bluuebunny/crossref_metadata_2025 huggingface dataset:
The Effect of Speed Ratio on the First Rotating Stall Stage in Contra-Rotating Compressor
Authors: Gao Limin , Li Xiaojun , Xie Jian , Liu Bo
Published: July 2013

Since the structure and aerodynamic advantages of contra-rotating technology, it has been considered as an important approach to further improve the thrust-weight ratio of aircraft engine. In the present work, the impact of rotating speed ratio on the first rotating stall stage of a CRAC which consist of two counter-rotating rotors is investigated numerically. To detect the stall margin of CRAC exactly, the back pressure dichotomy method is developed, the grid indepence is verified and the performance is measured. A large number computation is carried out to explore the influence of rotating speed ratio on the performance of contra-rotating compressor. Finally the flow filed near blade tip is analyzed to find the relation between the rotating speed ratio and the first stall rotor. The result shows: (1) The work conditions of ROT1 have a significant impact on the aerodynamic performance of ROT2, while ROT2 play a little impact on the performance of ROT1. (2) At the condition of rotating speed ratio R2:R1≥0.9, the second rotor will be the first stall stage as the mass flow is decreased. (3) When the ROT2 rotating speed is slower than the ROT1, the intensity of tip leakage in ROT2 declines obviously with decreasing the rotating speed ratio, but the intensity of tip leakage in ROT1 has little changes. At the condition of R2:R1<0.9, the first rotor will be the first stall stage.
Secondary Flow Loss Reduction Through Blowing for a High-Lift Front-Loaded Low Pressure Turbine Cascade
Authors: Benton Stuart , Bons Jeffrey P. , Sondergaard Rolf
Published: July 2013

Efforts to increase individual blade loading in the low pressure turbine have resulted in blade geometries optimized for midspan performance. Many researchers have shown that increased blade loading and a front-loaded pressure distribution each contribute separately to increased losses in the endwall region. A detailed investigation is performed of the baseline endwall flow of the L2F profile, a high-lift, front loaded profile. In-plane velocity vectors and total pressure loss maps are obtained in five planes oriented normal to the blade surface, for three Reynolds numbers. A row of pitched and skewed jets are introduced near the endwall on the suction surface of the blade. The flow control method is evaluated for four momentum coefficients at the high Reynolds number, with a maximum reduction of 42% in the area averaged total pressure loss coefficient. The same blade is also fitted with midspan vortex-generator jets and is tested at a Reynolds number of 20,000, resulting in a 21% reduction in area averaged total pressure loss.
Hub Extension in an Axial Gas Turbine Diffuser
Authors: Hirschmann A. , Volkmer S. , Casey M. , Montgomery M.
Published: July 2013

Diffusers downstream of axial gas turbines typically support a hub carrying the turbine rotor bearing. The blunt end of the hub is one of the main causes of loss production. However, a long hub can theoretically reduce the losses and improve the pressure recovery of the diffuser and thus the overall performance of the gas turbine. At the end of the hub the flow experiences a sudden expansion, which can be described by the Borda Carnot equation for incompressible fluid. The pressure loss at the end of the hub depends on the dynamic head and the area ratio of the expansion. With a long hub the blunt end moves further downstream into the diffuser and therefore the flow velocity and the area ratio of the step are reduced, resulting in a lower pressure loss.

Two different casing geometries, each with a short and a long hub configuration are examined experimentally and numerically using CFD (Computational Fluid Dynamics). The geometrical configurations lead to a separation at the casing for a uniform total pressure profile at the inlet which shows the highly loaded state of the diffuser configuration. Experiments have been conducted at a high subsonic inlet Mach number comparable to real turbine exit flows and a low Mach number representing incompressible flow. It is shown that in some cases the long hub has a benefit to pressure recovery. However, the Mach number at the inlet of the diffuser influences the intensity of this effect. A high inlet Mach number increases the losses in the inlet section of the diffuser and at the struts supporting the hub and hence decreases the diffuser performance. Consequently the performance depending on the particular diffuser design can be decreased with an extended hub configuration.
Aerodynamic Investigation of a Nozzle Clearance Effect on Radial Turbine Performance
Authors: Roumeas M. , Cros S.
Published: July 2013

Liebherr Aerospace designs and develops cooling air systems, with notably axial fan and radial turbine or compressor. The development of new architectures (especially electrical system) now requires improving the turbine performances on the whole operating range. To reach that industrial request, a variable nozzle area can be used, performed by changing the nozzle blade angle for a given blade height. For the blades to be moveable, tip and hub clearances must be present (and thus modeled) in the nozzle. The impact of that clearance on the turbine performance, and moreover on the flow field, is studied here by 3D numerical way. The clearance mass flow leads to the development of a marginal vortex along the nozzle blade chord that tends to increase the total pressure loss in the nozzle on the one hand, and to modify the flow angle on the rotor inlet on the other hand. The vortex development induces an efficiency loss that must be taken into account during the design.
On LES Methods Applied to Seal Geometries
Authors: Tyacke James , Jefferson-Loveday Richard , Tucker Paul
Published: July 2013

