Numerical study on performance of supersonic inlets with various three-dimensional bumps

Numerical investigations were performed with a supersonic inlet system installed with a three-dimensional bump which was substituted for a diverter or conventional ramp-type compression systems at Mach 2. The modified inlets were designed to have two oblique shocks and a terminal normal shock followed by a subsonic diffuser, with a circular cross-section throughout. A numerical analysis was conducted to understand the three-dimensional flow field including shock/boundary layer interactions that occur around a three-dimensional bump and to evaluate the performance of the supersonic inlets. The current numerical simulations showed a bump-type inlet based on a conventional ramp-type inlet can provide an improvement in total pressure recovery downstream of the shock/boundary-layer interaction over a ramp-type inlet.     

A numerical study of shock wave/boundary layer interaction in a supersonic compressor cascade

A numerical analysis of shock wave/boundary layer interaction in transonic/supersonic axial flow compressor cascade has been performed by using a characteristic upwind Navier-Stokes method with various turbulence models. Two equation turbulence models were applied to transonic/supersonic flows over a NACA 0012 airfoil. The results are superion to those from an algebraic turbulence model. High order TVD schemes predicted shock wave/boundary layer interactions reasonably well. However, the prediction of SWBLI depends more on turbulence models than high order schemes. In a supersonic axial flow cascade at M=1.59 and exit/inlet static pressure ratio of 2.21, k-μ and Shear Stress Transport (SST) models were numerically stables. However, the k-μ model predicted thicker shock waves in the flow passage. Losses due to shock/shock and shock/boundary layer interactions in transonic/supersonic compressor flowfields can be higher losses than viscous losses due to flow separation and viscous dissipation.     

Numerical simulation of flows around axisymmetric inlet with bleed regions

A numerical simulation of flows in an axisymmetric supersonic inlet with bleed regions is performed. An existing code which solves the Reynolds Averaged Navier-Stokes equations and the two-equation turbulence model equations is converted into an axisymmetric code. In addition, a bleed boundary condition model has been applied to the code. In this paper, the modified code is validated by comparing numerical results against experimental data and other computational results for flows on a bump and over an oblique shock with bleed region. Using the code, numerical simulation is performed for the flows in an inlet with multiple bleed regions.     

A numerical study on the effect of the flow control using bleeding systems in a bump-type inlet

In many supersonic inlets, several oblique shock waves are followed by a terminal normal shock wave. The normal shock-wave/turbulent boundary-layer interaction is critical with respect to its influence on the development of boundary layer throughout the subsonic diffuser and the total pressure recovery at the engine face. In the current study, the bump-type inlet was designed for two oblique shock waves and a normal shock wave. In addition, a porous surface was installed underneath the root of the normal shock wave. The effect of flow control on the interaction between the normal shock wave and turbulent boundary layer in supersonic inlets by using the bleeding system was investigated numerically and was evaluated with respect to the inlet performance parameters.

     

CFD investigation of pressure depressions in aerostatic circular thrust bearings

Under certain operating conditions, the pressure distributions of aerostatic thrust bearings experience an undesirable pressure depression which decreases their load carrying capacity. Many investigators reasoned this phenomenon to the occurrence of shock waves in the bearing clearance. Recently, some investigators reasoned this phenomenon to the transition from laminar to turbulent flow and claimed that no shock wave is generated at the boundary between supersonic and subsonic flows. As such, there is a contradiction between these two opinions. In this paper, the rationale of the pressure depression phenomenon in aerostatic thrust bearings is investigated using computational fluid dynamics (CFD) simulations. The turbulent full Navier-Stokes equations for steady, three-dimensional, compressible flows are numerically solved in this study. Two circular bearing configurations are analyzed. The obtained results showed that the predicted pressure distributions along the fluid film compare well with the corresponding experimental data of other investigators. The present computational methodology allowed a clear capturing of the coherent structures of the flowfield in the bearing inlet region which include the coalescing of compression waves into shock waves and the region of shock/boundary layer interaction (pseudo-shock). The thorough understanding of this phenomenon is the first step towards the development of its appropriate control methods.     

