自由射流预混合微火焰的燃烧特性

对静止空气中自由射流微喷管甲烷/空气预混合火焰的燃烧特性进行了实验研究,考察了火焰高度特征及其相关影响因素,详细探讨了尺度变化对微火焰熄火极限的影响.结果表明:微喷管射流预混合火焰为层流火焰,火焰高度与微喷管出口流速成止比,火焰高度随当量比减小而减小;同一当量比下,无量纲参数H/d(火焰高度/微喷管直径)与出口Re数呈线性关系.微尺度效应导致预混合火焰淬熄速度明显增大,同时可燃极限当量比远大于1,微预混合火焰发生淬熄的主要原凶是微尺度下热量和质量扩散作用明显增强.     

来流马赫数影响高负荷扇形扩压叶栅气动性能的研究

为研究亚音速高负荷扇形扩压叶栅NACA0065-K48的变工况性能,采用数值方法研究了来流马赫数对叶栅气动性能和流场结构的影响,来流马赫数的取值范围为0.3～0.8。计算结果表明:随着来流马赫数的增大,叶栅静压比不断提高,但总压损失先减后增,马赫数在0.5～0.7范围内叶栅具有较好的综合气动性能。扇形叶栅下角区分离程度大于上角区的不平衡流动现象会随着马赫数的增大而加剧,通道涡则是这一发展趋势的主导,且马赫数达到0.7之后,下角区通道涡与集中脱落涡趋于融合。此外,下端壁分离螺旋点的形成与发展是控制下角区分离程度和损失大小的关键。     

喷嘴与安装孔之间的缝隙对微型燃气轮机燃烧室影响的数值分析

为了探究喷嘴与燃烧室壁面安装孔间的缝隙对微型燃气轮机燃烧室流动及燃烧特性的影响,运用三维数值计算软件,对30 kW微型燃气轮机燃烧室在不同面积缝隙下的燃烧过程进行了数值计算,得到了燃烧室内的流场及温度场,并对比分析了燃烧室各处的气体流量分配、燃烧室内部温度分布以及污染物排放量。计算结果表明:缝隙面积的变化对燃烧室内气量分配的影响是全局性的,随着缝隙面积的增大,缝隙内的气体流量增加,燃烧室其它各处的流量则相应减小。在贫燃的条件下,这一过程使得燃烧室内部的整体温度逐渐减低,随之C0的排放量小幅增大。此外,一定范围内的缝隙能够在降低燃烧室整体温度的同时维持火焰形态,有效降低NOx的排放量。     

不同喷管间距甲烷微火焰阵列温度场和燃尽特性研究

为了优化微阵列火焰燃烧加热系统,在相同的燃料负荷和喷管物理条件下,构建了甲烷预混和微火焰阵列燃烧模型,并研究了不同喷管中心间距对温度场和燃尽率特性的影响规律。研究结果表明,由若干微小喷管火焰优化组成的阵列可形成温度均匀的加热场;随着喷管中心间距减小,火焰间相互影响程度增加,均温加热场的温度提高;喷管中心间距继续减小,微喷管阵列火焰开始聚并、向大火焰转变,燃烧反应区间变长、均温场处的燃尽率下降,微喷管火焰丧失微火焰特性;因此确定微火焰阵列加热场喷管中心间距这一重要参数时,需综合考虑温度均匀性、热负荷、燃尽率和污染物等因素。     

喷管内传递特性对微尺度扩散火焰的影响

对均匀空气流中微尺度甲烷扩散燃烧进行了数值模拟,重点考察微喷管内的流动和传热传质对微尺度燃烧特性的影响.结果表明,在低流速下,内径为0.3 mm的微喷管内进气速度为1.0 m/s时燃料与空气的混合已经发生,混合气被管外的热量预热,同时火焰的热损失增加.在喷管直径一定时,减小燃料喷出速度,传热传质现象对微尺度甲烷扩散火焰特性的影响增强;当进气速度为0.5 m/s时,甲烷在微喷管内开始燃烧,放出热量.在进行微尺度解析计算时,必须包含一定的喷管区域.     