Nine Large Eddy Simulation (LES) methods are used to simulate flow through two labyrinth seal geometries and are compared with a wide range of Reynolds-Averaged Navier-Stokes (RANS) solutions. These involve one-equation, two-equation and Reynolds Stress RANS models. Also applied are linear and nonlinear pure LES models, hybrid RANS-Numerical-LES (RANS-NLES) and Numerical-LES (NLES). RANS is found to have a maximum error and a scatter of 20%. A similar level of scatter is also found among the same turbulence model implemented in different codes. In a design context, this makes RANS unusable as a final solution. Results show that LES and RANS-NLES is capable of accurately predicting flow behaviour of two seals with a scatter of less than 5%. The complex flow physics gives rise to both laminar and turbulent zones making most LES models inappropriate. Nonetheless, this is found to have minimal tangible results impact. In accord with experimental observations, the ability of LES to find multiple solutions due to solution non-uniqueness is also observed.
Applications of EARSM Turbulence Models to Internal Flows
Authors: Mehdizadeh O. Z. , Temmerman L. , Tartinville B. , Hirsch Ch.
Published: July 2013

Turbulence modeling remains an active CFD development front for turbomachinery as well as for general industrial applications. While DNS and even LES still seem out of reach within the typical industrial design cycle due to their high computational cost, RANS-based models remain the workhorse of CFD. Currently, the most widely used models are Linear Eddy-Viscosity Models (LEVM), despite their known limitations for certain flow complexities. Therefore, extending the reliability of eddy-viscosity models to more complex flows without significantly increasing the computational cost can immediately contribute to more reliable CFD results for wider range of applications. This, in turn, can further reduce the need for costly tests and consequently can reduce the product development cost.

A promising approach to achieve this goal is using Explicit Algebraic Reynolds Stress Models (EARSM), obtained through a simplification of the full Differential Reynolds Stress Models (DRSM), and can be perceived as an extension of LEVMs by including the non-linear relation between the turbulence stress tensor, the mean-flow gradient and the turbulence scales. These models are thus less demanding than DRSM, yet capable of capturing more complex turbulence features, compared to LEVM, such as anisotropy in the normal stresses. This may be particularly important in corner flows, for instance, in the hub-blade regions or in diffusers.

This work explores the application of EARSM models to a double diffuser and a high-performance centrifugal compressor stage (HPCC). The results are compared to available experimental data [1,2] showing the importance of including the anisotropy of turbulence in the model, particularly in presence of turbulent corner flows in a diffuser. Furthermore, the EARSM results are also compared to results from the commonly used SST turbulence model. The CFD comparison includes details of the flow structure in the diffuser, where the most noticeable impact from the use of EARSM turbulence models is expected.
Surface Thin Film Gauge Measurements in a Two-Stage Low Pressure Turbine at Low Reynolds Number
Authors: Kürner Matthias , Rose Martin G. , Staudacher Stephan , Gier Jochen , Fiala Andreas , Patzer Bastian
Published: July 2013

A two-stage low pressure axial turbine has been tested in cooperation between the Institute of Aircraft Propulsion Systems (ILA) and MTU Aero Engines GmbH (MTU). The experimental results taken in the Altitude Test Facility are used to assess blade row performance of vane 2 at mid height over a range of Reynolds numbers from as low as 35,000 up to 88,000. Both Mach and Reynolds similarity are preserved. Surface thin film gauges at midspan on vane 2 suction side are used to analyse the unsteady behaviour of the boundary layer. Unsteady data from area traverses downstream of vane 2 using X-hotfilm probes complement the analysis describing the unsteady wake evolution at mid height. The nature of the unsteady transitional low Reynolds number boundary layer is discussed.
The Effect of Tip Leakage Vortex for Operating Range Enhancement of Centrifugal Compressor
Authors: Tomita Isao , Ibaraki Seiichi , Furukawa Masato , Yamada Kazutoyo
Published: July 2013

Recently, the application of turbochargers is increasing because they are effective in improving fuel consumption of engines. One of the most important turbocharger characteristics is compressor operating range, since it has been used in various driving patterns with the advent of variable geometry turbochargers. Owing to the complicated phenomena such as rotating stall occurring at low flow rate condition, flow analysis is very difficult and details of flow structure have not been fully understood for a long time since the early 1970’s.