来流附面层对吸附式压气机叶栅影响的数值研究

数值研究低速条件下来流附面层特性对吸附式压气机叶栅气动性能的影响,在不同正冲角时,设计不同方案的附面层厚度分布,对比分析叶栅出口气动参数的分布以及叶栅内的三维流场结构,讨论不同来流附面层情况下在压气机叶栅中采用附面层吸除的效果。结果表明,数值模拟结果与试验数据有较好的一致性;附面层厚度的增加导致角区流动三维性增强,且变冲角性能下降,二次流横向作用和角区范围的增加使得附面层抽吸效果减弱;当来流附面层厚度增加时,通过增加抽吸量可以有效降低强吸附式压气机叶栅的损失,变冲角性能得到改善。     

Non-classical Elastohydrodynamic Lubricating Film Shape Under Large Slide-roll Ratios

The authors showed in previous experiments with high viscosity polymeric lubricants that a non-classical elastohydrodynamic (EHL) film, which featured an inlet dimple, could be generated under pure sliding conditions. The phenomenon was tentatively attributed to boundary slippage. In this paper, much greater sliding is introduced in the experiments to gain further insight into film formation under boundary slippage. By putting all of the results on a load versus entrainment speed chart, it is found that the required conditions for the formation of the inlet dimple fall into an open triangular region in the chart. The existence of the inlet dimple can be maintained for a larger speed range with a higher load. The minimum speed required (the lower speed bound for the dimple existence) decreases only marginally with an increase in load but the speed of the disappearance of the dimple (the upper speed bound) increases with an increasing load. Interferograms show that with an increase in the slide-roll ratio, i.e., expanded boundary slippage, a bump occurs before the exit constriction, which indicates an obvious drop in film thickness, and the location of the minimum film thickness in the whole EHL contact moves from the outlet constriction to the center of the bump. The observed inlet dimple and bump have already been described in the previous numerical results that consider boundary slippage, and provide more justification for the boundary slippage postulation in the experimental films.     

高压下环面节流静压气体轴承适用的润滑方程探讨

采用纯黏性润滑方程和基于层流模式、SST k-ω湍流模式的N-S方程,对环面节流静压气体润滑推力轴承内的压力分布进行了研究,分析了随着气膜厚度的变化,轴承流场内压力的变化及其变化机理。实验证明,在轴承厚度很小时两种方程求得的压力值与实验结果一致,然而随着气膜厚度的增大,采用纯黏性润滑方程计算所得结果的偏差很大,而采用N-S方程计算所得结果与实验结果基本一致,但在逆压力梯度段存在偏差;SST k-ω 湍流模式能较好地处理湍流剪切应力在逆压梯度边界层内的输运和激波与边界层的相互作用,准确模拟出气膜入口附近复杂的流动状态。      

Effect of boundary layer swirl on supersonic jet instabilities and thrust

This paper reports the effects of nozzle exit boundary layer swirl on the instability modes of underexpanded supersonic jets emerging from plane rectangular nozzles. The effects of boundary layer swirl at the nozzle exit on thrust and mixing of supersonic rectangular jets are also considered. The previous study was performed with a 30° boundary layer swirl (S=0.41) in a plane rectangular nozzle exit. At this study, a 45° boundary layer swirl (S=1.0) is applied in a plane rectangular nozzle exit. A three-dimensional unsteady compressible Reynolds-Averaged Navier-Stokes code with Baldwin-Lomax and Chien’sk-ε two-equation turbulence models was used for numerical simulation. A shock adaptive grid system was applied to enhance shock resolution. The nozzle aspect ratio used in this study was 5.0, and the fully-expanded jet Mach number was 1.526. The “flapping” and “pumping” oscillations were observed in the jet’s small dimension at frequencies of about 3,900Hz and 7,800Hz, respectively. In the jefs large dimension, “spanwise” oscillations at the same frequency as the small dimension’s “flapping“ oscillations were captured. As reported before with a 30° nozzle exit boundary layer swirl, the induction of 45° swirl to the nozzle exit boundary layer also strongly enhances jet mixing with the reduction of thrust by 10%.     

