Stall characteristics and tip clearance effects in forward swept axial compressor rotors

Tilting the blade sections to the flow direction (blade sweep) would increase the operating range of an axial compressor due to modifications in the pressure and velocity fields on the suction surface. On the other hand, blade tip gap, though finite, has great influence on the performance of a turbomachine. The present paper investigates the combined effect of these two factors on various flow characteristics in a low speed axial flow compressor. For this present study, nine computational domains were modeled; three rotor sweep configurations (0°, 20° and 30°) and for three different clearance levels for each rotor. Commercial CFD solver ANSYS CFX 11.0 is used for the simulations. Results indicated that tip chordline sweep is found to improve the stall margin of the compressor by modifying the suction surface boundary layer migration phenomenon. Diffusion Factor (DF) contours showed the severity of stalling with unswept rotor. For the swept rotors, the zones of high probable stall are less severe and they become less in size with increasing sweep. Increment in the tip gap is found to gradually affect the performance of unswept rotor, while the effect is very high for the two swept rotors for the earlier increments. As a minimum clearance is unavoidable, swept rotors suffer relatively higher deviation from the idealistic behavior than the unswept rotor due to tip clearance.     

基于数值优化的跨音速压气机动叶三维设计

实验验证了三维粘性流场求解程序,然后采用基于梯度法的数值优化程序对跨音压气机动叶积叠线进行优化设计,得到了弯掠结合的三维叶片,并对其进行了数值模拟及详细的流场分析.结果表明采用弯掠的三维动叶可以有效的改变叶片排内三维激波结构,降低尾迹损失,显著提高动叶的整体绝热效率,并使动叶具有更加良好的变工况性能.     

Control of transonic flow fields using local occurrence of non-equilibrium condensation

Control of supersonic flow fields with shock wave is important for some industrial fields. There are many studies for control of the supersonic flow fields using active or passive control. When non-equilibrium condensation occurs in a supersonic flow field, the flow is affected by latent heat released. Many studies for the condensation have been conducted and the characteristics have been almost clarified. Further, it was found that non-equilibrium condensation can control the flow field. In these studies, the condensation occurs across the passage of the flow field and it causes the total pressure loss in the flow field. However, local occurrence of non-equilibrium condensation in the flow field may change the characteristics of total pressure loss compared with that by the condensation across the passage of the nozzle and there are few for researches of locally occurred non-equilibrium condensation in supersonic flow field. The purpose in the present study is to clarify the effect of local occurrence of non-equilibrium condensation on the transonic flow field in a nozzle with a circular bump. As a result, local occurrence of non-equilibrium condensation reduced the shock strength and total pressure loss in the transonic flow field by flowing the moist air from trailing edge of the circular bump to the mainstream.     

Effects of aspect ratio in a transonic shock tube airfoil flow

In the present study, the flow visualizations were performed around the NACA 0012 models which differ in aspect ratios. We discussed the effects of the aspect ratio in the test models. Additionally the unsteady, two-dimensional, compressible Euler equations were solved for the NACA 0012 airfoil. Experiments were performed utilizing the conventional gas driven shock tube as the intermittent transonic wind tunnel. The aspect ratios of the models are about 0.86 and 1.5, respectively. The Mach numbers M 2 are about 0.84. The Reynolds numbers of the present experimental conditions were constant that Re based on chord length is about 4.0×10 5 . The results are as follows: in different aspect ratios, the difference of the shock wave location is confirmed though the Mach number and Reynolds number are same. It indicates the different correction Mach number by the effects of the side wall boundary layer though the nominal Mach number measured the same value. Also, on the difference of shock wave location for the effects of the aspect ratio, the tend of CFD shows the qualitative agreement with the result of an experiment.     

Investigation on multiple cylindrical holes casing treatment for transonic axial compressor stability enhancement

Casing treatment is one possible way of regaining axial compressor operating range. However, most of casing treatments extend the operating range with the cost of efficiency penalty. A new form of multiple cylindrical holes casing treatment (MHCT) with pre-swirl blowing for the NASA Rotor-37 has been designed based on profound understanding of the stall inception. Unsteady numerical simulations have been performed for Rotor-37 with and without MHCT. Parametric studies of the total extraction holes area and their axial locations show that the compressor performance deteriorates as the area ratio increases but the stall margin is extended and there is an optimum extraction holes axial location for stall margin extending. The better configuration of MHCT could extend the stall margin by 6.2% with only 0.23% peak efficiency reduction. Detailed analysis of the physical mechanism behind the stall margin improvement shows that the casing treatment could eliminate the passage blockage by suppressing breakup of tip leakage vortex and decrease the blade load in tip region, which both contribute to improve stall margin of transonic axial compressors.     

