 0 R16 O+NO2＝NO+O2 3.92E+12 0 -1
R8 H+OH+M＝H2+M 8.91E-07 -1.2 2.58 R17 NO2+H＝NO+OH 8.85E+13 0 0
R9 O+O+M＝O2+M 3.13E-11 0 -7.48 R18 NO+O+M＝NO2+M 6.20E-07 -2. 6.49
Ambient temperature: 300k, and (c) respectively; ambient pressure: 101325 Pa; total pressure of the engine: 1.8*107Pa; total temperature of the engine: 3400 k.
The numerical simulation method for liquid rocket afterburning reaction calculation utilizes 30 threads of a V4-2697A calculation platform to calculate 989 ten thousand grids, the total calculation time is 46 hours, and Mach number cloud pictures, temperature cloud pictures, CO and CO are output2、H2、H2An O molar fraction cloud chart is shown in the following figure, wherein figure 4 shows that the tail flame of the liquid carrier rocket impacts a Mach number re-ignition flow field, plumes form two Mach waves before impacting a diversion trench, the third Mach wave appears on a diversion surface, figure 5 shows that the tail flame of the liquid carrier rocket impacts a temperature re-ignition flow field, a high-temperature region mainly appears in an accumulation region where fuel gas directly impacts on the diversion surface and a mixing region where the fuel gas is contacted with air, and figures 6, 7, 8 and 9 respectively show that the tail flame of the liquid carrier rocket impacts CO and CO2、H2、H2O mole fraction post-combustion flow field, CO and H compared to a case where post-combustion is not taken into consideration2A significant reduction in CO2And H2O is significantly increased, and FIG. 10 is a numerical value of the present inventionThe results of the calculation by the method are compared with the tests (Cao D, He G, Qin F, Dan M., "Local super and sub communication model transmission in a super jet film," Proceedings of the communication institute, Vol.37,2018, pp.3723-3731.), and the results are well matched with the tests, thereby showing that the invention has the characteristic of high precision. The numerical simulation method can improve the calculation precision and reduce the calculation cost, and the calculation result can provide a theoretical basis for the thermal protection design of the liquid carrier rocket in the launching stage.

Claims (5)

