bluuebunny/crossref_metadata_2025 huggingface dataset:
Three Dimensional Aerodynamic Optimization for an Ultra-Low Aspect Ratio Transonic Turbine Stator Blade
Authors: Hasenja¨ger Martina , Sendhoff Bernhard , Sonoda Toyotaka , Arima Toshiyuki
Published: November 2008

A modern numerical stochastic optimization method, namely the evolution strategy (ES), was applied to an ultra-low aspect ratio transonic turbine stator blade in order to seek a new aerodynamic design concept for lower secondary flow losses. The low stator blade count is selected to avoid the direct viscous interaction of the stator wake with the downstream rotor blade. This led to the ultra-low aspect ratio stator blade. In the optimization, two kinds of objective functions were used, that is, (1) minimization of the “aerodynamic loss” (a single objective), (2) minimization of the “aerodynamic loss” and of the “variation of circumferential static pressure distribution” downstream of the stator blade (multi-objective optimization). In the case of the single objective, the aerodynamic loss is improved by an extreme aft-loaded airfoil with a noticeable bent part near the trailing edge, although the circumferential static distribution is slightly worse than that of the baseline. In the case of the multi-objective optimization, we observe a trade-off relation between aerodynamic loss and variation of static pressure distribution which is not easily resolved. A new design concept to achieve lower aerodynamic loss for ultra-low aspect ratio transonic turbine stator blades is discussed.
Effects of Fin Shapes and Arrangements on Heat Transfer for Impingement/Effusion Cooling With Crossflow
Authors: Hong Sung Kook , Rhee Dong-Ho , Cho Hyung Hee
Published: November 2008

The present paper has investigated the effects of fin on the flow and heat/mass transfer characteristics for the impingement/effusion cooling with crossflow. The fins of circular or rectangular shape are installed between two perforated plates and the crossflow passes between these two plates. The blowing ratio is changed from 0.5 to 1.5 for a fixed jet Reynolds number of 10,000. A naphthalene sublimation method is used to obtain the local heat/mass transfer coefficients on the effusion plate. A numerical calculation is also performed to investigate the flow characteristics. Flow and heat/mass transfer characteristics are changed significantly due to installation of fins. In the injection region, wall jet spreads more widely than the case without fins because fin prevents the wall jet from being swept away by the crossflow. In the effusion region, higher heat/mass transfer coefficient is obtained due to the flow disturbance and acceleration by the fin. As the blowing ratio increases, the effects of fin against the crossflow become more significant and then the higher average heat/mass transfer coefficients are obtained. Especially, the cases with rectangular fins have about 40%∼45% enhancement at the high blowing ratio of M = 1.5. However, the increase of blockage effect gives more pressure loss in the channel.
Redesign of an 11-Stage Axial Compressor for Industrial Gas Turbine
Authors: Terauchi Koji , Kariya Daisuke , Maeda Seiichirou , Yoshiura Kenichirou
Published: November 2008

Outlined in this paper are the details of redesign of an 11-stage axial compressor for 18MW class industrial gas turbine. Although the design value of adiabatic efficiency is 85%, the efficiency achieved in the prototype engine test was 1 point short of the target. According to the test results and CFD analyses, it was estimated that the stage mismatching due to an excessive work of intermediate stages was a major cause of the losses. To improve the matching, redesign of airfoils were carried out to some blades and vanes. The results of Multistage CFD showed that the matching was improved and the efficiency was increased, which was also confirmed by the engine test.
Advances in Unsteady Turbomachinery Aerodynamics in Japan: Professor Gallus’ Contribution
Authors: Watanabe Toshinori , Outa Eisuke , Adachi Tsutomu , Ariga Ichiro , Kashiwabara Yasushige , Mizuki Shimpei , Tanaka Hideo
Published: November 2008

The present paper describes recent advances in turbomachinery aerodynamic research in Japan achieved by the researchers who studied under the supervision of the late Professor Gallus. His research work, way of thinking, and personality exerted great influence in Japan on the styles of research in the field of unsteady aerodynamics of turbomachines. The profound contribution by him to the research and development of Japanese gas turbine technologies is highly appreciated. The paper presents research results on the issues concerning cascade flutter suppression, rotating stall analysis and control for axial compressors, unsteady flow analysis and control for radial compressors and diffusers, as well as design studies of axial turbomachines.
Influence of Geometric Scaling on the Stability and Range of a Turbocharger Centrifugal Compressor
Authors: Schleer Matthias , Abhari Reza S.
Published: November 2008