静压气体轴承中的激波与边界层相互影响的研究进展

综述空气动力学理论中激波—边界层相互影响的一般特性以及静压气体轴承中的激波与边界层相互影响理论与试验研究。指出随着计算技术和微流体测试技术的发展,气膜内的压力分布、速度分布和马赫数分布,根据流场各段不同的流动特点,将可以采用合适的湍流模型进行计算机模拟,并采用相关试验得到验证。静压气体轴承中的激波—边界层相互影响将得到更深入的研究,必将促进对较高供气压力和较大气膜间隙条件下,静压气体轴承气膜入口、出口转角区和气膜内轴承特性的全面了解。     

Comparison of turbulence models in shock-wave/boundary-layer interaction

This paper presents a comparative study of a fully coupled, upwind, compressible Navier-Stokes code with three two-equation models and the Baldwin-Lomax algebraic model in predicting transonic/supersonic flow. Thek - ε turbulence model of Abe performed well in predicting the pressure distributions and the velocity profiles near the flow separation over the axisymmetric bump, even though there were some discrepancies with the experimental data in the shear-stress distributions. Additionally, it is noted that this model hasy* in damping functions instead of y⁺. The turbulence model of Abe and Wilcox showed better agreements in skin friction coefficient distribution with the experimental data than the other models did for a supersonic compression ramp problem. Wilcox’ s model seems to be more reliable than the other models in terms of numerical stability. The two-equation models revealed that the redevelopment of the boundary layer was somewhat slow downstream of the reattachment portion.     

An experimental study on the effect of square grooves on decay characteristics of a supersonic jet from a circular nozzle

In this experimental work, the effect of square grooves on the structure of a supersonic jet emanating from a circular nozzle has been investigated at three different nozzle inlet total pressures i.e 360 kPa, 550 kPa and 720 kPa. The nominal exit Mach number is 1.8. A new empirical relation for predicting the supersonic core length for grooved nozzle has been suggested. Further, a new parameter “groove effectiveness” has also been suggested to quantify the effect of the groove by using the total pressure data in the supersonic core length. Experimental results suggest that at higher nozzle inlet total pressure, the groove effectiveness plays a minor role. From the jet centreline total pressure data, supersonic core length, the locations at which 50 % and 90 % decay occurs have been obtained. It has been observed that higher groove effectiveness is associated with smaller values of supersonic core length, L_50% and L_90%. Schlieren images of the jet structure shows unsymmetrical shock pattern of jets emanating from a single grooved nozzle.     

环面节流平面气浮轴承大间隙下流场数值模拟

利用三维粘性可压缩平均Navier-Stokes方程,采用SSTk-ω湍流模式,使用非结构化网格和二阶精度的有限体积法,对环面节流圆盘推力轴承大间隙下的流场进行了数值模拟并与层流模式数值结果、实验进行了对比,结果表明,SSTk-ω湍流模式数值模拟的压力分布与实测的压力分布符合程度比较高。马赫数和速度数值计算结果表明,流场内存在激波和激波边界层的相互干扰,导致壁面边界层流动出现了分离,压力出现激烈变化。     

Flow structure in critical flow Venturi nozzle and its effect on the flow rate

The flow fields in toroidal Venturi-nozzles, shaped according to the ISO-9300 Standard, have been investigated using numerical flow simulation. The present study was aimed at clarifying some of the phenomena associated with unchoking the flow in the throat. To this end, the shock structure has been studied for different Reynolds numbers and exit pressure ratios. The flow simulations were carried out in two and three dimensions. The flow fields were always unsteady, displaying a complex shock–boundary layer interaction.     

主动流动控制增大高超飞行器来流捕获量研究进展

高超声速进气道来流捕获量是确保进气道起动的重要因素,而在实际飞行环境中很难使捕获量达到设计值,总结了国内外研究团队通过实验和数值方法,对磁流体控制进气道斜激波位置、虚拟唇口增大进气质量流量、表面放电形成等离子体用于高速气流控制等方面进行的研究.他们对设计方案进行了不断地优化,并取得了增大进气道来流捕获量主动流动控制方法...     

Application of the scaling law for swept shock/boundary— layer interactions

An experimental study providing additional knowledge of quasi-conical symmetry in swept shock wave/turbulent boundary-layer interactions is described. When a turbulent boundary layer on the flat plate is subjected to interact with a swept planar shock wave, the interaction flowfield far from fin leading edge has a nature of conical symmetry, which topological features of the interaction How appear to emanate from a virtual conical origin. Surface streakline patterns obtained from the kerosene-lampblack tracings have been utilized to obtain representative surface features of the flow, including the location of the virtual conical origin. The scaling law for the sharp-fin interactions suggested by previous investigators has been reexamined for different freestream Mach numbers. It is noticed that the scaling law reasonably agrees with the present experimental data, however, that the law is not appropriate to estimate the location of the virtual conical origin. Further knowledge of the correlation for the virtual conical origin has thus been proposed.     