In this study, two compressors with different operating range width were investigated with experimental and computational flow analysis. In the compressor with narrow operating range, the amplitude of blade passing pressure fluctuation decreases rapidly and rotating stall occurs near surging. On the other hand, in the compressor with wide operating range, the blockage by the tip leakage vortex breakdown play a role in stabilizing the flow filed and keeping up a high performance even at low flow rates.
Aerodynamic Effects of an Unshrouded Low Pressure Turbine on a Low Aspect Ratio Exit Guide Vane
Authors: Selic Thorsten , Lengani Davide , Marn Andreas , Heitmeir Franz
Published: July 2013

This paper presents the effects of an unshrouded low pressure turbine (LPT) onto the following exit guide vane row (EGV). The measurement results were obtained in the subsonic test turbine facility at Graz University of Technology by means of a fast response pressure probe in planes downstream of the rotor as well as oil flow visualisation. The test rig was designed in cooperation with MTU Aero Engines and represents the last 1.5 stages of a commercial aero engine. Considerable efforts were put into the adjustment of all relevant model parameters to reproduce the full scale LPT situation.

Different tip clearances were evaluated by means of CFD obtained using a commercial Navier-Stokes code and validated with experimental results. The goal is to evaluate the effect of the varying leakage flow on the flow in the low aspect ratio EGV. Special attention is given to the impact on the development of secondary flows as well as the flow structures downstream of the EGV. The effect of the leakage flow causes a change of the flow structure of the EGV, particularly losses. Considering the largest investigated tip-clearance, the losses increased by 71% when compared to a zero-leakage case.
Experimental and Numerical Study of a Subsonic Aspirated Cascade
Authors: Godard Antoine , Bario François , Burguburu Stéphane , Lebœuf Francis
Published: July 2013

This paper presents the validation of a design method for aspirated compressor blades, combining a passive separation control by blade shaping with an active flow control by aspiration.

In a first part, a linear aspirated cascade designed according to this method was built and tested at low speed, without and with aspiration. The latter was only applied on the suction surfaces of the blades. Particle Image Velocimetry measurements performed at mid-span of the cascade, in the central passage, showed a complete reattachment of the separated boundary layer on the suction side of the blade. A flow deflection of approximately 65 degrees was achieved requiring an aspirated mass flow rate of 3.3%. However, boundary layer reattachment is effective in a zone centered at mid-span covering 30% of blade span. Flow visualization revealed large corner separation in the presence of aspiration. This is due to the re-establishment of strong pressure gradient on sidewalls of the cascade. No flow control was applied on these zones for optical access purpose. These secondary-flow regions reduced the diffusion occurring within the cascade by nearly 60% in comparison with the design intent. They also increased the expected level of total pressure losses measured by wake traverses downstream of the cascade.

In a second part, numerical simulations of the aforementioned experiment were carried out to help the understanding of the experimental results. The simulations were able to reproduce correctly the characteristic flow features, without and with aspiration, observed and measured during the experiment. Thus, they confirmed the potential of this design method developed for aspirated compressor blades, as well as CFD capabilities to simulate the influence of technological effects like suction slots. A uniform and a non-uniform aspiration distribution along the blade span direction were considered during simulations. Suction distribution was found to have a significant impact on the control by aspiration. This design feature, in addition to flow control on endwalls, has to be taken into account in the three-dimensional design of highly loaded aspirated compressor blades.
Analysis of Combustor/Vane Interaction With Decoupled and Loosely Coupled Approaches
Authors: Salvadori Simone , Riccio Giovanni , Insinna Massimiliano , Martelli Francesco
Published: July 2013

Numerical techniques are commonly used during both design and analysis processes, mainly considering separated components. Technological progress asks for advanced approaches that allow to analysing the interaction between the components, especially when considering combustor/turbine interaction. Hot spots and inlet swirl profiles generated by the combustor have been demonstrated to affect high-pressure turbine performances and reliability.

This work deals with the investigation of the effects of realistic boundary conditions for the high-pressure turbine vane, also proposing an approach for coupled simulation of the combustor/vane interaction. The method consists in a loosely coupled approach for the data exchange on the combustor/vane interface section. Data from the combustor exit section (stagnation conditions, velocity profile and turbulent quantities) are provided to the vane inlet and vice versa (for the static pressure). The proposed method is applied to a test case consisting of a redesigned combustor and the vane of the MT1 test case from QinetiQ.

A preliminary analysis was dedicated to define the combustor geometry and the operating conditions. Then, the MT1 working conditions have been rescaled and coupled with the combustor, maintaining the stage geometry and the experimental non-dimensional parameters. Second order accurate steady simulations were performed for both combustor and high-pressure turbine vane.

Calculations with a uniform profile and a theoretical nonuniform inlet profile (deriving from the EU funded TATEF2 project) have been considered as representative of commonly used approaches. The results obtained for the stator in terms of isentropic Mach number and Nusselt number along blades surfaces and inner end-wall are compared with each other and with the available experimental data.

Due to the large dimension of computational grids a parallel approach was applied. The activity was carried out using the IBM PLX supercomputer in the frame of the FrUIT project supported by CINECA.
Reconstruction of the Unsteady Pressure Field in a Low Speed Linear Cascade
Authors: Förster Felix , Sims-Williams David , Ingram Grant
Published: July 2013

This paper describes investigations into unsteady flow phenomena in a high turning linear cascade with a nominal steady flow. A fast response five hole probe was used to obtain time-accurate measurements. Since a single probe was used to traverse an entire measurement plane, the results were obtained as discrete points and were not synchronous. One method used to address this issue is to traverse the measurement plane point by point and record the signal of a reference probe simultaneously. Synchronization can be performed during post processing by analyzing the reference signal for a common periodic component and phase-shifting all signals with respect to the reference phase of that flow feature.