In this work the stability behavior of small-scale centrifugal compressor is evaluated in detail and the influence of design features typical for small-scale applications is shown. The impeller used in this study represents the design features of machines typically used in small turbochargers like a low blade count, high blade loading and a large relative tip gap. The work is evaluating data obtained in an enlarged research facility and in an actual scale turbocharger application. Both facilities are using a geometrical similar impeller and all nondimensional parameter are maintained. The Reynolds number is adjusted by changing the inlet pressure and thus the density of the air. This setup allows measurements with high accuracy on the enlarged research stage and simple parameter studies on the small-scale model. Comparing the operating characteristics of both scales shows the validity of this approach. For the range of Reynolds numbers present, the stability of the compressor is not affected by the geometric scaling. As the user of the compressor system wishes to operate at a wide range and under varying load demands but always in stable condition the knowledge of the stability margin and the kind of instabilities is vital. An analysis of the instable phenomena limiting the range of the centrifugal compressors is shown. The analyses are interpreting the pressure fluctuations gained with high response pressure transducer located in the diffuser for the characterization of the system stability. A similar overall compressor characteristic and stability range is obtained for both scales investigated. The flow structure within the diffuser is shown in a time-resolved manner using a 3D Laser Doppler Anemometer. It is shown how the flow structure is affected by the leakage flow through the tip gap. It is forming a strong jet-wake pattern resulting in a non-uniform flow and sheared velocity triangles.
A Reduced-Order Through-Flow Program for Choked and Cooled Axial Turbines
Authors: Genrup Magnus , Carlsson Ivan , Engdar Ulf , Assadi Mohsen
Published: November 2008

A reduced-order through-flow program has been developed at the Lund Institute of Technology. The goal of the work was to develop and verify a program suitable for scientific studies of axial turbines. The program has built-in capability for multiple choked stages, and turbine cooling, as well as flexible modules for state, losses, deviation, and blockage. The code uses Matlab™ as a platform for numerics, pre- and post-processing. The turbine modules are available from the Institute free of charge for scientific use. Matlab is a commercially available mathematical package and is used as a numerical tool by many universities and companies. Today, several codes are available for turbine analysis at different levels. Most of the codes are proprietary and not available outside the companies that have developed them. There are, however, commercially available codes, but the user does not normally have accesses to the source code. This poses serious problems when such codes are used for scientific studies, and open easily modified code is indeed a desirable feature. The present paper describes in detail the calculation methods used to simulate performance of a cooled and choked turbine at off-design conditions. The algorithms necessary for finding the turbine choke point will, for example, be described in detail. The way in which the loss, deviation (including secondary and tip clearance), and blockage are included in a quasi 1-D calculation environment is also presented. The code is suitable for further development (e.g. streamline curvature through-flow), since it is based on modules for e.g. state, combustion, loss, deviation, diffuser, numerics, and in-and output data processing. It is fairly easy to transform the Matlab program into e.g. Fortran. However, the use of the original platform simplifies plotting of turbine characteristics, velocity triangles, etc. The program is validated against test-rig data from an industrial two-stage power turbine.
Experimental Investigation of the Effects of a Moving Shock Wave on Compressor Stator Flow
Authors: Langford Matthew D. , Breeze-Stringfellow Andrew , Guillot Stephen A. , Solomon William , Ng Wing F. , Estevadeordal Jordi
Published: November 2008

Linear cascade testing was performed to simulate the flow conditions experienced by stator blades in an axial compressor with supersonic relative Mach numbers at the inlet to the downstream embedded rotors. Experiments were conducted in a transonic blow-down wind tunnel with a nominal inlet Mach number of 0.65. A single moving normal shock introduced at the exit of the stator cascade simulated the bow shock from a downstream rotor. The shock was generated using a shock tube external to the wind tunnel. Pressure measurements indicated that the stator matched its design intent loading, turning and loss under steady flow conditions. Effects of the passing shock on the stator flow-field were investigated using shadowgraph photography and Digital Particle Image Velocimetry (DPIV). Measurements were taken with three different shock strengths. In each case, the passing shock induced a vortex around the trailing edge of the stator. The size and strength of these vortices were directly related to the shock strength. A suction side separation on the trailing edge of the stator was observed and found to correlate with the vortex blockage.
Local Heat/Mass Transfer Characteristics on a Rotating Blade With Flat Tip in a Low Speed Annular Cascade: Part 1 — Near-Tip Surface
Authors: Rhee Dong-Ho , Cho Hyung-Hee
Published: November 2008