基于MEMS技术的石蜡驱动气体微流量调节阀研究

针对微纳卫星冷气推进系统和空间电推进系统变推力的任务需求,提出一种基于MEMS(微机电系统)技术的相变材料气体微流量调节阀设计方案以实现对推进工质的流量调节。利用MEMS加工工艺制得微致动器和微流量调节阀,通过表面轮廓仪测量微致动器调节膜片形变特性与输入电流的相互关系,通过间接法测量流经该调节阀的气体质量流量,结果表明:当输入电流小于12 mA,微致动器无明显形变,微阀气体流量无明显变化;当输入电流大于12 mA,微致动器形变明显增加,微阀气体流量显著减小;当输入电流大于48 mA,微致动器形变基本不变,微阀气体流量降至最低,接近关闭状态。因此气体微流量调节阀的有效工作电流区间为12～48 mA。     

超临界机组流动加速腐蚀边界层流动模拟与分析

流动加速腐蚀(FAC)是电厂高压漏水、管道断裂等主要事故的原因,采用Fluent软件模拟管道的流动加速腐蚀规律,将微观与宏观研究联系起来,分析了温度、流速、管型与边界层流动及流动加速腐蚀之间的规律。结果表明:与其他温度相比,150℃时管道边界层处的Fe2+质量浓度增大,边界层的厚度更薄,边界层的腐蚀较严重,对流动加速腐蚀的影响也较大,为最易腐蚀温度;腐蚀速率随流速的增大而增大;T型三通管相比直管易发生腐蚀,且在流向改变处腐蚀最严重。     

射流预冷对高空高马赫数下压气机性能影响的数值研究

为研究射流液滴与空气双向耦合流动对压气机内部流场及其工作性能的影响,以NASA Stage35为模型,基于相似理论得到高空高马赫数下相似流场和压气机进出口条件,并对多工况下不同喷水量和液滴粒径下的射流冷却湿压缩过程进行分析。研究表明:射流预冷技术可有效抑制压气机进气同比温升。液滴与空气双向耦合流动使得压气机内部流场发生变化,有效降低叶片载荷的同时使动叶内的激波后移。在空气质量流量的0～2%的射流范围内,随着喷雾粒径的增大,压气机压比先增大后减小;比耗功量随喷雾量的增多而减少。25 km高空3.5Ma工况下,5μm粒径且喷雾量为空气质量流量的2%时,液滴蒸发率超过50%,压气机出口温度下降约20%,实际比压缩耗功同比减少约12%,压气机等熵效率提升约8%。     

微小尺度典型冷却结构的相似流动特性研究

以航空发动机空气系统中典型结构之一的孔元件为对象,研究微小尺度相似放大流动特性,研究相似流动中Re(雷诺数)和Ma(马赫数)对孔流动特性的影响。以空气为工作介质,通过数值计算小孔与相似放大孔的流量系数,对比内部流场及速度场。研究结果表明:Ma对突缩处的流动特性影响较大,Re对突扩处的流动特性影响较大;在Ma和Re相同的情况下,小孔与相似放大孔流量系数基本吻合,突扩和突缩处的流线分布、整体的速度云图分布也基本一致;如果Ma和Re有一个不相同,小孔与相似放大孔结果会产生较大差异。总之,在相似放大研究微尺度结构的流动特性时,不仅要保证几何结构和单值性条件相似,而且需保证无量纲参数Re和Ma分别相等。     

预旋对迷宫密封内流动传热特性影响的研究

采用三维周期性模型对发散型光滑面迷宫密封内传热及流动特性进行了研究,得出了两种流量下有无进口预旋时密封的间隙热系数随周向马赫数的变化关系,并与实验值、经验公式、二维轴对称模型得到的结果进行了比较,结果表明:该模型能较好地模拟有进口预旋时迷宫密封内的传热特性.在相同的流量和进口预旋比条件下,间隙热系数随转速的增加而增大;在相同流量和转速下,施加进口预旋能明显降低密封内总温升、减小间隙热系数,但不会影响子午面上的速度场;在相同转速和预旋比条件下,随着流量的增大,间隙热系数减小,子午面上速度增大,但流场结构不会发生变化.     

Temperature-Invariant Scaling for Compressible Turbulent Boundary Layers with Wall Heat Transfer

In a recent paper by Zhang et al. in 2012, a Mach number-invariant scaling was proposed to account for the effect of variation of free-stream Mach number in supersonic turbulent boundary layers. The present work focuses on the effect of variation of wall temperature with strong heating and cooling at the wall. Direct numerical simulation is used to study scaling and turbulence structure of a spatially evolving Mach 2 supersonic boundary layer at a friction Reynolds number of 500. A new scaling law is proposed to account for temperature-dependent fluid-property variations. This universal scaling appears superior to the existing models with the novelty that it applies not only for the mean-velocity profile but also extends to the turbulent transport, production, and dissipation terms in the budget of the turbulent kinetic energy.     