This paper presents results from the application this unsteady reconstruction technique to the so-called Durham Cascade 28% downstream of the trailing edge in order to provide insight into the nature of the dominant unsteady mode. It was found that there was significant temporal variation in the strength of the passage vortex but the position of the vortex remains fixed.
Design Optimization and Experimental Verification of Centrifugal Compressors With Curvilinear Element Blades
Authors: Sugimura Kazuyuki , Kobayashi Hiromi , Nishida Hideo
Published: July 2013

This paper describes the optimal design and experimental verification of centrifugal compressors with leaned curvilinear element blades. The design targets were a fully shrouded centrifugal impeller and a low-solidity vaned diffuser.

First, a new method of defining the curvilinear element blade was developed for centrifugal turbomachinery using coordinate transformations between a revolutionary flow-surface coordinate system and a cylindrical coordinate system. All the blade sections in the transferred cylindrical coordinate system were moved and stacked spanwise to form a new leaned blade surface. The inverse transformation results in a curvilinear element blade in the original coordinate system. The direction of movement of the blade section could be any of the existing definitions such as sweep, dihedral, or tangential lean. For simplicity, we have defined a “lean profile” as a general expression for the spanwise distribution profile of movement of the blade sections.

Model compressors with curvilinear element blades were then designed at a suction flow coefficient of 0.073. Optimal lean profiles for the impeller and vaned diffuser that maximized adiabatic efficiency and uniformity of outflow were explored using a multi-objective genetic algorithm, Kriging surrogate model, and steady Reynolds-averaged Navier Stokes simulations. We chose efficiency-weighted solutions since the efficiency and uniformity had a positive correlation. A sensitivity analysis showed that tangential leans near the endwalls are keys for improving the impeller’s efficiency. The chosen optimal impeller had a concave blade suction surface and a concave leading edge. Although clear patterns in geometrical features for optimal diffuser vanes could not be captured, we found that the dihedral profile had a predominant effect on efficiency.

The model compressors was experimentally measured and compared with traditional compressors as to their aerodynamic performance. The experimental apparatus was composed of a suction nozzle, impeller, diffuser, and return channel. The results demonstrated that the models’ adiabatic efficiency was higher by 1.2–1.4% at the design point. Although the stall margins slightly decreased (by 0.7%) in the experiments, the surge margins expanded (by 3.7–6.7%).
Corner Separation Control in a Highly Loaded Compressor Cascade Using Plasma Aerodynamic Actuation
Authors: Wu Yun , Zhao Xiao-hu , Li Ying-hong , Li Jun
Published: July 2013

Corner separation, which forms over the suction surface and endwall corner of a blade passage, causes significant total pressure loss in highly loaded compressors. Plasma flow control, based on the plasma aerodynamic actuation, is a novel active flow control technique to improve aircrafts’ aerodynamic characteristics and propulsion efficiency. This paper reports computational and experimental results on using three types of plasma aerodynamic actuation (PAA) to control the corner separation in a highly loaded, low speed, linear compressor cascade. Reynolds-Averaged Navier-Stokes simulations were performed to optimize the PAA arrangement. The PAA was generated by a microsecond or nanosecond dielectric barrier discharge in wind tunnel experiments. The total pressure loss coefficient distribution was adopted to evaluate the corner separation control effect. The control effect of pitch-wise PAA on the endwall, in terms of relative reduction of the pitch-wise averaged total pressure loss coefficient in the wake, is much better than that of stream-wise PAA on the suction surface. When both pitch-wise PAA on the endwall and stream-wise PAA on the suction surface are turned on simultaneously, the control effect is the best among all three types of PAA. The main effect of pitch-wise PAA on the endwall is to inhibit the crossflow from neighboring pressure surface to the suction surface, whilest the main effect of stream-wise PAA on the suction surface is to inhibit the boundary layer accumulation and separation. Compared to microsecond discharge PAA, nanosecond discharge PAA is more effective at higher freestream velocity. The mechanisms for nanosecond discharge and microsecond discharge PAA are different for corner separation control.
Experimentally Observed Unsteady Work at Inlet to and Exit From an Axial Flow Turbine Rotor
Authors: Rose Martin G. , Jenny Philipp , Gier Jochen , Abhari Reza S.
Published: July 2013