The present study focuses on local heat/mass transfer characteristics on the near-tip region of a rotating blade. To investigate the local heat/mass transfer on the near-tip surface of the rotating turbine blade, detailed measurements of time-averaged mass transfer coefficients on the blade surfaces were conducted using a naphthalene sublimation technique. A low speed wind tunnel with a single stage annular turbine cascade was used. The turbine stage is composed of sixteen guide plates and blades with spacing of 34 mm, and the chord length of the blade is 150 mm. The mean tip clearance is about 2.5% of the blade chord. The tested Reynolds number based on inlet flow velocity and blade chord is 1.5×105 and the rotational speed of blade is 255.8 rpm for the design condition. The result at the design condition was compared with the results for the stationary blade to clarify the rotational effect, and the effects of incoming flow incidence angle were examined for incidence angles ranging from −15 to +7 degree. The off-design test condition is obtained by changing the rotational speed maintaining a fixed incoming flow velocity. Complex heat transfer characteristics are observed on the blade surface due to the complicated flow patterns, such as flow acceleration, laminarization, transition, separation bubble and tip leakage flow. The blade rotation causes an increase of the incoming flow turbulence intensity and a reduction of the tip gap flow. At off-design conditions, the heat transfer on the turbine rotor changes significantly due to the flow acceleration/deceleration and the incoming flow angle variation.
Towards Excellence in Turbomachinery CFD: A Hybrid Structured-Unstructured RANS Solver
Authors: Yang Hong , Nuernberger Dirk , Kersken Hans-Peter
Published: November 2008

A three-dimensional hybrid structured-unstructured RANS solver has been developed to simulate flows in complex turbomachinery geometries. It is built by coupling an existing structured CFD solver with a newly developed unstructured-grid module via a conservative hybrid-grid interfacing algorithm, so that it can get benefits from the both structured and unstructured grids. The unstructured-grid module has been developed with consistent numerical algorithms, data structure, user interface and parallelization to those of the structured one. The numerical features of the hybrid RANS solver are its second-order accurate upwind scheme in space, its SGS implicit formulation of time integration, and its accurate modeling of steady/unsteady boundary conditions for multistage turbomachinery flows. The hybrid-grid interfacing algorithm is essentially an extension of the conservative zonal approach that has been previously applied on the mismatched zonal interface of the structured grids, and it is fully conservative and also second-order accurate. Due to the mismatched grids are allowed at the block interface, users would have great flexibility to build the hybrid grids even with different structured and unstructured grid generators. The performance of the hybrid RANS solver is assessed with a variety of validation and application examples, through which the hybrid RANS solver has been demonstrated to be able to cope with the flows in complex turbomachinery geometries and to be promising for the future industrial applications.
Aerodynamic Concept for Very Large Steam Turbine Last Stages
Authors: Stu¨er Heinrich , Truckenmu¨ller Frank , Borthwick Don , Denton John D.
Published: November 2008

A move to ever larger low pressure steam turbines inevitably leads to an increase in the blade height to hub diameter ratio and the extent and magnitude of supersonic flow. This paper looks at the options available to the designer to mitigate the aerodynamic challenges associated with an increase in turbine size. Combinations of the effects of lean and sweep are explained using simple arguments based on flow field streamlines, then demonstrated in a stationary blade design by numerical and experimental methods. Developments in the principal regions of transonic and supersonic flow are shown by stationary blade hub and rotating blade tip examples. A stationary blade design which optimises the flow field effects shown is presented. The paper concludes by evaluating the unsteady flow in a very large final stage design.
The Design, Development and Evaluation of 3D Aerofoils for High Speed Axial Compressors: Part 2 — Simulation and Comparison With Experiment
Authors: Woollatt G. , Lippett D. , Ivey P. C. , Timmis P. , Charnley B. A.
Published: November 2008

The focus of this paper is to report on measurements from and simulation of Cranfield University’s 3-stage high-speed axial compressor test rig. This newly built rig is supported by European Commission funding and has tested a set of conventionally stacked 2D rotor and stator blades (Reference 1). The results were used to evaluate and to assess the performance of several commercially available CFD codes leading to the collaborative design of an advanced three-dimensional blade set. The philosophy behind the advanced design is described. The datum test results show that the state of the art, highly loaded, datum compressor is well matched with limited potential for loss reduction. A comparison is made between the measured results and a series of numerical analyses using the various CFD codes. Although the codes showed reasonable qualitative agreement with each other and the measured data, there were significant differences in the predicted performance of the datum build. Further the codes were unable to grade candidate redesigns consistently on a quantitative basis and therefore increased the difficulty of selecting suitable ‘3d’ features. Generic studies involving sweep, lean and recambering are used to evolve a design philosophy for the advanced three-dimensional design. Over cambering of the end-wall sections, coupled with a suitable stack of the blades, enables the blade count to be reduced. In the presence of a clearance combinations of sweep and lean are used to modify the loading in the clearance gap, thereby influencing the associated losses. The application of three-dimensional features redistributes the flow. The opportunity is therefore taken to rematch the sections based on the predicted results of the CFD codes. The above philosophy is used in the redesign of the datum compressor. Overall characteristics and exit traverse results from the test of the advanced build are compared to those from the datum build.
Development and Validation of the New ALSTOM GT24/GT26 Upgrade Compressor
Authors: Ko¨rba¨cher Heiko , Pfeiffer Andreas , Rothe Thomas
Published: November 2008

