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1.
This study is concerned with the experimental and theoretical investigation of the combustion instabilities in a premixed swirl combustor. It is focused on the effects of the swirl mixing distance on the intrinsic thermoacoustic mode. The swirler as an origin of the swirling flow is also the source of the flow disturbance, which has effects on the flame response. The location of the swirler is varied in the experiment to study the effect on combustion instabilities and flame transfer functions. ... 相似文献
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Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions 总被引:2,自引:0,他引:2
To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions.The overall model is benchmarked under various impingement angles, jet momentum and offcenter ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines. 相似文献
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Numerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant
Multi-phase flowfield simulation has been performed on solid rocket motor and effect of multi-phases on the performance prediction of the solid rocket motor(SRM) is investigation.During the combustion of aluminized propellant,the aluminum particles in the propellant melt and formliquid aluminum at the burning propellant surface.So the flow within the rocket motor is multi phase or two phase because it contains droplets and smoke particles of Al2O3.Flowsi mulations have been performed on a large scale motor,to observe the effect of the flowfield onthe chamber and nozzle as well.Uniform particles diameters and Rosin-Rammler diameter distribution method that is based on the assumption that an exponential relationship exists betweenthe droplet diameter,dand mass fraction of droplets with diameter greater thandhave been used for the si mulation of different distribution of Al2O3 droplets present in SRM.Particles sizes in the range of 1-100μm are used,as being the most common droplets.In this approachthe complete range of particle sizes is dividedinto a set of discrete size ranges,eachto be defined by single streamthat is part of the group.Roe scheme-flux differencing splitting based on approxi mate Riemann problem has been used to si mulate the effects of the multi-phase flowfeild.This is second order upwind scheme in which flux differencing splitting method is employed.To cater for the turbulence effect,Spalart-All maras model has been used.The results obtained show the great sensitivity of this diameters distribution and particles concentrations to the SRMflowdynamics,primarily at the motor chamber and nozzle exit.The results are shown with various sizes of the particles concentrations and geometrical configurations including models for SRM and nozzle.The analysis also provides effect of multi-phase on performance prediction of solid rocket motor. 相似文献
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This paper presents how the combustion performance of nano-sized aluminum(nAl)powder in carbon dioxide are affected by silica. The ignition and combustion performance of nAl powder with silica addition were studied by a high-temperature tube furnace. An s-type thermocouple and a high-speed motion acquisition instrument were performed to evaluate the ignition temperature, maximum combustion temperature, maximum change of rate of temperature, and combustion propagation speed. The combustion effici... 相似文献
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Experimental investigation on aero-heating of rudder shaft within laminar/turbulent hypersonic boundary layers 总被引:1,自引:0,他引:1
The aero-heating of the rudder shaft region of a hypersonic vehicle is very harsh, as the peak heat flux in this region can be even higher than that at the stagnation point. Therefore, studying the aero-heating of the rudder shaft is of great significance for designing the thermal protection system of the hypersonic vehicle. In the wind tunnel test of the aero-heating effect, we find that with the increase of the angle of attack of the lifting body model, the increasement of the heat flux of the rudder shaft is larger under laminar flow conditions than that under turbulent flow conditions. To understand this, we design a wind tunnel experiment to study the effect of laminar/turbulent hypersonic boundary layers on the heat flux of the rudder shaft under the same wind tunnel freestream conditions. The experiment is carried out in the ?2 m shock tunnel(FD-14 A) affiliated to the China Aerodynamics Research and Development Center(CARDC). The laminar boundary layer on the model is triggered to a turbulent one by using vortex generators, which are 2 mm-high diamonds. The aero-heating of the rudder shaft(with the rudder) and the protuberance(without the rudder) are studied in both hypersonic laminar and turbulent boundary layers under the same freestream condition. The nominal Mach numbers are 10 and 12, and the unit Reynolds numbers are2.4 × 10~6 m~(-1) and 2.1 × 10~6 m-1. The angle of attack of the model is 20°, and the deflection angle of the rudder and the protuberance is 10°. The heat flux on the model surface is measured by thin film heat flux sensors, and the heat flux distribution along the center line of the lifting body model suggests that forced transition is achieved in the upstream of the rudder. The test results of the rudder shaft and the protuberance show that the heat flux of the rudder shaft is lower in the turbulent flow than that in the laminar flow, but the heat flux of the protuberance is the other way around,i.e., lower in the laminar flow than in the turbulent flow. The wind tunnel test results is also validated by numerical simulations. Our analysis suggests that this phenomenon is due to the difference of boundary layer velocities caused by different thickness of boundary layer between laminar and turbulent flows, as well as the restricted flow within the rudder gap. When the turbulent boundary layer is more than three times thicker than that of the laminar boundary layer, the heat flux of the rudder shaft under the laminar flow condition is higher than that under the turbulent flow condition. Discovery of this phenomenon has great importance for guiding the design of the thermal protection system for the rudder shaft of hypersonic vehicles. 相似文献