An improvement on the efficiency of a single rotor transonic compressor by reducing the shock wave strength on the blade suction surfaces

It is well known that increasing the rotational velocity is an effective way to increase the total pressure ratio. With increasing flow velocity especially under the condition of transonic flow in the supersonic region, where exist strong shock waves, the shock wave loss becomes main and important. Simultaneously, there occurs boundary layer separation due to the shock wave / boundary layer interaction. In the present paper the transonic compressor blades were studied and analyzed to find a proper and simple way to reduce the shock wave loss by optimizing the suction surface configuration or controlling the gradient of isentropic Mach number on the suction surface. A Navier-Stokes solver combined with a modified design algorithm was developed and used. The NASA single rotor for transonic flow compressor was served as a numerical example to show the effectiveness of this method. Two cases for both original and modified rotors were analyzed and compared.     

正倾斜叶片压气机叶栅二次流的数值研究

应用Beam-Warming近似隐式因子分解格式以及MML代数湍流模型,采用拟压缩性方法求解雷诺平均拟压缩N－S方程组,对正倾斜叶片压气机叶栅内三维粘性流场进行了数值研究,并与直叶栅进行了对照。结果发现,正倾斜叶栅中上、下通道涡的发生、发展过程与直叶栅存在明显的差异,这导致正倾斜侧二次流减弱,负倾斜侧二次流高损失区扩大,流动状况恶化,叶栅顶部区域的附面层分离发展成一个向叶栅中部扩展的更大的区域。计算与实验结果比较,两者吻合较好。     

The effect of non-equilibrium condensation on the drag coefficient in a transonic airfoil flow

In this study,a transonic flow past NACA0012 profile at angle of attack α=00 whose aspect ratio AR is 1.0 with non-equilibrium condensation is analyzed by numerical analysis using a TVD scheme and is investigated using an intermittent indraft type supersonic wind tunnel.Transonic flows of 0.78-0.90 in free stream Mach number with the variations of the stagnation relative humidity(Φ0)are tested.For the same free stream Mach number,the increase in Φ0 causes decrease in the drag coefficient of profile which is composed of the drag components of form,viscous and wave.In the case of the same M∞ and T0,for more than Φ0=30%,despite the irreversibility of process in non-equilibrium condensation,the drag by shock wave decreases considerably with the increase of Φ0.On the other hand,it shows that the effect of condensation on the drag coefficients of form and viscous is negligible.As an example,the decreasing rate in the drag coefficient of profile caused by the influence of non-equilibrium condensation for the case of M∞=0.9 and Φ0=50% amounts to 34%.Also,it were turned out that the size of supersonic bubble(that is,the maximum height of supersonic zone)and the deviation of pressure coefficient from the value for M=1 decrease with the increase of Φ0 for the same M∞.     

A study of the unsteady tip flow field of a transonic compressor

An experimental investigation on the unsteady tip flow field of a transonic compressor rotor has been performed.The casing-mounted high frequency response pressure transducers were arranged along both the blade chord and the blade pitch.The chord-wise ones were used to indicate both the ensemble averaged and time varying flow structure of the tip region of the rotor at different operating points under 95% design speed and 60% design speed.The pitch-wise circumferential transducers were mainly used to analyze the unsteadiness frequency of the tip leakage flow in the rotor frame at the near stall condition.The contours of casing wall pressure show that there were two clear low pressure regions in blade passages,one along the chord direction,caused by the leakage flow and the other along the tangential direction,maybe caused by the forward swept leading edge.Both low pressure regions were originated from the leading edge and formed a scissor-like flow pattern.At 95% design speed condition,the shock wave interacted with the low pressure region and made the flow field unsteady.With the mass flow reduced,the two low pressure regions gradually contracted to the leading edge and then a spike disturbance emerged.     

Aerodynamic sweeping study and design for transonic compressor rotor blades

<正>The objective of the present paper is to study the sweep effect on the blade design performance of a transonic compressor rotor.The baseline to be modified and swept is a designed well efficient transonic single rotor compressor. The first part of the present study is concerning the sweep effect with straight leading edge.In this case fixing the hub section the swept blade is formed by tilting the leading edge with whole blade forwards and backwards axially.The second part is to use an optimization strategy with simple gradient-based optimum-searching method and multi-section blade parameterization technique to search and generate an optimal swept rotor with curved arbitrary leading edge.Its adiabatic efficiency is a little bit greater than that of the reference un-swept rotor.     