Since 1996 more than 75 ALSTOM GT24/GT26 heavy-duty gas turbines have been installed worldwide for the 50 Hz and 60 Hz market. The GT24/GT26 features a unique advanced sequential combustion system, a one stage high-pressure turbine and a 4-stage low-pressure turbine. The 22-stage compressor delivers a pressure ratio of 32:1. Typical power output is up to 281MW (GT24: 188MW) in single-cycle (with quench cooler) and more than 400MW in combined cycle operation (CCPP). In 2001, ALSTOM launched a performance improvement and upgrade program for these heavy-duty gas turbines after experiencing some problems with the advanced technology. Several upgrade packages were rolled out into the field after extensive testing at the Birr Test Center; Switzerland. A new GT24/GT26B-compressor is part of the upgrade and retrofit packages that were developed. These were extensively tested and validated during several test campaigns at the Test Centre in 2002. The GT24 compressor is a scaled version of the GT26 compressor. The compressor tests were completed in 2002 and since then fifteen engines have received the compressor upgrade and further installations are planned. This paper gives an overview of the compressor development, validation and the impact on the gas turbine performance of the GT24/GT26 upgrade compressor.
A Numerical Study of 3D Turbulent Cooling Jet Interaction Over a Range of Blowing Ratios, Turbulence Intensity and Turbulent Length Scale
Authors: Ratzlaff Jon , Orkwis Paul D. , Sekar Balu
Published: November 2008

Three-dimensional simulations of an unloaded cooled vane have been conducted for blowing ratios of 0.67, 1.02, and 1.4. For each blowing ratio, three free stream turbulence intensities of 1%, 10%, and 20% have been simulated. A brief investigation into the effects of length scale has also been performed at a turbulence intensity of 10% via a 40% reduction in length scale of. Three rows of cooling holes were simulated for a total of 31 cooling holes. The flow through each hole and the feed plenum were simulated. The first two rows of holes were inclined downward at 60° to the horizon, while the third row exited axially. The cases were run at Mach number 0.23 and Reynolds number based on the blade leading edge diameter, or thickness, of 4.1×104 with a main flow total temperature of 705.6K° and a cooling flow total temperature of 360K°, providing a cold to hot gas density ratio of approximately 2. Surface contours of film cooling effectiveness and static temperature, plots of η vs. s, exit plane static temperature contours, and exit plane plots of mass averaged total temperature are presented along with detailed streamline maps to show the propagation of cooling flows through the passage. The results indicate that cooling effectiveness was greatest for the 1.02 blowing ratio case. Higher blowing ratios resulted in streaks of uncooled blade surface between cooling holes in the showerhead region caused by cooling jet coupling and interactions, and the misplacement of the holes for this condition. These cooling patterns resulted in a cell-like structure of cooling flows in the downstream wake for the lowest turbulence intensity, although this was not observed with higher turbulence. Lastly, cooling flows impacted the lower wall of the passage for all cases. This occurred when cooling flows were either entrained by the corner separation for the two lower blowing ratio cases, or impacted the lower surface before the separation, as observed for high blowing and low turbulence. In the latter case this resulted in suppression of the corner separation in the trailing edge region of the blade.
The Effect of IGV Wake Impingement on the Flow Structure and Turbulence Around a Rotor Blade
Authors: Soranna Francesco , Chow Yi-Chih , Uzol Oguz , Katz Joseph
Published: November 2008