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In order to accurately predict the heat and mass transfer behaviors and analyze key factors affecting pressurization process in the hydrogen tank, a comprehensive 2 D axial symmetry Volume-Of-Fluid(VOF) model is established by Computational Fluid Dynamics(CFD) method.The effects of phase change, turbulence and mass diffusion are included in the model and relationships between physical properties and temperature are also comprehensively considered. The phase change model is based on Hertz-Knudsen equation and the mass transfer time relaxation factor is determined by the NASA's experimental data. The mass diffusion model is included in gaseous helium pressurizing. The key factors including the inlet temperature, inlet mass flow rate, injector types and pressurizing gas kinds are quantitatively analyzed. Compared with the experiment, the simulation results show that the deviation of pressurizing gas mass consumption, condensing mass and ullage temperature are 3.0%, 7.5% and 4.0% respectively. The temperature stratification is existed along the axial direction in the surface liquid region and the ullage region, and the bulk liquid is in subcooled state during pressurizing. The location of phase change mainly appears near the vapor–liquid interface, and the mass transfer expressing as condensation or vaporization is mainly determined by the heat convection and molecular concentration near the vapor–liquid interface.The key factors show that increasing the inlet temperature and inlet mass flow rate could shorten the pressurizing time interval and save the pressurizing gas mass. The proportion of the total energy addition of the tank absorbed by the ullage region, the liquid region and the tank wall respectively is greatly influenced by the injector types and more heat transferred into the ullage would result in a faster pressure rising rate. Gaseous hydrogen pressurization has a higher efficiency than gaseous helium pressurization. The simulation results presented in this paper can be used as a reference for design optimization of the pressurization systems of cryogenic liquid launch vehicles so as to save the mass of pressurizing gases and shorten the pressurizing time interval. 相似文献
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A new water-cooling Gardon-type heat power measuring apparatus is designed to meet the need of heat power source management and distribution. The steady state measurement mathematic model of the apparatus is built up in theory and the system amplification coefficient is defined as the ratio of the heat power to the temperature difference of the device, with which the value of the measured source power can be calculated easily with the corresponding temperature difference. In order to obtain an optimal heat power measuring system, the coefficients that can influence the relationship between the amplification coefficient, the temperature difference, and the heat power are analyzed. On the basis of these analyses, a set of experimental device is constructed and a number of experiments are carried on. Compared with the input heat power sample data, the error of the experimental measuring results is less than ±2%, and the experimental measuring values are in good agreement with the calculated theoretical ones. The heat power measuring apparatus can be applied in heat flux or heat power measurement in other fields due to its simple structure and high accuracy. 相似文献
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《中国航空学报》1990,(4)
The two-dimensional turbulent evaporating gas-droplet two-phase flows in an afterburner diffusor of turbo-fan jet engines are simulated by the k-ε turbulence model and the particle trajectory model. Comparison of predicted gas velocity and temperature distributions with experimental results for the cases without liquid spray shows pretty good agreement. Gas-droplet two-phase flow predictions give plausible droplet trajectories, fuel-vapor concentration distribution, gas-phase velocity and temperature field in presence of liquid droplets. One run of computation with this method is made for a particular afterburner. The results indicate that the location of the atomizers is not favorable to flame stabilization and combustion efficiency. The proposed numerical modeling can also be adopted for optimization design and performance evaluation of afterburner combustors of turbo-fan jet engines. 相似文献
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The infrared(IR) irradiance signature from rocket motor exhaust plumes is closely related to motor type,propellant composition,burn time,rocket geometry,chamber parameters and flight conditions.In this paper,an infrared signature analysis tool(IRSAT) was developed to understand the spectral characteristics of exhaust plumes in detail.Through a finite volume technique,flow field properties were obtained through the solution of axisymmetric Navier-Stokes equations with the Reynolds-averaged approach.A refined 13-species,30-reaction chemistry scheme was used for combustion effects and a k-e-Rtturbulence model for entrainment effects.Using flowfield properties as input data,the spectrum was integrated with a line of sight(LOS) method based on a single line group(SLG) model with Curtis-Godson approximation.The model correctly predicted spectral distribution in the wavelengths of 1.50–5.50 lm and had good agreement for its location with imaging spectrometer data.The IRSAT was then applied to discuss the effects of three operating conditions on IR signatures:(a) afterburning;(b) chamber pressure from ignition to cutoff;and(c) minor changes in the ratio of hydroxyl-terminated polybutadiene(HTPB) binder to ammonium perchlorate(AP) oxidizer in propellant.Results show that afterburning effects can increase the size and shape of radiance images with enhancement of radiation intensity up to 40%.Also,the total IR irradiance in different bands can be characterized by a non-dimensional chamber pressure trace in which the maximum discrepancy is less than 13% during ignition and engine cutoff.An increase of chamber pressure can lead to more distinct diamonds,whose distance intervals are extended,and the position of the first diamond moving backwards.In addition,an increase in HTPB/AP causes a significant jump in spectral intensity.The incremental rates of radiance intensity integrated in each band are linear with the increase of HTPB,and the growth rates of radiance intensities in some bands reach up to 50% as HTPB weight increases by 3%. 相似文献