1. The numerical simulation method for the liquid rocket afterburning reaction calculation is characterized by comprising the following steps of:
step 1, establishing a three-dimensional geometric model of a gas impact flow guide device in a liquid carrier rocket launching stage;
step 2, carrying out structural grid division on the liquid rocket impact flow guide device by adopting a grid block scheme;
step 3, establishing a liquid rocket fuel gas impact model: establishing a rocket gas impact model based on a three-dimensional compressible Reynolds average N-S equation and a Realizblek-epsilon two-stroke turbulence model;
step 4, establishing a liquid rocket fuel gas chemical dynamics model: carrying out numerical simulation on the re-combustion reaction of the rocket gas based on the finite rate chemical dynamics model, and establishing a rocket gas chemical dynamics model;
step 5, carrying out parallel calculation on the gas impact flow field of the liquid carrier rocket, and outputting a Mach number cloud picture, a temperature cloud picture, CO and CO2、H2、H2And O mole fraction cloud.
2. The numerical simulation method for liquid rocket afterburning reaction calculation according to claim 1, wherein step 1 of establishing a three-dimensional geometric model of the gas impact diversion device at the launch stage of the liquid carrier rocket specifically comprises the following steps:
1.1, drawing a three-dimensional model according to a liquid carrier rocket 1:1, wherein the following parameters are required: the carrier rocket comprises a carrier rocket warhead curvature, a rocket body height and radius, a spray pipe length, a spray pipe inlet radius, a spray pipe throat radius and a spray pipe outlet radius;
1.2, drawing a three-dimensional model according to the flow guide device 1:1, wherein the following parameters are required: length and width of the inlet of the flow guide device, length and width of the outlet of the flow guide device and curvature of the bottom surface of the flow guide device.
3. The numerical simulation method for liquid rocket afterburning reaction calculation according to claim 1, wherein step 2 adopts a grid blocking scheme to perform structural grid division on the liquid rocket impact diversion device, and specifically comprises the following steps:
2.1, carrying out grid partitioning on the three-dimensional model of the liquid rocket impact flow guide device, and dividing the whole calculation area into a carrier rocket sub-area, a transition sub-area of a rocket engine and the flow guide device and a flow guide device sub-area;
and 2.2, encrypting the sub-domain of the transition of the rocket engine and the flow guiding device and the sub-domain grids of the flow guiding device, and gradually transitioning the sub-domain grids of the carrier rocket from dense to sparse.
4. The numerical simulation method for liquid rocket afterburning reaction calculation according to claim 1, wherein step 3 is to establish a liquid rocket fuel gas impact model, and the specific steps are as follows:
3.1, establishing a three-dimensional compressible Reynolds average N-S equation of the liquid rocket fuel gas:
in a three-dimensional rectangular coordinate system, the mass, momentum, and energy equations can be expressed as follows:
Figure FDA0002331902150000011
wherein U is a liquid rocket fuel gas flow variable; F. g, H flux vectors in x, y and z directions of gas flow of liquid rocket, Fv、Gv、HvFlux vectors of the liquid rocket fuel gas sticking in the x, y and z directions are respectively;
3.2, establishing a turbulence model of the liquid rocket gas impact flow guide device by adopting a Realizable k-epsilon two-equation model:
the equation of the turbulence kinetic energy k and the equation of the turbulence dissipation rate epsilon are
Figure FDA0002331902150000021
Figure FDA0002331902150000022
Wherein G iskFor the generation term of the turbulent kinetic energy k due to the mean velocity gradient, mutFor turbulent viscosity, σkAnd σεPrandtl numbers of turbulence energy k and turbulence dissipation rate epsilon respectively, α and β are index symbols in tensor, the value range is (1, 2 and 3), and C1And C2Are the equation constant coefficients.
5. The numerical simulation method for liquid rocket afterburning reaction calculation according to claim 1, wherein step 4 is to establish a liquid rocket fuel gas chemical dynamics model, and specifically comprises the following steps:
4.1, quantification based on mass action, for any one chemical reaction equation:
Figure FDA0002331902150000023
wherein i is a liquid rocket fuel gas component, r is a chemical reaction step, v'irIs the reaction stoichiometry, v ″, of component i in reaction step rirIs the formation of the stoichiometric coefficient, M, of component i in the reaction step riIs the relative molecular mass of component i;
4.2 production ratio ω of component i in reaction step rirComprises the following steps:
Figure FDA0002331902150000024
in the formula (I), the compound is shown in the specification,
Figure FDA0002331902150000025
of component iMolecular weight, CiIs the molar concentration of component i, Kfr、KbrRespectively, the forward reaction constant and the reverse reaction constant of the reaction step r.
CN201911342880.XA 2019-12-23 2019-12-23 Numerical simulation method for liquid rocket afterburning reaction calculation Withdrawn CN111210503A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201911342880.XA CN111210503A (en) 2019-12-23 2019-12-23 Numerical simulation method for liquid rocket afterburning reaction calculation

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201911342880.XA CN111210503A (en) 2019-12-23 2019-12-23 Numerical simulation method for liquid rocket afterburning reaction calculation

Publications (1)

Publication Number Publication Date
CN111210503A true CN111210503A (en) 2020-05-29

Family

ID=70788179

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201911342880.XA Withdrawn CN111210503A (en) 2019-12-23 2019-12-23 Numerical simulation method for liquid rocket afterburning reaction calculation

Country Status (1)

Country Link
CN (1) CN111210503A (en)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113722830A (en) * 2021-09-03 2021-11-30 华南理工大学 Solid rocket engine C/C composite material nozzle ablation behavior modeling simulation method
CN115081108A (en) * 2022-05-23 2022-09-20 中国人民解放军国防科技大学 Full-flow-field numerical simulation method and system for hypersonic cruise aircraft
CN116629028A (en) * 2023-07-19 2023-08-22 东方空间技术(山东)有限公司 Method and device for determining parameters of flow guide groove of petal-shaped launching pad
CN117216900A (en) * 2023-09-11 2023-12-12 中国人民解放军国防科技大学 IRC method-based liquid oxygen kerosene engine combustion chamber modeling method
CN117711511A (en) * 2024-02-04 2024-03-15 中国空气动力研究与发展中心计算空气动力研究所 Mars gas finite rate chemical reaction model construction method and model data system