Blade parameterization and aerodynamic design optimization for a 3D transonic compressor rotor

The present paper describes an optimization methodology for aerodynamic design of turbomachinery combinedwith a rapid 3D blade and grid generator(RAPID3DGRID),a N.S.solver,a blade parameterization method(BPM),a gradient-based parameterization-analyzing method(GPAM),a response surface method(RSM)withzooming algorithm and a simple gradient method.By the use of blade parameterization method a transonic com-pressor rotor can be expressed by a set of polynomials,and then it enables us to transform coordinate-expressedblade data to parameter-expressed and then to reduce the number of parameters.With changing any one of theparameters and by applying grid generator and N.S.solver,we can obtain several groups of samples.Here onlyten parameters were considered to search an optimized compressor rotor.As a result of optimization,the adiabaticefficiency was increased by 1.73%.     

Detailed analysis of the flow in the inducer of a transonic centrifugal compressor

Numerical and experimental investigations were conducted in a transonic centrifugal compressor stage composed of a backswept splittered unshrouded impeller and a vaned diffuser. A detailed analysis of the flow in the inducer (i.e. the entry zone of the impeller between the main blade leading edge and the splitter blade leading edge) is proposed from choke to surge. Steady and unsteady simulations were performed using the code elsA, which uses a multi-domain approach on structured meshes and solves the compressible RANS equations, associated with a two-equation turbulence model k-l in the rotating frame of reference. The 1MW LMFA-ECL test rig was used for carrying out the tests in the compressor stage. Unsteady pressure measurements up to 150 kHz and Laser Doppler Anemometry measurements were performed in the inducer. A good agreement is obtained between the experimental and numerical data even if an over dissipation is noticed in the numerical results. The change in flow pattern from choke to surge is mainly due to a change in the tip leakage flow trajectory which straightens, leading to a flow blockage of an individual passage near shroud. A spectral analysis shows that only the blade passing frequency and its harmonics compose the various spectra obtained from choke to surge.     

Numerical and performance analysis of one row transonic rotor with sweep and lean angle

In this study,aerodynamic behaviors of swept and leaned blades were investigated.Axial and tangential blade curvatures impacts on compressor＇s operating parameters were analyzed separately.A commercial CFD program which solves the Reynolds-averaged Navier-Stokes equations was used to find out the mentioned impact and the complicated flow field of transonic compressor-rotors.The CFD method that was used for solving flow field＇s equation was validated by experimental data of NASA Rotor 67.The results showed that the compressor with curved rotors had higher efficiency,rotor pressure ratio and stable operating range compared to the compressor with un-curved rotors.Using curved rotors mostly had higher impact on the overall stable operating range compared to the other operating parameters.Operating range involves choking point and stall point that were changed separately by using of bended blade.For finding the detailed impact of sweep and lean angle on transonic blades,various forms of lean and sweep angles were exerted to basic rotor.It was found that sweep angles increased overall operating range up to 30%,efficiency up to 2%and pressure ratio up to 1%.Leaning the blades increased the safe operating range,the pressure ratio and efficiency by 14%,4%and 2%respectively.     

Numerical investigation of the unsteady tip leakage flow and rotating stall inception in a transonic compressor

<正>It is well known that tip leakage flow has a strong effect on the compressor performance and stability. This paper reports on a numerical investigation of detailed flow structures in an isolated transonic compressor rotor-NASA Rotor 37 at near stall and stalled conditions aimed at improving understanding of changes in 3D tip leakage flow structures with rotating stall inception.Steady and unsteady 3D Navier-Stokes analyses were conducted to investigate flow structures in the same rotor.For steady analysis,the predicted results agree well with the experimental data for the estimation of compressor rotor global performance.For unsteady flow analysis, the unsteady flow nature caused by the breakdown of the tip leakage vortex in blade tip region in the transonic compressor rotor at near stall condition has been captured with a single blade passage.On the other hand, the time-accurate unsteady computations of multi-blade passage at near stall condition indicate that the unsteady breakdown of the tip leakage vortex triggered the short length-scale-spike type rotating stall inception at blade tip region.It was the forward spillage of the tip leakage flow at blade leading edge resulting in the spike stall inception. As the mass flow ratio is decreased,the rotating stall cell was further developed in the blade passage.     

Singular Surfaces in the Theory of Thermoelasticity with Microtemperatures

This work concerns the behavior of shock waves and higher-order discontinuities which propagate in a thermoelastic body with inner structure and microtemperatures. The response of the material is assumed to be linear. The propagation conditions and growth equations that govern the propagation of singular surfaces of order r ≥ 1 are derived. The coupling between the discontinuities in the mechanical and thermal fields is discussed.     