Study on drug powder acceleration in a micro shock tube

Recently, micro shock tubes have been widely used in the medical engineering. The needle-free drug delivery device which mainly consists of a micro shock tube and an expanded nozzle has been produced to inject drug powders into human and animal bodies without any sharp metal needles. The drug powders were delivered by obtaining high momentum, which can be done by accelerating drug powders in the micro shock tube and supersonic nozzle. The particle-gas flows are induced by the incident shock wave developing by rupturing the diaphragm in the micro shock tube and again accelerated in the supersonic nozzle. The momentum of injected drug particles should be strictly controlled otherwise patients will suffer from skin injury or hurt. Even though micro shock tubes have been investigated in the past several decades, the detailed studies on particle-gas flows in the micro shock tube were rare to date due to the micro size and difficult experimental operation on micro shock tubes. In this paper, the experimental and numerical studies were carried out on investigating particle-gas flows in a designed micro shock tube. Particle tracking velocimetry (PTV) was performed to calculated particle average velocity at the exit of the supersonic nozzle. The nozzle flows were analyzed by obtaining instantaneous particle fields. The particle number density ratio was also investigated in the test section. The numerical simulations were performed by calculating unsteady Naver-Stokes equations on compressible flows and using fully implicit finite volume schemes. Discrete phase model (DPM) was used for simulating particle-gas flows in the micro shock tube. Particle diameter and density were varied to investigate their effects on the particle-gas flows. Unsteady particle-gas flows and shock wave propagation were obtained in details in the micro shock tube for present experimental and numerical studies.     

空调室外机流场特性研究

分析了空调器室外机组内的空气流场特性,采用体积法对三维不可压缩Navier-Stokes方程在旋转坐标系中进行非定常求解。在随叶片旋转的区域与固定区域的交接面处,采用滑移网格技术求解动静非定常作用。计算中采用两类入口边界条件,即流量入口和压力入口,并将流场数值结果与微粒图像测速实验数据进行比较。比较结果表明,在流量入口边界条件下数值模拟计算的结果接近实验结果,流场计算可以为气动声学分析奠定基础。     

Negative electrode structure design in InSb focal plane array detector for deformation reduction

When an indium antimonide (InSb) infrared focal plane array (IRFPA) is subjected to a thermal shock test, most of the cracks originate from the region over the negative electrode, which restricts its final yield. In light of the proposed equivalent modeling, three negative electrode structures are assessed to eliminate the accumulated deformation around the negative electrode. Simulation results show that when a thicker indium bump array is connected directly with negative InSb material, the accumulated thermal deformation is the minimum, the top surface of InSb chip is the smoothest, and the square checkerboard buckling pattern, present clearly in both gold buffer layer and sparse thicker indium bump array structure, seems to be unclear. All these mean that a thicker indium bump array structure is a good choice, which will benefit to reduce fracture probability of InSb IRFPAs under thermal shock.     

Analysis of two dimensional and three dimensional supersonic turbulence flow around tandem cavities

The supersonic flows around tandem cavities were investigated by two-dimensional and three-dimensional numerical simulations using the Reynolds-Averaged Navier-Stokes (RANS) equation with thek-ω turbulence model. The flow around a cavity is characterized as unsteady flow because of the formation and dissipation of vortices due to the interaction between the freestream shear layer and cavity internal flow, the generation of shock and expansion waves, and the acoustic effect transmitted from wake flow to upstream. The upwind TVD scheme based on the flux vector split with van Leer’s limiter was used as the numerical method. Numerical calculations were performed by the parallel processing with time discretizations carried out by the 4th-order Runge-Kutta method. The aspect ratios of cavities are 3 for the first cavity and 1 for the second cavity. The ratio of cavity interval to depth is 1. The ratio of cavity width to depth is 1 in the case of three dimensional flow. The Mach number and the Reynolds number were 1.5 and 4.5 × 10⁵, respectively. The characteristics of the dominant frequency between twodimensional and three-dimensional flows were compared, and the characteristics of the second cavity flow due to the first cavity flow was analyzed. Both two dimensional and three dimensional flow oscillations were in the ‘shear layer mode’, which is based on the feedback mechanism of Rossiter’s formula. However, three dimensional flow was much less turbulent than two dimensional flow, depending on whether it could inflow and outflow laterally. The dominant frequencies of the two dimensional flow and three dimensional flows coincided with Rossiter’s 2nd mode frequency. The another dominant frequency of the three dimensional flow corresponded to Rossiter’s 1st mode frequency.