World literature has introduced the aerodynamic importance of unsteadiness in turbines. In particular the unsteady static pressure field determines the work of the machine. The unsteadiness can redistribute the total pressure in a cascade with wake interaction. It has been shown that differences in work between wake and free-stream can act to rectify the wakes and boost efficiency. In this paper FRAP (Fast Response Aerodynamic Probe) data is used to study the nature of the unsteady work in the flow at entry to and exit from a rotating turbine blade. The topic is addressed experimentally, theoretically and computationally. It is found at both rotor inlet and exit that upstream wakes influence the unsteady work distribution. The relationship between the unsteady work in the absolute frame, the relative frame and the momentum of the fluid circumferentially is derived and verified experimentally. Computational results (URANS) are compared to the experimental results: reasonable agreement is found at rotor exit but significant differences at rotor inlet are found. The CFD has failed to capture the von Karman vortices and has dramatically lower levels of unsteady work. The experimental unsteady work distribution suggests possible effects of wake bending and vortex instability.
The Impact of Leakage Flow Tangential Velocity on Secondary Losses in a Shrouded Compressor Cascade
Authors: Kim J. W. , Lee J. S. , Song S. J. , Kim T. , Shin H-. W.
Published: July 2013

Experimental and numerical studies have been performed to investigate the effects of the leakage flow tangential velocity on the secondary flow and aerodynamic loss in an axial compressor cascade with a labyrinth seal. Six selected leakage flow tangential (vy/Uhub = 0.15, 0.25, 0.35, 0.45, 0.55 and 0.65) have been tested. In addition to the classical “secondary” flow, shroud trailing edge vortex and shroud leading edge vortex are examined. The overall loss decreases with increasing leakage flow tangential velocity. Increased leakage flow tangential velocity underturns the hub endwall flows through the blade passage, weakening the suction side hub corner separation. Due to the suction effect of the downstream cavity, increasing leakage flow tangential velocity weakens the shroud trailing edge vortex. Also, increasing leakage flow tangential velocity strengthens the shroud leading edge vortex, weakening the pressure side leg of the horseshoe vortex, and, in turn, the passage vortex. Thus, the overall loss is reduced with increasing leakage flow tangential velocity.
Experimental Investigation on Tip Clearance Flow of Compressor Cascade With Blade Tip Winglet
Authors: Han Shaobing , Zhong Jingjun , Lu Huawei , Kan Xiaoxu , Gao Haiyang
Published: July 2013

This paper presents the results of experimental research of flow in an axial compressor cascade with different types of winglet on the blade tip, which consists of a suction-side winglet, pressure-side winglet and the combined winglet. The detailed tip leakage flow field with different winglet was described with total pressure loss coefficient, secondary streamline and axial vorticity on the cascade exit flow field. The mechanisms of the three passive control methods were illuminated. The result indicated that the tip clearance flow strengths could be reduced in all the three control strategies. The compressor aerodynamic performance could be improved via the addition of tip winglets. The suction-side winglet had the best effect on the cascade flow field, and the strength of leakage vortex and the associated mixture losses were reduced.
Multi Disciplinary Design Optimization and Performance Evaluation of a Single-Stage Transonic Axial Compressor
Authors: Kang Young-Seok , Park Tae-Choon , Yang Soo-Seok , Lee Sae-Il , Lee Dong-Ho
Published: July 2013

The multi-disciplinary optimization (MDO) method, which integrates aerodynamic performance and structural stability, was utilized in the development of a single-stage transonic axial compressor. Numerical simulations and compressor tests were also carried out to evaluate the aerodynamic performance and safety factor of the optimized compressor. The rotor has 60 design parameters with twelve most sensitive design variables selected for design optimization. The stator was redesigned according to the rotor outlet flow angle variation to match the stator incidence angle by −1∼0 degrees, while maintaining the stage outlet flow angle. The design goal is to maximize both the stage efficiency and the safety factor from the baseline scratch compressor design. The object function is composed of the normalized efficiency and safety factor with weight factors. Initially, an approximation model was created to search for the global optimization within given ranges of variables and considering several design constraints. The genetic algorithm was used to explore the Pareto front of the optimization to find the maximum objective function values. The final design was chosen after a second stage gradient-based optimization process to improve the accuracy of the optimization.

The CFD results showed that more blade loading is burden to the hub region by increasing the incidence angle. The fore part blade loading gradually decreases along the span-wise direction. In addition, normal shock, which spreads along the hub to the blade tip, is confined in the rotor flow passage and pressure surface shock coincidence point moves to be closer to the blade leading edge, indicating an increase in the amount of blade loading. FEA analyses showed that the blade root stress has been drastically relieved, because the optimized blade has trapezoid-shaped hub design relative to the baseline design. The final design achieved efficiency gain of 3.69% and showed a higher safety factor by 2.3 times relative to the baseline model, while maintaining its stage mass flow rate and total-to-total pressure within the design constraints. The compressor performance test data showed good agreement with the optimized design and CFD results. However, there is room for improvement in the optimization process to reflect off-design performance so as to secure more stable compressor operation ranges.
An Evaluation of Mixing Effect on the Recycle Stage of Syngas Compressor Based on Numerical and Experimental Investigations
Authors: Ding Liang , Wang Tong , Fu Yao , Yang Bo , Gu Chuangang
Published: July 2013