Effect of Mach number on thermoelectric performance of SiC ceramics nose-tip for supersonic vehicles

This paper focus on the effects of Mach number on thermoelectric energy conversion for the limitation of aero-heating and the feasibility of energy harvesting on supersonic vehicles. A model of nose-tip structure constructed with SiC ceramics is developed to numerically study the thermoelectric performance in a supersonic flow field by employing the computational fluid dynamics and the thermal conduction theory. Results are given in the cases of different Mach numbers. Moreover, the thermoelectric performance in each case is predicted with and without Thomson heat, respectively. Due to the increase of Mach number, both the temperature difference and the conductive heat flux between the hot side and the cold side of nose tip are increased. This results in the growth of the thermoelectric power generated and the energy conversion efficiency. With respect to the Thomson effect, over 50% of total power generated converts to Thomson heat, which greatly reduces the thermoelectric power and efficiency. However, whether the Thomson effect is considered or not, with the Mach number increasing from 2.5 to 4.5, the thermoelectric performance can be effectively improved.     

The research of laminar-turbulent transition in hypersonic three-dimensional boundary layer

The results of experimental investigation of laminar-turbulent transition in three-dimensional flow under the high continuous pressure gradient including the flow with local boundary layer separation are presented. The experimental studies were performed within the Mach number range from 4 to 6 and Reynolds number 10-60×106 1/m, the angles of attack were 0°and 5°. The experiments were carried out on the three-dimensional convergent inlet model with and without sidewalls. The influence of artificial tubulator of boundary layer on transition and flow structure was studied. The conducted researches have shown that adverse pressure gradient increase hastens transition and leads to decrease of transition area length. If pressure gradient rises velocity profile fullness increases and profile transformation from laminar to turbulent occurs. As a result of it the decrease of separation area length occurs. The same effect was reached with Reynolds number increase. These results are compared with the data on two-dimensional model with longitudinal curvature.     

The effects of outlet boundary conditions on simulating supersonic microchannel flows using DSMC

High speed gas flows through two-dimensional microchannels have been investigated using the Direct Simulation Monte Carlo (DSMC) method, where the pressure boundary condition has been implemented using the theory of characteristics as an alternative to the vacuum boundary. Two species, nitrogen and helium, have been used to conduct the flow simulations. It was found that the pressure boundary condition cannot only predict the flow with exit-plane pressure equal to the back pressure, which the vacuum boundary condition fails to do, but can also simulate the flow with expansion waves outside the channel. Therefore, it is considered to be more appropriate. Two inlet Mach numbers, 4.15 and 3.39, have been employed for the nitrogen flow cases with an inlet Knudsen number (Kn) of 0.062. It has been shown that for cases with an inlet Mach number equal to 4.15, the back pressure only has an effect on flow in the latter half of the microchannel, where the wall heat flux can be enhanced by increasing the back pressure. At an inlet Mach number of 3.39, the wall heat flux has the same trend as that in the higher Mach number case, though its magnitude is considerably lower. In addition, no significant effect of a step change in wall temperature distribution on the total heat exchange between the wall and the bulk flow was detected for the same inlet Mach number and back pressure.     

Numerical characterization of 3D nonreacting supersonic cavity combustor with inlet Mach number variation

A numerical investigation of a cavity-based supersonic combustor with non-reacting upstream hydrogen fuel injection is conducted to study the effects of inlet Mach number (Ma) on flow structure and fuel-air mixing. Three different freestream Mach number cases (1.5, 2.5 and 3.5) are investigated at a constant fuel flow rate, injected at the sonic condition by considering governing equations for compressible, turbulent flow using Shear Stress Transport (SST) k-ω model. The complex flow structure is investigated by identifying various flow features namely, upstream three-dimensional bow shock, compression waves, Mach reflection, vortex in the separated boundary layer and horseshoe vortices at the downstream of the injection port. Besides this, the flow physics involved in these complex flow features are unravelled. Moreover, the performance of the combustor is characterized quantitatively in terms of mixing efficiency, total pressure loss and coefficient of pressure. However, the mixing efficiency and total pressure loss for the operating condition of Ma = 1.5, exhibits better performance than that of the other Mach number cases (2.5, 3.5) due to decrease in inclination angle of reattachment shock from 47.6° to 29.9°. The present numerical investigation also demonstrates that the three-dimensional simulation is essential in the characterization of fuel-air mixing in supersonic cavity-based combustors.     