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The aim of the present paper is to study the numerical solutions of the steady MHD two dimensional stagnation point flow of an incompressible nano fluid towards a stretching cylinder.The effects of radiation and convective boundary condition are also taken into account.The model used for the nanofluid incorporates the effects of Brownian motion and thermophoresis.The resulting nonlinear momentum,energy and nano particle equations are simplifed using similarity transformations.Numerical solutions have been obtained for the velocity,temperature and nanoparticle fraction profles.The influence of physical parameters on the velocity,temperature,nanoparticle fraction,rates of heat transfer and nanoparticle fraction are shown graphically. 相似文献
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《中国航空学报》2023,36(5):66-77
The combustion of aluminum particles in solid rocket motor plays an important role in energy release of propellants. However, due to the limited residence time, aluminum particles may not be burned completely, thus hindering the improvement of specific impulse. This study aims to explore the characteristics of aluminum combustion efficiency and its influencing factors by experiments and numerical simulations, providing a guideline for engine performance improvement. As an input of simulation, the initial agglomerate size was measured by a high pressure system. Meanwhile, the size distribution of the particles in plume was measured by ground firing test to validate the numerical model. Then, a two-phase flow model coupling combustion of micro aluminum particle was developed, by which the detailed effects of particle size, detaching position and nozzle convergent section structure on aluminum combustion efficiency were explored. The results suggest that the average combustion temperature in the chamber drops with increasing initial particle size, while the maximum temperature increases slightly. In the tested motors, the aluminum particle burns completely as its diameter is smaller than 50 μm, and beyond 50 μm the combustion efficiency decreases obviously with the increase of initial size. As the diameter approaches to 75 μm, the combustion efficiency becomes more sensitive to particle size. The combustion efficiency of aluminum particle escaping from end-burning surfaces is significantly higher than that from internal burning surface, where the particle combustion efficiency decreases during approaching the convergent section. Furthermore, the combustion efficiency decreases slightly with increasing nozzle convergent section angle. And theoretically it is feasible to improve combustion efficiency of aluminum particles by designing the convergent profile of nozzle. 相似文献
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《Aerospace Science and Technology》2007,11(1):33-38
Two different approaches are used in this work to reduce the burning times of aluminium particles with the ultimate goal to improve the performances of solid propellants. One method is to coat the micro-sized particles by nickel, and the second is to decrease the particle sizes to nano-metric scales.A thin coating of Ni on the surface of Al particles can prevent their agglomeration and at the same time facilitates their ignition, thus increasing the efficiency of aluminized propellants. In this work, ignition and burning of single Ni-coated Al particles are investigated using an electrodynamic levitation setup and laser heating of the particles. The levitation experiments are used to measure the particle ignition delay time and burning time at different Ni contents in the particles.Decreasing the size of Al particles increases their specific surface, and hence decreases the burning time of the same mass of particles. In this investigation, a cloud of Al nano-particles formed in a combustion tube is ignited by an electric spark. The cloud experiments are used to measure comparative flame front propagation velocities for different Al particle sizes with and without organic coating.The results and their analysis show that both methods reduce the Al burning time. Ni coating reduces significantly the ignition time of micro-sized Al particles and hence the total burning time compared to non-coated particles. Nano-sized particle clouds burn faster than micro-sized Al particle clouds. 相似文献
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《中国航空学报》2017,(6)
Particle size and oxygen content are two of the key factors that affect the ignition and combustion properties of aluminum particles. In this study, a laser ignition experimental system and flame test system were built to analyze the ignition and combustion characteristics and the flame morphology of aluminum particles. A thermobalance system was used to analyze the thermal oxidation characteristics. In addition, the microstructure of aluminum was analyzed by scanning electron microscopy. It was found that the oxidized products were some of the gas phase products agglomerated. Smaller particle size samples showed better combustion characteristics. The combustion intensity, self-sustaining combustion time and the burn-off rate showed a rising trend with the decrease in the particle size. Increasing the oxygen content in the atmosphere could improve the ignition and combustion characteristics of the samples. Four distinct stages were observed in the process of ignition and combustion. Small particle size samples had a larger flame height and luminance, and the self-sustaining combustion time was much longer.Three distinct stages were observed during the thermal oxidation process. The degree of oxidation for small-sized samples was significantly higher than that for the larger particle size samples.Moreover, it was observed that the higher the oxygen content, the higher the degree of oxidation was. 相似文献
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