针对径向落差比ΔR/L为 0.38 的双涵道中介机匣通道内二次流损失的问题,通过改变轮毂局部型面补偿支板带来面积堵塞的构型设计,提高了流场的品质.结果表明:设计方案减小了通道内二次流带来的损失,使中介机匣内涵道总压恢复系数提高到0.993 9,内涵道总压损失系数降低了 9.05%.

在微型发动机燃烧室初步设计方案的基础上,结合流体网络法、化学反应器网络模型法和遗传算法进行微型发动机燃烧室一维优化设计,对微型发动机燃烧室的燃烧效率进行优化,程序可在预定范围内自动调整火焰筒内外双壁面各排孔孔径,在保证总压恢复系数大于95%的前提下使燃烧效率增加 1.7%,达到了微型发动机燃烧室的一维优化设计要求.

为了给出航空发动机折流燃烧室火焰筒的构型设计方法,按照总体尺寸设计和具体构型设计的步骤进行,对折流燃烧室进行初步设计,依次进行一维流量分配计算和三维流场计算,结果满足初步设计的流量分配预期结果和燃烧效率与总压恢复系数要求.

为了确定某航空发动机中介机匣内部的流场旋涡结构,对其进行数值模拟,采用Q准则对流场进行分析,探索流场中的旋涡结构.研究发现,中介机匣中的旋涡模型主要由马蹄涡、壁角涡、轮毂/支板角区分离涡及机匣出口分离涡等4个旋涡构成.壁角涡与轮毂/支板角区分离涡位于轮毂端壁与支板的角区,主要因为角区的压力场导致了通道内较大的流动损失.机匣出口分离涡由机匣壁面较大的凸曲率产生,导致支板尾缘靠近机匣壁面的流场产生低速区.

基于压流修正的航空发动机燃烧室流量分配的计算方法可以快速、准确地计算出燃烧室的流量分配,在此基础上加入燃烧室沿程热力参数计算以及火焰筒一维壁温计算模块,开发出能够快速完成燃烧室一维性能计算的程序.采用所开发的程序对微型蒸发管燃烧室进行一维计算,可以快速计算出燃烧室的流量分配、沿程热力参数分布以及火焰筒的一维壁温分布.将程序计算结果与采用Fluent软件计算的数值仿真结果进行对比,两者相对偏差不大,表明所开发的燃烧室一维计算程序能够快速、准确地评估燃烧室的一维性能.

为了解旋转爆震波在单侧部分受限情况下产生的波系结构及环缝内压力波动特点,以恰化学当量比的氢气/氧气混合物作为反应物,采用9组分19步的基元反应模型,忽略粘性、热传导和扩散等输运效应,对爆震波在开缝式弯管内的传播情况进行了数值模拟.研究结果表明,爆震波在开缝式弯管内传播时,爆震波及爆震区内的反射激波后的高压区都会在环缝内产生透射激波,但环缝内激波不单纯依赖弯管内高压区的侧向膨胀产生,且环缝内的一些反射激波产生的波后高压区还会反过来在爆震区内产生透射激波;环缝内第二道斜激波会被接触面分为两部分,在接触面前方的惰性气体一侧具有更大的倾斜角;环缝位于外侧时爆震波的倾斜程度比位于内侧时更大;当环缝处于外侧和内侧时,会由于激波相交、反射产生比第一道斜激波更大的压力峰值,内侧开缝时环缝内最大压力峰值约为第一道斜激波后的两倍,且内侧开缝时环缝内压力大于4倍填充压力的高压区域范围比外侧开缝时大36.4％.

为研究微型燃烧室蒸发管的雾化蒸发性能,试验研究了进气温度、气油比(AFR)、管壁温度和进口空气流速对燃油蒸发率的影响.试验结果表明:进气温度和进口空气流速是影响蒸发效率的两个主要因素;当气油比减小到3.0时,管内两相流型由膜态沸腾向过渡态沸腾转变,该状态下燃油与管壁的换热效率最低.蒸发管数值仿真引入离散相模型(DPM)和液滴碰壁飞溅模型,蒸发效率计算结果与试验数据呈现相同趋势.在此基础上研究了气动参数对燃油雾化的影响.计算结果表明,进口空气流速的提高可以改善燃油雾化细度,但不利于液滴分布的均匀性,索太尔平均直径(SMD)与进口空气流速的-1.69次方成正比.