The flow structure and turbulence around the leading and trailing edges of a rotor blade operating downstream of a row of Inlet Guide Vanes (IGV) are investigated experimentally. Particle Image Velocimetry (PIV) measurements are performed in a refractive index matched facility that provides unobstructed view of the entire flow field. Data obtained at several rotor blade phases focus on modification to the flow structure and turbulence in the IGV wake as it propagates along the blade. The phase-averaged velocity distributions demonstrate that wake impingement significantly modifies the wall-parallel velocity component and its gradients along the blade. Due to spatially non-uniform velocity distribution, especially on the suction side, the wake deforms while propagating along the blade, expanding near the leading edge and shrinking near the trailing edge. While being exposed to the non-uniform strain field within the rotor passage, the turbulence within the IGV wake becomes spatially non-uniform and highly anisotropic. Several mechanisms, which are consistent with rapid distortion theory (RDT) and distribution of turbulence production rate, contribute to the observed trends. For example, streamwise (in rotor frame reference) diffusion in the aft part of the rotor passage enhances the streamwise fluctuations. Compression also enhances the turbulence production very near the leading edge. However, along the suction side, rapid changes to the direction of compression and extension cause negative production. The so-called wall blockage effect reduces the wall-normal component.
Effect of Blade Passage Surface Heat Extraction on Axial Compressor Performance
Authors: Shah P. N. , Tan C. S.
Published: November 2008

Axial compressor performance with heat extraction via blade passage surfaces (compressor cooling) is compared to its adiabatic counterpart, using computational experiments and meanline modeling. For a multistage compressor with an adiabatic design point, results at fixed corrected rotor speed indicate that if available, compressor cooling would: (i) raise the overall pressure ratio (at a given corrected flow), (ii) raise the maximum mass flow capability, (iii) raise the efficiency, defined as the ratio of isentropic work for a given pressure ratio to actual shaft work, and (iv) provide rear stage choking relief at low corrected speed. In addition, it is found that, if available, cooling in the front stages is better than in the rear stages. This is primarily a thermodynamic effect that results from the fact that, for a given gas, the compression work required to achieve a given pressure ratio decreases as the gas becomes colder. Heat transfer considerations indicate that the engineering challenges lie in achieving high enough heat transfer rates to provide significant impact to the compressor’s performance.
An Integrated Design System for Fans, Compressors, and Turbines: Part 4 — Through-Flow Solver
Authors: Medlock Allen , Miller Max J. , Konan S. Murthy , Chambers Ben , Nguyen Bao Q.
Published: November 2008

A new aerodynamic design system has been developed that includes a through-flow solver for fans, axial compressors and turbines, and radial compressors and turbines. Three earlier papers gave an overview of the system and described the interactive interface and geometry generators. This paper focuses on several special features in the through-flow solver that provide increases in aerodynamic designer productivity. Some of the key features are stations decoupled from flow paths, ability to accept a wide variety of input parameters, use of gas property routines, ability to inject flow non-uniformly with a different composition than the main flow gas composition, ability to access information from several airfoil geometry generator solutions, and clear, comprehensive error handling. These special features and others have provided major savings from productivity improvements and reductions in design cycle time.
Influence of Surface Roughness on the Aerodynamic Losses of a Turbine Vane
Authors: Zhang Qiang , Goodro Matt , Ligrani Phillip M. , Trindade Ricardo , Sreekanth Sri
Published: November 2008

The effects of surface roughness on the aerodynamic performance of a turbine vane are investigated for three Mach number distributions, one of which results in transonic flow and matches an arrangement employed in an industrial application. Four turbine vanes, each with the same shape and exterior dimensions, are employed with different rough surfaces. The non-uniform, irregular, three-dimensional roughness on the tested vanes is employed to match the roughness which exists on operating turbine vanes subject to extended operating times with significant particulate deposition on the surfaces. Wake profiles are measured for two different positions downstream the vane trailing edge. The contributions of varying surface roughness to aerodynamic losses, Mach number profiles, normalized kinetic energy profiles, Integrated Aerodynamics Losses (IAL), area-averaged loss coefficients, and mass-averaged loss coefficients are quantified. Total pressure losses, Mach number deficits, and deficits of kinetic energy all increase at each profile location within the wake as the size of equivalent sandgrain roughness increases, provided the roughness on the surfaces is uniform. Corresponding Integrated Aerodynamic Loss IAL magnitudes increase either as Mach numbers along the airfoil are higher, or as the size of surface roughness increases. Data are also provided which illustrate the larger loss magnitudes which are present with flow turning and cambered airfoils, than with symmetric airfoils. Also described are wake broadening, profile asymmetry, and effects of increased turbulent diffusion, variable surface roughness, and streamwise development.
Efficiency Improvement and Noise Reduction Through Stator-Stator Clocking Effect of a Two-Stage Turbine
Authors: Blaszczak Jaroslaw R.
Published: November 2008

