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Transverse slot injection scheme is very important for the mixing process between the air and the fuel in supersonic flows. The effect of the turbulence model and slot width on the transverse slot injection flow field has been investigated numerically based on the grid independency analysis, and the predicted results have been compared with the experimental data available in the open literature. The obtained results show that the grid scale makes only a slight difference to the wall pressure profiles for all jet-to-crossflow pressure ratios employed in this study, and the wall pressure profile with low jet-to-crossflow pressure ratio is predicted accurately by the RNG k–ε turbulence model, the SST k–ω turbulence model for the flow field with high jet-to-crossflow pressure ratio. High jet-to-crossflow pressure ratio can increase the jet penetration depth in supersonic flows, and the gradient of the length of the upstream separation region is larger than that of the height of the Mach surface. At the same time, when the jet-to-crossflow pressure ratio is maintained constant, the jet penetration depth increases with the increase of the slot width. 相似文献
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为了提高超燃冲压发动机隔离段耐反压能力以及缩短其长度,在前期后掠斜楔数值研究基础上,设计了一种带后掠斜楔的隔离段,斜楔放置在隔离段进口的下壁面上,距隔离段进口长度约15%处,在非对称的隔离段进口来流速度为1.98马赫数的条件下完成吹风实验.实验结果表明,隔离段添加后掠斜楔后的最大承受反压从来流静压的3.55倍上升到3.90倍,提高了9.89%.相同反压下,带后掠斜楔的短隔离段长度缩短了15%.相同长度的带后掠斜楔的隔离段出口平均总压恢复系数由基准隔离段的0.694上升到0.710,提高了2.3%. 相似文献
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The dynamics of a two dimensional plane jet injected at the base of a step, parallel to the wall, in backward facing step flow geometry is numerically studied. The objective of this work is to gain insight into the dynamics of the igniter flow field in solid fuel ramjet motors. Solid fuel ramjets operate by ingestion of air and subsequent combustion with a solid fuel grain such as polyethylene. The system of governing equations is solved with a finite volume approach using a structured grid in which the AUSM+ scheme is used to calculate the convective fluxes. The Spalart and Allmaras turbulence model is used in these simulations. Experimental data have been used to validate the flow solver and turbulence model simulation results. The comparison of the numerical results and experimental data has validated the use of the adopted turbulence model for the study of this type of problem. A special attention is paid to the igniter jet exit location. It is shown that the wall jet igniter, issuing from the base of the step, drastically changes the structure of recirculating region of backward facing step flow and produces large and damaging shear stress on the fuel surface. Moving the igniter jet exit location to the top of the backward facing step changes the flow field favorably, by reducing the fuel surface shear stress by an order of magnitude and maintaining the recirculating region behind the step, which can provide proper residence time for the fuel–air mixture chemical reactions. 相似文献
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The mixing process between the injectant and the supersonic crossflow is one of the important issues for the design of the scramjet engine, and the efficiency mixing has a great impact on the improvement of the combustion efficiency. A hovering vortex is formed between the separation region and the barrel shock wave, and this may be induced by the large negative density gradient. The separation region provides a good mixing area for the injectant and the subsonic boundary layer. In the current study, the transverse injection flow field with a freestream Mach number of 3.5 has been optimized by the non-dominated sorting genetic algorithm (NSGA II) coupled with the Kriging surrogate model; and the variance analysis method and the extreme difference analysis method have been employed to evaluate the values of the objective functions. The obtained results show that the jet-to-crossflow pressure ratio is the most important design variable for the transverse injection flow field, and the injectant molecular weight and the slot width should be considered for the mixing process between the injectant and the supersonic crossflow. There exists an optimal penetration height for the mixing efficiency, and its value is about 14.3 mm in the range considered in the current study. The larger penetration height provides a larger total pressure loss, and there must be a tradeoff between these two objection functions. In addition, this study demonstrates that the multi-objective design optimization method with the data mining technique can be used efficiently to explore the relationship between the design variables and the objective functions. 相似文献