Take the rotating detonation engine combustion chamber wall which under great heat flux as the research background. The characteristics of impingement cooling under different parameters were studied by numerical simulation, then analyzing and getting a better Impingement cooling structure. It is found that under the same air mass flow rate and constant heat flux, when the diameter of impingement hole is d=4mm, the impingement distance is H=4mm, the hole spacing is Xn=Yn= 35.168mm, the cooling effect is the best. With the cooling structure, the cooling of the heated wall with constant heat flux of 500000 W/m(2) was studied. The results show that when the Reynolds number is Re = 50000, the wall is cooled and the wall temperature is reduced to a reasonable range.

Improve the accuracy of ignition calculate simulation through dynamic thickening of the of flame front in order to achieve Large Eddy Simulation (LES) solution scales. Taking the WALL model which takes the sticky near the wall into account as the sub-grid turbulence model, simulating the process of energy releasing of spark plug by the way of adding Gauss heat source to the energy function. 40ms of the ignition process is calculated based on the cold state flow field, and compare the numerical simulation results with corresponding experimental results abroad. It shows: The simulation accuracy of cold flow field of LES is good, and be able to capture small-size vortex; Large Eddy Simulation based on dynamic thickening flame model is an effective method to predict the dynamic diffusion processes of CH4 non-premixed flame; Recirculation zone plays an important role of fuel distribution, flame spread and stability.

Particle image velocimetry (PIV) was used for learning the 5 mm diameter warped hole flow field in jet-in-crossflow. The result shows that: the jet angle differs from the warped hole angles, the larger of the warped hole angle means the larger jet angle; there will be a useful parameter named effective diameter from the experiment data. Jet flow trajectory curves were fitted in condition of two types of velocity, four different warped hole angles and three tunnels with different depth. The curve is well fitted with the real jet flow, contributing to the future flow field design in micro combustors or other small tunnels.

To reduce flow loss and simplify the structure of the inlet particle separator(IPS), hybrid optimization algorithm was introduced into the IPS optimization design. Firstly, four order spline curve was adopted as parameterized description of the IPS model and an automated simulation process was established. Optimal Latin hypercube design(OLHD) method was used to obtain the sample points, and an elliptical radial basis function(EBF) surrogate model was established to substitute the simulation code. Then non-dominated sorting genetic algorithm (NSGA-II) was used for multi-objective optimization based on surrogate model, and finally non-linear programming by quadratic Lagrangian (NLPQL) algorithm was applied to local optimization. As a result, a new configuration was obtained, with the scavenge efficiency of C-spec distribution improved to 100% and 86.7% for AC coarse distribution fine sand and total pressure reduced to less than 0.6% with smaller main outlet pressure distortion.

The effect of geometry and aerodynamic factors on the flow in trapped vortex cavity was evaluated.The effect of geometry was analyzed by altering the ratio of height to length（H/L） of the cavity,injection position and main-flow inlet angle;the impact of aerodynamics was analyzed by varying main-flow and second inject flow inlet velocity.Total pressure loss was measured by total pressure probe,and the velocity field was measured by particle image velocimetry（PIV）.The results reveal that：（1） When the position of the injection holes is in the bottom or the middle position,large scale vortexes can be formed in the trapped vortex cavity.A couple of opposite swirling vortexes are formed in the middle position;and only a single vortex can be formed in the bottom.（2） For the middle injection configuration,increasing the main-flow mach number and inlet angle can enhance the vortex intensity in the upside of the cavity.（3） The total pressure loss increases apparently with the growing main-flow inlet Mach number.

The concept of new form of combustor named as head-swirl micro-combustor was studied to improve the performance of the micro-combustor.The experiment on the ambient temperature and pressure was carried out,the resistance characteristic shows that the combustor couldn't get into modeling state.The characteristics of the limit blow-out show that,the head swirl takes great advantage in keeping the fire stabilization,and the loading parameter is more satisfactory than that without the head swirl,the combustion efficiency is 0.8 at most,the major reason to the low efficiency is attributed to the short residual time for gas to stay in the combustor.The structure of head-swirl is successfully used in the microcombustor design.