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113722830A (en) * 2021-09-03 2021-11-30 华南理工大学 Solid rocket engine C/C composite material nozzle ablation behavior modeling simulation method
CN113722830B (en) * 2021-09-03 2023-04-11 华南理工大学 Solid rocket engine C/C composite material nozzle ablation behavior modeling simulation method
CN115081108A (en) * 2022-05-23 2022-09-20 中国人民解放军国防科技大学 Full-flow-field numerical simulation method and system for hypersonic cruise aircraft
CN116629028A (en) * 2023-07-19 2023-08-22 东方空间技术(山东)有限公司 Method and device for determining parameters of flow guide groove of petal-shaped launching pad
CN116629028B (en) * 2023-07-19 2023-09-22 东方空间技术(山东)有限公司 Method and device for determining parameters of flow guide groove of petal-shaped launching pad
CN117216900A (en) * 2023-09-11 2023-12-12 中国人民解放军国防科技大学 IRC method-based liquid oxygen kerosene engine combustion chamber modeling method
CN117216900B (en) * 2023-09-11 2024-04-02 中国人民解放军国防科技大学 IRC method-based liquid oxygen kerosene engine combustion chamber modeling method
CN117711511A (en) * 2024-02-04 2024-03-15 中国空气动力研究与发展中心计算空气动力研究所 Mars gas finite rate chemical reaction model construction method and model data system
CN117711511B (en) * 2024-02-04 2024-05-17 中国空气动力研究与发展中心计算空气动力研究所 Mars gas finite rate chemical reaction model construction method and model data system

Similar Documents

Publication Publication Date Title
CN111210503A (en) Numerical simulation method for liquid rocket afterburning reaction calculation
Wang et al. Unified Navier-Stokes flowfield and performance analysis of liquid rocket engines
Chan et al. Numerically simulated comparative performance of a scramjet and shcramjet at Mach 11
CN109359325B (en) Simulation method for multi-nozzle rocket flow field and convection/radiation coupling heat exchange
Smart Scramjet inlets
Dudebout et al. Numerical simulation of hypersonic shock-induced combustion ramjets
Dalle et al. Hypersonic vehicle flight dynamics with coupled aerodynamic and reduced-order propulsive models
Hassantabar et al. Investigating the effect of engine speed and flight altitude on the performance of throttle body injection (TBI) system of a two-stroke air-powered engine
Siddiqui et al. Design and analysis on scramjet engine inlet
Yungster et al. Analysis of a new rocket-based combined-cycle engine concept at low speed
Mogavero et al. Hybrid Propulsion Parametric and Modular Model: a novel engine analysis tool conceived for design optimization
Haws et al. Computational investigation of a method to compress air fluidically in supersonic inlets
Kodera et al. Numerical analysis of transient phenomena to ramjet mode in a RBCC combustor
Ispir et al. Analysis of a combined cycle propulsion system for STRATOFLY hypersonic vehicle over an extended trajectory
Shyji et al. Numerical studies on thrust augmentation in high area ratio rocket nozzles by Secondary injection
Molchanov An effective numerical method for simulating chemically and thermally nonequilibrium gas flows
Savino et al. Numerical analysis of supersonic combustion ramjet with upstream fuel injection
Luo et al. Thermodynamic performance modeling, optimization and numerical simulation of RBCC ejector mode
Chapman et al. Simulations of non-reacting ethylene/air supersonic flow in a cavity flame holder at Mach 2 and Mach 3
Kodera et al. Numerical analysis of scramjet mode operation of a RBCC engine
Pao et al. Establishing Approaches to Modeling the Ares IX and Ares I Roll Control System with Free-Stream Interaction
McKamey et al. A one-dimensional engineering model for the evaluation of rocket-based combined cycle engine performance
Ferguson et al. CFD analysis of an inlet-isolator combination for dual mode scramjet applications
Couture et al. Comparison of scramjet and shcramjet propulsion for an hypersonic waverider configuration
DellaFera Optimization of hypersonic airbreathing propulsion systems through mixed analysis methods

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
WW01 Invention patent application withdrawn after publication

Application publication date: 20200529

WW01 Invention patent application withdrawn after publication