The initiation and propagation of detonation in supersonic combustible flow with boundary layer

Adaptive simulations solving the Navier-Stokes equations have been conducted in order to get a better understanding on the detonation initiation and propagation in a stoichiometric H₂/O₂/Ar supersonic mixture with boundary layer. The detonation is initiated by a continuous hot jet. When reflecting on the wall, the jet induced bow shock interacts with the boundary layer and forms the shock boundary layer interaction phenomena, while in Euler result the bow shock forms Mach reflection. The investigation shows that the Navier-Stokes simulation result is structurally in better agreement with the experiment compared with that of the inviscid Euler simulation result. The bow shock interacts with the separation shock, forming the shock induced combustion behind the interaction zone. Then the combustion front couples with shock and forms Mach stem induced detonation. The Mach stem induced detonation continues to getting higher and propagating upstream, initiating the main flow. The initiated partial detonation exists with the separation shock induced combustion front, forming an “oblique shock induced combustion-partial detonation” structure in the main flow. The investigation on the influence of free stream Mach number further confirms that the boundary layer has an important influence on detonation initiation. The parametric studies also show that there exists a free stream Mach number range to initiate the partial detonation in supersonic combustible flow successfully.     

Simulation and experimental measurements of 10-kW PEMFC passive hydrogen recovery system

Simulations are performed to examine the performance of a vacuum ejector in the hydrogen recovery loop of a 10-kW PEMFC system. The simulations commence by examining the effects of the primary flow fluid pressure and secondary flow temperature on the recirculation ratio and hydrogen stoichiometric ratio. Further simulations are then performed to investigate the temperature, pressure, velocity and Mach number distributions within the ejector for various values of the primary flow inlet pressure and temperature. A prototype ejector is fabricated using a 3D printing technique. Experiments are performed to evaluate the gas tightness and gas recovery performance of the ejector under realistic operating conditions. The simulation results show that the recirculation ratio and hydrogen stoichiometric ratio increase with a decreasing primary flow inlet pressure and secondary flow inlet temperature. As the primary flow inlet pressure increases, the pressure, velocity, and Mach number in the mixing chamber increase, and the hydrogen recovery performance decreases. Furthermore, as the temperature of the primary fluid flow increases, the stability of the isentropic flow condition within the ejector is enhanced. The experimental results show that the prototype vacuum ejector has a maximum gas leakage of just 0.7 psi and a minimum hydrogen recirculation rate of 59.3%. Consequently, it has significant potential for passive hydrogen recovery in large-scale fuel cell systems.     

Simulation of compressible flow in high pressure buried gas pipelines

The aim of this work is to analyze the gas flow in high pressure buried pipelines subjected to wall friction and heat transfer. The governing equations for one-dimensional compressible pipe flow are derived and solved numerically. The effects of friction, heat transfer from the wall and inlet temperature on various parameters such as pressure, temperature, Mach number and mass flow rate of the gas are investigated. The numerical scheme and numerical solution was confirmed by some previous numerical studies and available experimental data. The results show that the rate of heat transfer has not a considerable effect on inflow Mach number, but it can reduce the choking length in larger f_DL/D values. The temperature loss will also increase in this case, if smaller pressure drop is desired along the pipe. The results also indicate that for f_DL/D = 150, decreasing the rate of heat transfer from the pipe wall, indicated here by Biot number from 100 to 0.001, will cause an increase of about 7% in the rate of mass flow carried by the pipeline, while for f_DL/D = 50, the change in the rate of mass flow has not a considerable effect. Furthermore, the mass flow rate of choked flow could be increased if the gas flow is cooled before entrance to the pipe.     

Performance of 2D scheme and different models in predicting flow turbulence and heat transfer through a supersonic turbine nozzle cascade

A number of turbulence models were employed to investigate the heat transfer and aerodynamic characteristics through a nozzle cascade of a high-pressure gas turbine. Isentropic Mach number and Nusselt number around the vane were predicted and compared to existing experimental data obtained at a supersonic flow condition. According to the result presented by different models, possible source of the prediction error was identified; and the performance of different turbulence closures in predicting the heat transfer characteristics around the vane was discussed. It shows that the calculated heat transfer result was affected directly by the predicted turbulence transportation throughout the boundary layer. Finally considering the computational cost and the performance of the models, the suitable model(s) are recommended for the further 3D applications.     

Numerical and Experimental Studies of 3D Hypersonic Inlet

The results of numerical and experimental studies of a new configuration of 3D hypersonic inlet with the minimum throat area, which was called a convergent inlet, are presented in this paper. It is shown that the use of this inlet configuration allows one to reduce the drag and thermal protection of surfaces of a hypersonic engine within the entire range of flight velocities. The calculations were performed within the framework of inviscid gas model by the method of finite volumes. The flow and inlet characteristics, taking account of viscosity, were also calculated using the boundary layer equations. The experimental studies were performed within the Mach number range from 2 to 10.7 and Reynolds number based on the model inlet height of Re=1-5×106. The results included the flow parameters on the external compression surface and in the inlet duct, the Mach number in the throat, the air flow rate, the total pressure recovery coefficient, the inlet drag, and the boundary layer characteristics on compressi