With the requirement of the chemical synthetic process, the reaction pressure of the synthesis gas is over 14MPa. Meanwhile it is necessary to employ a recycle stage to improve the synthesis efficiency. The performance of the recycle stage would be influenced by the upstream from the previous stage and the complementary gas. However, it is difficult to understand the performance of the recycle section separately because of the lack of the testing data in practice. In fact, the inlet flow distortions caused by the two streams mixing process may result in significant deterioration in compressor performance. To evaluate the impact of the mixing process on the recycle stage, a 3D numerical flow passage model consisting of last stage of high pressure compressor (HPC) section, the mixing chamber and recycle stage, has been established in the paper. True thermodynamic properties of the mixture are considered for the numerical analyses at high pressure and the results are validated with the test performance data of an onsite synthesis gas compressor. In the mixing chamber, the flow directions, pressures, temperatures and velocities of the two upstreams are different, which result in the non-uniform distribution of the flow along radial direction at the inlet of recycle stage. From the numerical results, eight different types of losses are calculated to understand the performance of the flow passage. It is found that the loss of the mixing process and blade loading loss are the two main losses. Further, the uniform conditions and mixing inlet conditions are applied separately at the impeller inlet position to evaluate the recycle stage performance. Compared to the uniform inlet, under the mixing inlet condition the recycle stage pressure ratio and efficiency decreases, and the operating range becomes narrower. In order to understand this mechanism, a fan test rig is setup with the mixing inlet structure similar to the recycle stage. The performance of the fan with the mixing inlet structure is numerically investigated. Both of the testing and numerical results illustrate the deterioration of the performance with the mixing inlet structure. The effect of the different mixture ratios on the fan performance is evaluated. The results suggest that the best efficiency point of the fan is achieved at the flow ratio (minlet2/minlet1) of 5.
Numerical Investigation of the Effect of Rotor Blade Leading Edge Geometry on the Performance of a Variable Geometry Turbine
Authors: Huang Lei , Zhuge Weilin , Zhang Yangjun , Hu Liaoping , Yang Di , Wang Jiawei
Published: July 2013

Variable geometry turbines are more and more widely used in diesel engines to meet the requirements of the stringent emission standards. The VGTs mostly operate at off-design conditions. At highly off-design conditions, there exist complex secondary flow structures and severe flow separation in the rotor passage, which deteriorate the turbine performance largely.

The influence of rotor blade leading edge geometries on the VGT performance was studied by CFD simulations. The blade angle distribution along the leading edge was varied while keeping the radial-fiber rotor construction. The effects of inlet sweep angle distribution and lean angle of the blade leading edge on the turbine flow fields and performance were investigated under different operating conditions.

Results show that the turbine with backswept leading edge has better performance at low U/C, while the turbine with forward swept leading edge has a higher efficiency under high flow rate conditions. With the same sweep angle distribution, the leading edge lean affects the flow fields in the rotor passage as well as the turbine performance significantly. The influence of blade lean on the turbine performance varies according to different swept blading and operating conditions.
Effects of Using Different Arrangements of Fanlets on Performance of Twisted Axial Aerial Fans
Authors: Abbaszadeh Morteza , Shahrooz Mina , Farhanieh Bijan
Published: July 2013

On the ground that axial turbo machines are used abundantly in propulsion, air conditioning, power generating and other systems, studying their performances to find ways to optimize them is of immense importance. In this research, effects of using different arrangements of fanlets planted on twisted axial aerial fans have been scrutinized, in order to minimize repercussions of the tip vortices on the fan blade. So a basic blade was designed in keeping with blade element theory to generate 2 and 4-blade fan models. Then each model was separately tested in three arrangements of: Up let, Down let and Fresh i.e. without fanlet. K.e turbulent modeling was employed to analyze designed models. In addition, Horizontal Element Separation known as H.E.S method was employed to plot distribution profiles of the aerodynamic forces acting on the blade to blade span as a post processing tool for Computational Fluid Dynamics (C.F.D.) tests. Convergence of C.F.D. outcome with results of the Blade Element Method for a two-blade fresh model proved C.F.D. method usage validity. Concluding results obviously demonstrate that performance of an axial aerial fan could be enhanced by maximum 5 per cent using fanlets. Studying and analyzing distribution profiles reveals why and how fanlet affects performance of axial fans.
Experimental Investigation and Computational Evaluation of Contoured Endwall and Leading Edge Fillet Configurations in a Turbine NGV
Authors: Turgut Özhan H. , Camcı Cengiz
Published: July 2013