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The single expansion ramp nozzle is an essential component for hypersonic vehicles to improve their internal/external integral level and produce most of the thrust force. The two-dimensional coupled implicit Reynolds Averaged Navier–Stokes (RANS) equations and the two-equation RNG k–ε turbulent model have been employed to numerically simulate the flow field in a single expansion ramp nozzle, and the interactions between the parametric parameters and the objective functions, namely the thrust force and the lift force, have been investigated by using the data mining technique coupled with a design of the experiment. The obtained results show that the physical model has a good two-dimensional structure, and the numerical results show very good agreement with the available experimental data. At the same time, the grid scale has only a slight impact on the pressure distribution. The influences of the horizontal length of the inner nozzle, the external expansion ramp and the internal cowl expansion on the thrust force performance are substantial, as are the effects of the internal cowl expansion and the external expansion ramp on the lift force performance. Further, optimized configurations of the single expansion ramp nozzle are obtained by using single- and multi-objective design optimization methods coupled with the Kriging surrogate model, and the optimized performances show very good agreement with the numerical predictions. The discrepancies between the optimized performances and the numerical predictions are less than 0.05%, and the method proposed in this paper is efficient in designing and optimizing the nozzle configuration. 相似文献
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A CFD study on drag reduction in supersonic flow with opposing jet has been conducted. Flowfield parameters, reattachment point position and surface pressure distributions are obtained and validated with experiments. From the analysis on the physical mechanism of drag reduction, it shows the phenomenon that, when the opposing jet blows, the high pressure region is located between the bow shock wave and the Mach disk, which makes the nose region much lower pressure. As the pressure ratio increases, the high pressure region is gradually pushed away from the surface. Larger the total pressure ratio is, the lower of the drag coefficient is. To study the effect of the intensity of opposing jet more reasonably, a new parameter RPA has been introduced by combining the flux and the total pressure ratio. The study shows that the same shock wave position and drag coefficient can be obtained with the same RPA with different fluxes and the total pressures, which means the new parameter could stand for the intensity of opposing jet and could be used to analyze the influence of opposing jet on flow field and aerodynamic force. 相似文献
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固体火箭燃气超燃冲压发动机具有高比冲、结构简单、流量易调节等优点,然而在超音速空气流的补燃室中,如何让燃料更好地与空气掺混,增加颗粒停留时间,在较短时间内释放出更多的燃烧焓成为目前研究的重点。采用Realiazble k-ε湍流模型,单步涡团耗散模型,在King的硼颗粒点火燃烧模型的基础上考虑了硼颗粒在高速气流当中的气动剥离效应,利用龙格-库塔算法迭代计算硼颗粒点火燃烧过程,对燃气进气方向与轴向夹角从45°~180°的10种进气方式下的补燃室进行了三维两相燃烧流动计算,分析了各种进气角下的燃气燃烧效率、硼颗粒燃烧效率以及总燃烧效率。结果表明:当一次燃气喷射角度与轴向夹角逐渐增加时,燃气与颗粒燃烧效率逐渐增加,并在180°时燃烧效率和比冲为最高。 相似文献
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高速气流场燃油雾化液滴分布数值研究 总被引:2,自引:0,他引:2
针对亚燃冲压发动机燃烧室内部流动特点,结合二元稳定器试验台高速气流场燃油雾化特性试验,建立试验件三维模型并对其喷雾两相流动进行数值模拟。主要研究了来流马赫数以及喷嘴条件变化时燃油雾化液滴与油气比的分布。分析认为,来流马赫数的增加使得雾化特征角缩小,可同时改善燃油蒸发并获得更加均匀的油气比分布。随着供油压力的提高,离心式与直流式喷嘴雾化特征角均增大,但供油压力不是影响直流喷嘴雾化锥角的主要因素。计算结果与试验结果对比定性符合良好,定量误差范围可以接受,验证了计算模型与计算方法的正确性,所得到的结果可应用于工程设计。 相似文献
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A systematic perturbation scheme is used to study the propagation of a weak shock wave attached to a slender body in a supersonic flow of plasma with thermal radiation and investigate as to how the coupling between the radiative transfer and magneto-hydrodynamic phenomena affects the flow field. The analytical solution of the flow field has been presented up to the second order of ε. The shape of the shock wave attached to the slender body of revolution is obtained, which however can be expressed explicitly in terms of known functions when the radiative decay length is of the same order as the typical body length. Also, the shock angle at the tip of the projectile is obtained. 相似文献
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带不同长度凹腔超声速燃烧数值研究 总被引:6,自引:0,他引:6
对带不同长深比凹腔的燃烧室三维燃烧流场进行数值模拟,研究了燃烧室流场结构。结果表明:液体碳氢燃料穿透深度较小;凹腔长深比对燃烧效率、总压损失影响较小,对燃烧室阻力影响显著。 相似文献
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The three-dimensional coupled implicit Reynolds Averaged Navier–Stokes (RANS) equations and the two equation standard k–ε turbulence model has been employed to numerically simulate the cold flow field in a typical cavity-based scramjet combustor. The numerical results show reasonable agreement with the schlieren photograph and the pressure distribution available in the open literature. The pressure distribution after the first pressure rise is under-predicted. There are five shock waves existing in the cold flow field of the referenced combustor. The first and second pressure rises on the upper wall of the combustor are predicted accurately with the medium grid. The other three shock waves occur in the core flow of the combustor. The location of the pressure rise due to these three shock waves could not be predicted accurately due to the presence of recirculation zone downstream of the small step. Further, the effect of length-to-depth ratio of the cavity and the back pressure on the wave structure in the combustor has been investigated. The obtained results show that there is an optimal length-to-depth ratio for the cavity to restrict the movement of the shock wave train in the flow field of the scramjet combustor. The low velocity region in the cavity affects the downstream flow field for low back pressure. The intensity of the shock wave generated at the exit of the isolator depends on the back pressure at the exit of the combustor and this in turn affects the pressure distribution on the upper wall of the combustor. 相似文献