Numerical investigation of tip clearance effects in an axial transonic compressor

Numerical investigations of the Darmstadt transonic single stage compressor (DTC), in the Rotor1-Stator1 configuration, aimed at advancing the understanding of the effect of different rotor tip gaps and transition modelling on the blade surfaces are presented. Steady three dimensional Reynolds Averaged Navier Stokes (RANS) simulations were performed to obtain the flow fields for the different configurations at different operating conditions using the RANS-Solver TRACE. The stage geometry and the multi-block structured grid were generated by G3DMESH and a grid sensitivity analysis was conducted. For the clearance gap region, a fully gridded special H-grid was chosen. Comparisons were made between the flow characteristic at design speed, representative for a transonic flow regime, and at 65% speed, representative for a subsonic flow regime. The computations were used to analyse the flow phenomena through the tip clearance region for the different configurations and their impact on the performance of the compressor stage.     

单个电磁振动风扇扇叶形状换热特性研究

电磁振动风扇主要由柔性扇叶和电磁驱动装置组成,工作原理是通过电磁力驱动扇叶振动,使周围的空气流动,达到散热的目的。扇叶优化是提高振动风扇散热能力的重要途径,在矩形扇叶的基础上提出锯齿形扇叶和折扇形扇叶。使用FLUENT软件进行数值模拟,研究单个矩形、锯齿形和折扇形扇叶诱导的流场特性和换热特性。结果表明：与矩形扇叶相比,锯齿形与折扇形扇叶均有增强散热的积极作用,可提升最高气流速度分别为0.4 和0.8 m/s,提升局部散热分别为5.8 %和9.0 %。     

An experimental study of the effect of 2.5% methane addition on self-ignition and flame propagation during high-pressure hydrogen release through a tube

This paper demonstrates experimental investigation on the self-ignition and subsequent flame propagation of high-pressure hydrogen-methane mixture release via a tube. The proportion of methane added to hydrogen is 2.5% (vol.). A transparent rectangular tube (d = 15 mm, L = 400 mm) is used in the experiments. It is shown that the minimum burst pressure required for self-ignition increases 1.57 times for only 2.5% methane addition from 2.89 MPa (pure hydrogen) up to 4.68 MPa (2.5% CH₄ addition). This is mainly caused by the following reasons: on the one hand, methane addition can result in the decease of shock intensity inside the tube, thereby lowering the temperature of the combustible mixture; on the other hand, the hydrogen-methane mixture has the higher minimum ignition energy than that of pure hydrogen. Besides, 2.5% methane addition can increase the initial ignition time, weaken the flame intensity and reduce the flame propagation velocity relative to tube wall inside the tube. Moreover, for cases with 2.5% methane addition, the complete flame throughout the tube is formed closer to the back end of the tube. When the self-sustained flame exits from the tube, the maximum overpressure in a confined space increases with 2.5% methane addition.     

Research progress on the self-ignition of high-pressure hydrogen discharge: A review

As one of the most promising fossil energy substitutes, hydrogen energy is receiving increasing attention, and it has been greatly developed in recent years. However, hydrogen safety issues limit the large-scale application of hydrogen energy. Since 1922, the issue of self-ignition of high-pressure hydrogen discharge has gradually become the focus of attention of scholars in the field of hydrogen energy. Particularly fruitful research results have been obtained in the past 20 years, showing that the minimum discharge pressure of hydrogen self-ignition is approximately 2 MPa. In particular, the discharge tube shape and bursting disc rupture have a significant effect on the characteristics of hydrogen self-ignition. Moreover, the study of the hydrogen self-ignition mechanism under special working conditions has been extended by shock-induced ignition theory. Initial conditions mainly affect the critical pressure of hydrogen self-ignition by changing the formation, development and propagation of shock waves. Finally, the deficiencies and future research trends in research methods, self-ignition characteristics, and dynamic mechanisms are analysed.     

Numerical analysis of the tip leakage flow field in a transonic axial compressor with circumferential casing treatment

This paper reports on numerical investigations aimed at understanding the influence of
circumferential casing grooves on the tip leakage flow and its resulting vortical structures.The results
and conclusions are based on steady state 3D numerical simulations of the well-known transonic axial
compressor NASA Rotor 37 near stall operating conditions.The calculations carried out on the casing
treatment configuration reveal an important modification of the vortex topology at the rotor tip
clearance.Circumferential grooves limit the expansion of the tip leakage vortex in the direction
perpendicular to the blade chord,but generate a set of secondary tip leakage vortices due to the
interaction with the leakage mass flow.Finally,a deeper investigation of the tip leakage flow is
proposed.