Secondary flow minimization is a crucial problem in a turbine passage. In the present paper, three different ways are employed to reduce the secondary flow related total pressure loss. These are nonaxisymmetric endwall contouring, leading edge (LE) fillet, and the combination of these two approaches. Experimental investigation and computational assessment are applied for the performance calculations. The experiments are carried out in an annular Axial Flow Turbine Research Facility (AFTRF) having a diameter of 91.66cm. For the experimental measurement comparison, a reference Flat Insert is installed in the nozzle guide vane (NGV) passage. It has a constant thickness with cylindrical surface and is manufactured by stereolithography (SLA) method. Also, Flat Insert has a backward facing step at the NGV exit, and the effect of this step is analyzed computationally. Four different LE fillets are designed, and they are attached to both cylindrical Flat Insert and the contoured endwall. Total pressure measurements are taken at rotor inlet plane with Kiel probe. The probe traversing is completed with one vane pitch and from 8% to 38% span. For one of the designs, area-averaged loss is reduced by 15.06%. The simulation estimated this reduction as 6.95%. Computational evaluation is also performed at the NGV exit plane. The most effective design reduced the mass-averaged loss by 1.84% on the whole passage. The computational study did not include the rim seal flow between the vane and rotor domain and also rotor simulation was absent. The difference between the measurements and the simulation comes from these two important effects.
Computational Validation of the Flow Through a Turbine Stage and the Effects of Rim Seal Cavity Leakage on Secondary Flows
Authors: Turgut Özhan H. , Camcı Cengiz
Published: July 2013

A computational validation study related to aerodynamic loss generation mechanisms is performed in an axial flow turbine. The 91.66 cm diameter axial flow turbine research facility has a stationary nozzle guide vane assembly and a 29 bladed HP turbine rotor. The NGV inlet and exit Reynolds numbers based on midspan axial chord are around 300000 and 900000, respectively. GRIDPRO is used as the structured grid generator. y+ values are kept below unity. The finite-volume flow solver ANSYS CFX with SST k–ω turbulence model together with the transitional flow model is employed. Experimental flow conditions are imposed at the boundaries. The computational predictions are compared to experimental data at NGV exit plane and rotor inlet plane. NGV exit plane measurements come from a previous experimental study with a five-hole probe and the data at rotor inlet plane is taken by the current authors using a Kiel probe with 3.175mm head diameter. The comparison of rotor-stator interface models shows that the stage model, which calculates the circumferentially averaged fluxes and uses as the boundary condition at the interface plane, agrees well with the experimental total pressure coefficient data at the NGV exit. The difference between the NGV only simulation and the rotor-stator simulation is emphasized. The effect of rim seal flow on the mainstream aerodynamics is investigated. This validation study shows that the effect of future geometrical modifications on the endwalls and the vane will be predicted reasonably accurately.
Casing Treatment and Inlet Swirl of Centrifugal Compressors
Authors: Chen Hua , Lei Vai-Man
Published: July 2013

Ported shroud is a cost effective casing treatment that can greatly improve stability of centrifugal compressors. It is widely used in turbochargers and other applications where compressors with wide flow range are required. This paper reviews the development of the ported shroud concept from its first conception in the 1980s to its current various configurations, explores the underline mechanisms that deliver the performance improvement. It is explained that by removing stagnant fluid from impeller inducer shroud end wall boundary-layer region and recirculating it to the impeller inlet, blade loading near the inducer shroud is increased with improved inlet suction. For transonic flow, ported shroud weakens the shock wave and reduces flow separation on the inducer suction surface. It is argued that the effectiveness of the ported shroud is a balance of blade loading and the flow loss inside the ported shroud cavity. The loss needs to be minimised if ported shroud is to be more effective. Blade loading may be increased by various methods such as using high inducer blade turning and using full bladed impellers. The blade loading can also be improved by removing flow swirl in ported shroud flow by vanes, or imposing negative swirl by vanes in ported shroud. Circumferential flow variation caused by volute housing can be taken into account by variable pitch vanes or by variable port position.
Influence of Mach Number and Aerodynamic Loading on Rotating Instability in an Annular Compressor Cascade
Authors: Beselt Christian , Peitsch Dieter
Published: July 2013

An experimental study was conducted to investigate the influence of Mach number and aerodynamic loading on Rotating Instability (RI). Investigations were performed on an annular compressor cascade with radial clearance at the hub. In addition to steady flow field measurements, time resolved measurements of static wall-pressure at the hub, close to the tip clearance, were performed. For a specific range of incidence angles, RI could be detected at all flow speeds (Ma = 0.2–0.7). Shortly before the inception of RI, a sudden increase in pressure loss coefficient and outflow angle was detected by steady flow field measurements. The time resolved measurements showed an abruptly broadband increase of amplitude in the frequency range, in which RI occurs at higher loadings. RI propagates with 25–33% of the circumferential component of the upstream flow speed. Furthermore it is possible to scale the power spectral density function of the unsteady data with Strouhal number.
A Path Towards the Aerodynamic Robust Design of Low Pressure Turbines
Authors: Bertini Francesco , Credi Martina , Marconcini Michele , Giovannini Matteo
Published: July 2013

Airline companies are continuously demanding lower-fuel-consuming engines and this leads to investigating innovative configurations and to further improving single module performance. In this framework the Low Pressure Turbine (LPT) is known to be a key component since it has a major effect on specific fuel consumption (SFC).