The objective of the presented test program was to further experimentally investigate vane-indexing effect influence on the performance, noise and vibration of two-stage low-pressure turbine. Keeping the inlet conditions strictly constant during the tests, two turbine stages were experimentally investigated. Herein, some flow measurement results and the external characteristics for different circumferential positions of the stator vanes are described. Comparisons were made with numerical simulation and they showed good agreement. Experimental data and numerical simulations of stator vane surface pressures are presented to determine how the flow phenomena were affected by indexing of the airfoils for two cases: for nominal rotational speed and for off-design turbine conditions. In addition, correlation to acoustic noise and machine vibration level is presented. They have been found to be clocking dependent. The experimental investigations have been carried out on a two-stage turbine research facility at the Institute of Turbomachinery of the Technical University of Lodz, Poland.
Influence of Sweep on the 3D Shock Structure in an Axial Transonic Compressor
Authors: Bergner Jo¨rg , Kablitz Stephan , Hennecke Dietmar K. , Passrucker Harald , Steinhardt Erich
Published: November 2008

To improve the understanding of how sweep affects the 3D shock structure and the shock-vortex interaction, two transonic compressor bladings are measured with a 3D Laser-2-Focus setup, and the flow is analyzed for comparable operating conditions. Rotor 1 features radially stacked profiles, while rotor 3 introduces forward sweep. Performance comparisons of these rotors [1] indicated that forward sweep, as introduced in this design, has a beneficial effect on performance and stall margin. The detailed experimental analysis of the 3D shock structure and the influence of tip leakage vortex interaction at design speed is the scope of this paper.
Film Cooling Performance in a Simulated Turbine Blade Tip Geometry
Authors: Yoon Jae H. , Martinez-Botas Ricardo F.
Published: November 2008

One of the most problematic areas in gas turbine engines is the blade tip region, especially near the trailing edge, where it is very difficult to provide sufficient cooling. In all configurations with unshrouded tips, a clearance gap exists between the turbine blade and the outer shroud. The pressure difference between the suction and pressure sides of the blade drives a sink-like flow through this gap. The combination of leakage flow from the freestream and coolant flow induces high convective heat transfer coefficients on the blade tip surface. The resultant thermal loading can be significant and detrimental to the turbine blade tip durability, leading to early failure. Film cooling can be provided by means of a series of holes located on the tip itself providing protection not dissimilar to film cooling of the main blade. However, the interaction of coolant and the separation bubble resulted in a significantly different film cooling performance to that of non-tip cases. An experimental investigation of the simulated turbine blade tip is presented in here. The first section discusses PIV flow field measurements, the second covers the measurement of film cooling effectiveness and the third heat transfer measurements. All three parts investigate the effect of using different film cooling injection points and blowing ratios for injection on the blade tip itself, close to the pressure surface corner. Additionally, the effect of varying the corner radii between the pressure surface and the tip is reported. The experimental method uses the steady state liquid crystal technique. A Reynolds number of 30,000 based on the clearance gap hydraulic diameter for the main flow was used.
Numerical Designed Experiment to Optimize a Ported Shroud to Extend the Operability Margin of a Centrifugal Compressor
Authors: Slovisky J. , Mansour M. L. , Barton M. T. , Palmer D. L.
Published: November 2008

This paper describes the Computational Fluid Dynamic (CFD) numerical optimization of a modern centrifugal compressor impeller with a ported shroud for increased surge margin. The vent configuration selected was a full circumference, constant-width slot. A multiblock, steady flow three dimensional (3D) viscous RANS model (ADPAC) is used with parallel processing capability to increase computational speed. Grid generation is performed in an automated fashion to enable the timely optimization of the ported shroud configuration. A designed experiment (DoE) approach is used to minimize the number of vent configurations to be modeled, to ensure that factor interaction effects are captured, and to facilitate the definition of an optimum vent configuration. The DoE is a 2 factor, 2 level full factorial experiment with a center point included to detect possible curvature in the solution surface. The factors optimized are slot width and the flow-wise location of the slot. The numerical technique verifies the ability of the ported shroud to extend compressor stall margin at the part-speed operating condition, while maintaining acceptable high speed performance, in good agreement with test data for a similar impeller with a ported shroud. The use of a DoE method coupled with CFD modeling identified an optimized vent configuration with a minimum of time and effort. The CFD results also provide enhanced understanding of the device physics.
Inverse BEM Method to Identify Surface Temperatures and Heat Transfer Coefficient Distributions at Inaccessible Surfaces
Authors: Kassab Alain J. , Divo Eduardo A. , Chyu Minking K. , Cunha Frank J.
Published: November 2008