In this paper, influences of various swirlers on flow fields and combustion performance in annular combustors were numerically investigated. The block-structured grids of annular combustors with swirlers (dual-stage axial, swirl cup and three-stage axial) were generated by solving a system of elliptic partial differential equations with the zone data transmit method. Three-dimensional two-phase combustion integrated flow fields of three different annular combustors including swirler, flame tube, dump diffuser, inner and outer cooling annuluses were calculated in arbitrary curvilinear coordinates. The gas and liquid phases were treated with Eulerian and Lagrangian methods, respectively. Combustor performances were estimated by the empirical-analytical design method. Calculations are in agreement with experimental data. Predictions indicate that the numerical procedures are reliable for numerical investigation of influence of different swirlers on combustion performance in annular combustors.

CFD simulations were performed on four 2-D hypersonic inlets with different aspect ratios in constant area.Effects of aspect ratio were analyzed at designed Mach 6 and other Mach numbers with different back pressures.The investigation suggest that with the reducing of length and wetted area while rising of aspect ratio,total pressure recovery coefficient of external compress ramp,mass flow ratio and coefficient of internal drag decrease;Inlet with low aspect ratio,as the huge area of three-dimensional-flow near the side wall,performance of inlet is affected notable and the starting ability become worse;When the inlets work in peak back pressure with design Mach number,the pressure that the 2-D hypersonic inlets can endure would reduce with the rising of aspect ratios.

To discuss the operating characteristics of the inlet which is used in Turbojet/Ramjet Combined Cycle engine at Mach 0 to 5,the profile of an over/under TBCC variable geometry inlet was given.Using CFD tools,the two-dimensional flow-field of the inlet was shown with properties of flow at different operating points.At every flight Mach number,the inlet can start without boundary layer suction.Based on the flow-field of cruise and transition states,the operating characteristics at different back pressure were analyzed.The result indicates that the back pressure at exit has effect on the performance of the inlet,especially at the transition state——change of back pressure between over and under exit can bring on aerodynamic coupling effects in the two ducts.The aerodynamic characteristic diagram varied with flight conditions of the inlet was presented.The factors affecting the performance of the inlet and the aerodynamic characteristics at typical flight conditions were analyzed.

A micro combustor experimental and testing system was designed to learn the performance of centimeter-scale micro combustor with vaporizing tube,and also provide some technical data for subsequent design.Under room temperature and pressure,the performance of micro combustion combustor was tested,during which some data were made available,such as:resistance characteristic,weak extinction characteristic,combustion efficiency,specific volume calorific intensity and exit temperature distribution.

Viscous interaction phenomena in hypersonic flow are investigated by CFD on the flat plate,and the numerical results are in well agreement with the measured ones.A modified relation about induced pressure on wall is obtained and compared with experiment data and CFD results.This modified equation agrees well with all the related data we can obtain.Based on the CFD results,the distribution laws of pressure and other flow paramaters are analysed,and it is found that there is a peak value in pressure distribution and that the wall pressure is greater than the free flow pressure.Fit equations about the peak pressure and its position are also given based on the CFD results.

The development and the validation of a parallel finite difference NavierStokes solver called the NAPA code for unsteady,three-dimensional flow simulations on workstation clusters are described.The solver is parallelized by divided-zone technique and message passing interface(MPI) communication library and validated for flow problems in hypersonic inlet,around micro aviation vehicle and NACA0012 airfoil.Results of the NACA0012 problem by original and the parallelized NAPA code are agreement with the experimental data.All of the comparison show that the parallel implementation of NAPA code is successful.Experiments identify the efficiency and speedups of the parallel code.Results show that the parallel code has good performance.Influencing factors,such as the ratio of communication,load balance and communication mode are discussed.Finally,the code is tuned with Intel Cluster tools and Vtune and the algorithm with the low efficiency in the code is improved.The calculation time for the hypersonic flow is reduced by 55.33%.

Multi-zone coupling method is adopted to simulate three- dimensional two-phase reacting flow field in the annual combustor with a triple swirler in arbitrary curvilinear coordinates. The block-structured grid is generated by zone method and elliptic partial differential system. The k-e turbulent model, EBU-Arrhenius turbulent combustion model and six-flux radiation model and the particle trajectory model are used in calculation. On the non-staggered grid system, the discretization equations are solved with SIMPLE algorithm and PSIC method. The calculation results are reasonable, so it reveals that computer codes are reliable for the optimum design and development of combustor.