Modern aerodynamic design of LPTs for civil aircraft engines has reached high levels of quality, but new engine data, after first engine tests, often cannot achieve the expected performance. Further work on the modules is usually required, with additional costs and time spent to reach the quality level needed to enter in service. The reported study is aimed at understanding some of the causes for this deficit and how to solve some of the highlighted problems.

In a real engine, the LPT module works under conditions which differ from those described in the analyzed numerical model: the definition of the geometry cannot be so accurate, a priori unknown values for boundary conditions data are often assumed, complex physical phenomena are seldom taken into account, operating cycle may differ from the design intent due to a non-optimal coupling with other engine components. Moreover, variations are present among different engines of the same family, manufacturing defects increase the uncertainty and, finally, deterioration of the components occurs during service.

Research projects and several studies carried out by the authors lead to the conclusion that being able to design a module whose performance is less sensitive to variations (Robust LPT) brings advantages not only when the engine performs under strong off-design conditions but also, due to the abovementioned unknowns, near the design point as well.

Concept and Preliminary Design phases are herein considered, highlighting the results arising from sensibility studies and their impact on the final designed robust configuration. Module performance is afterward estimated using a statistical approach.
Experimental Investigation on the Performance of a Compressor Airfoil in Droplet Laden Flow
Authors: Ober Birger , Joos Franz
Published: July 2013

One way to improve performance and efficiency of gas turbines is to implement a wet compression process in the gas turbines’ axial compressor. Water droplet laden compressor flows promise benefits in efficiency and to some extent in performance and stability. A promising approach is the stabilizing influence on highly stressed airfoils as numerically predicted by Sun et al. [1]. First systematic experimental investigations have been made [2], but were restricted to a limited number of test cases. The ongoing experimental investigation focuses on the influence of a droplet laden flow on an axial compressors’ aerodynamics over the entire range of relevant incident flow angles. The result is a comparison of a dry air flow and a droplet laden air flow with respect to values of loss coefficient and the range of incident angles which result in a stable compressor flow.
On the Use of Synthetic Jets in Transonic Compressors
Authors: Biollo R. , Staffieri E. , Benini E.
Published: July 2013

Several fluidic-based methods have been investigated by compressor designers with the aim of improving the aerodynamic behavior of compressor blades. Recently, synthetic jets have been proposed as a promising fluidic-based solution for reducing flow separation inside turbomachines. With the characteristic of zero net mass flux and non-zero momentum flux, synthetic jets could be very effective because at the suction part of the cycle the low-momentum fluid is sucked into the device, whereas in the blowing part a high momentum jet accelerates it. However, to the authors’ knowledge, the use of synthetic jets has never been experimented in transonic compressors. In these machines, synthetic jets could have a great impact on the thickness and stability of the blade suction side boundary layer arising after the interaction with the shock. With the aim to investigate that, a rake of 600 Hz sinusoidal wave form synthetic jets (with no radial velocity component) positioned into a thin radial slot on the blade suction side of the NASA Rotor 37 has been modeled and simulated. Simulations have been carried out using a validated 3D CFD (Computational Fluid Dynamics) unsteady model.

The study gave a clear understanding of positive and negative impacts of the simulated device on the local flow field during the entire jets working cycle. The shock structure and boundary layer resulted significantly affected in the proximity of the synthetic jets, with different effects depending on slot location and jet channel angle. The local perturbation then affected the overall flowfield with impacts on the overall rotor performances. Despite the reported cases were selected for the purpose of the analysis only, and not for the aerodynamic enhancement of the rotor, slight benefits in terms of overall performance were observed for almost all the cases.
Influence of Inlet Guide Vane Wakes on the Passage Flow in a Transonic Axial Compressor
Authors: Leichtfuss S. , Biela C. , Schiffer H.-P. , Heinichen F.
Published: July 2013

Variable inlet guide vanes enhance the efficiency and stability of modern transonic compressors. The current quest for compact and highly efficient aero-engines requires for higher stage-loading and small axial gaps between adjacent blade rows, increasing interaction between blade rows and introducing further unsteadiness into the flow. Modeling these interactions is relevant to jet engine implementation.

Recent advancements in numerical simulation of unsteady flow in multiple blade rows create an additional way to investigate the flow patterns formed by the unsteady interaction. In the case of transonic compressors, the experimental and numerical database is small. The current article will contribute to this database by investigating the flow downstream of an inlet guide vane and the influence this flow has on the passage in a transonic compressor.

Unsteady flow phenomena are resolved by piezoresistive wall pressure tappings. Numerical and experimental data show that the interaction of blade rows influences the formation of the tip leakage vortex. Analyzing the vortex structure within the transonic compressor stage it is possible to show how the blade rows interact and how geometric modifications in the VIGV tip gap model influence simulation results.