The purpose of the inverse problem considered in this study is to resolve heat transfer coefficient distributions by solving a steady-state inverse problem. Temperature measurements at interior locations supply the additional information that renders the inverse problem solvable. A regularized quadratic functional is defined to measure the deviation of computed temperatures from the values under current estimates of the heat transfer coefficient distribution at the surface exposed to convective heat transfer. The inverse problem is solved by minimizing this functional using a parallelized genetic algorithm (PGA) as the minimization algorithm and a two-dimensional multi-region boundary element method (BEM) heat conduction code as the field variable solver. Results are presented for a regular rectangular geometry and an irregular geometry representative of a blade trailing edge and demonstrate the success of the approach in retrieving accurate heat transfer coefficient distributions.
Separation and Transition Control on an Aft-Loaded Ultra-High-Lift LP Turbine Blade at Low Reynolds Numbers: Low-Speed Investigation
Authors: Zhang Xue Feng , Vera Maria , Hodson Howard , Harvey Neil
Published: November 2008

An experimental study was conducted to improve the performance of an aft-loaded ultra-high-lift low-pressure turbine blade known as U2 at low Reynolds numbers. This was achieved by manipulation of the laminar-turbulent transition process on the suction surface. The U2 profile was designed to meet the targets of reduced cost, weight and fuel burn of aircraft engines. The studies were conducted on both low-speed and high-speed experimental facilities under the unsteady flow conditions with upstream passing wakes. The current paper presents the low-speed investigation results. On the smooth suction surface, the incoming wakes are not strong enough to suppress the separation bubble due to the strong adverse pressure gradient on the suction surface and the low wake passing frequency, which allows the separation between the wakes more time to re-establish. Therefore, the profile losses of this ultra-high-lift blade are not as low as conventional or high-lift blades at low Reynolds numbers even in unsteady flows. Two different types of passive separation control devices, i.e. surface trips and air jets, were investigated to further improve the blade performance. The measurement results show that the profile losses can be further reduced to the levels similar to those of the high-lift and conventional blades due to the aft-loaded nature of this ultra-high-lift blade. Detailed surveys of the blade surface boundary layer developments showed that the loss reduction was due to the suppression of the separation underneath the wakes, the effect of the strengthened calmed region and the smaller separation bubble between wakes.
Improved Diffuser/Volute Combinations for Centrifugal Compressors
Authors: Steglich T. , Kitzinger J. , Seume J. R. , Van den Braembussche R. A. , Prinsier J.
Published: November 2008

Internal volutes have a constant outer radius, slightly larger than the diffuser exit radius, and the circumferential increase of the cross section is accommodated by a decrease of the inner radius. They allow the design of compact radial compressors and hence are very attractive for turbochargers and high-pressure pipeline compressors where small housing diameters have a favorable impact on weight and cost. Internal volutes, however, have higher losses and lower pressure rise than external ones in which the center of the cross sections is located at a larger radius than the diffuser exit. This paper focuses on the improvement of the internal volute performance by taking into account the interaction between the diffuser and the volute. Two alternative configurations with enhanced aerodynamic performance are presented. A first one features a novel, non-axi-symmetric diffuser/internal volute combination. It demonstrates an increased pressure ratio and lower loss over most of the operating range at all rotational speeds. The circumferential pressure distortion at off design operation is slightly larger than in the original configuration with a concentric vaneless diffuser. Alternatively, a parallel-walled Low-Solidity Diffuser (LSD) with an internal volute allows a reduction of the unsteady load on the impeller and an improved performance close to the one of a vaneless concentric diffuser with a large external volute.
Wake-Induced Transitional Flow Over a Highly-Loaded LP Turbine Blade Through Large-Eddy Simulation
Authors: Sarkar S.
Published: November 2008

An attempt is made to describe the physical mechanism of transition of an inflexional boundary layer over the suction surface of a highly cambered low-pressure (LP) turbine blade influenced by the periodic passing wakes. Large-eddy simulations (LES) of wake passing over the T106 profile for a Reynolds number of 1.6×105 (based on the chord and exit velocity) are performed using wake data extracted from precursor simulations of cylinder replacing a moving bar in front of the cascade. The three-dimensional, time-dependent, incompressible Navier-Stokes equations in fully covariant form are solved using a symmetry-preserving finite difference scheme of second-order spatial and temporal accuracy. The present LES results are compared with experiments and DNS. The operating condition of a high-lift LP turbine blade leads to the formation of a separation bubble on the suction side. The interactions of incoming wake with this separation bubble complicate the transition process. Enhanced receptivity of inflexional boundary layer causes amplification of the perturbations produced by the passing wake leading to the formation of coherent vortices within the boundary layer. The transition mechanism during the wake-induced path is highly influenced by the convection and breakdown of these coherent vortices. Streamwise evolution of turbulent kinetic energy and production illustrates that these vortices play an important role in generation of turbulence and thus to decide the transitional length, which becomes time-dependent. LES results resolve a multimoded transition on the suction surface and the calmed region. The calmed region is nothing but an attached flow with low production as the boundary layer tends to relax after wake passing; the level of turbulent intensity suggests that the boundary layer is in a state of transition rather than laminarized.
Effects of Mid-Passage Gap, Endwall Misalignment and Roughness on Endwall Film-Cooling
Authors: Cardwell N. D. , Sundaram N. , Thole K. A.
Published: November 2008

To maintain acceptable turbine airfoil temperatures, film-cooling is typically used whereby coolant, extracted from the compressor, is injected through component surfaces. In manufacturing a turbine, the first stage vanes are cast in either single airfoils or double airfoils. As the engine is assembled, these singlets or doublets are placed in a turbine disk in which there are inherent gaps between the airfoils. The turbine is designed to allow outflow of high pressure coolant rather than hot gas ingestion. Moreover, it is quite possible that the singlets or doublets become misaligned during engine operation. It has also become of interest to the turbine community as to the effect of corrosion and deposition of particles on component heat transfer. This study uses a large scale turbine vane in which the following two effects are investigated: the effect of a mid-passage gap on endwall film-cooling and the effect of roughness on endwall film-cooling. The results indicate that the mid-passage gap was found to have a significant effect on the coolant exiting from the combustor-turbine interface slot. When the gap is misaligned, the results indicate a severe reduction in the film-cooling effectiveness in the case where the pressure side endwall is below the endwall associated with the suction side of the adjacent vane.
Experimental Investigation of Aerodynamic Losses of Different Shapes of a Shrouded Blade Tip Section
Authors: Nirmalan Nirm V. , Bailey Jeremy C.
Published: November 2008

An experimental investigation was conducted to study the effects on aerodynamic losses of different tip shroud shapes of a shrouded turbine blade. Pressures were measured on the airfoil surface near the tip and a plane downstream of the exit plane in a three-airfoil stationary cascade. The instrumented center airfoil and the two slave airfoils modeled the aerodynamic tip section of a blade and have the capability to vary tip clearance. The experiments were run at tip-clearances varying from 0.25% to 1.67% and at an exit Reynolds number of 1.25 × 106 and Mach Number of 0.95. The paper presents the influence of three tip-shroud shapes and five different tip-clearances on the aerodynamic losses.
Investigation on Cooling Performance of Impingement Cooling Devices Combined With Pins
Authors: Quan Dongliang , Liu Songling , Li Jianghai , Liu Gaowen
Published: November 2008

Integrated impingement and pin fin cooling devices have comprehensive advantages of hot-side film cooling, internal impingement cooling, large internal heat transfer area and enhanced heat exchange caused by the pin fin arrays, so it is considered a promising cooling concept to meet the requirements of modern advanced aircraft engines. In this paper, experimental study, one dimensional model analysis and numerical simulation were conducted to investigate cooling performance of this kind of cooling device. A typical configuration specimen was made and tested in a large scale low speed closed-looped wind tunnel. The cooling effectiveness was measured by an infrared thermography technique. The target surface was coated carefully with a high quality black paint to keep a uniform high emissivity condition. The measurements were calibrated with thermocouples welded on the surface. Detailed two-dimensional contour maps of the temperature and cooling effectiveness were obtained for different pressure ratios and therefore different coolant flow-rates through the tested specimen. The experimental results showed that very high cooling effectiveness can be achieved by this cooling device with relatively small amount of coolant flow. Based on the theory of transpiration cooling in porous material, a one dimensional heat transfer model was established to analyze the effect of various parameters on the cooling effectiveness. The required resistance and internal heat transfer characteristics were obtained from experiments. It was found from this model that the variation of heat transfer on the gas side, including heat transfer coefficient and film cooling effectiveness, of the specimen created much more effect on its cooling effectiveness than that of the coolant side. The heat transfer intensities inside the specimen played an important role in the performance of cooling. In the last part of this paper, a conjugate numerical simulation was carried out using commercial software FLUENT 6.1. The domain of the numerical simulation included the specimen and the coolant. Detailed temperature contours of the specimen were obtained for various heat transfer boundary conditions. The calculated flow resistance and cooling effectiveness agree well with the experimental data and the predictions with the one-dimensional analysis model. The numerical simulations reveal that the impingement of the coolant jets in the specimen is the main contribution to the high cooling effectiveness.
3D Fluid–Solid Heat Transfer Coupling of an Aero Engine Pre-Swirl System
Authors: Illingworth Justin B. , Hills Nicholas J. , Barnes Christopher J.
Published: November 2008

This paper examines the application of a newly developed code, which couples a commercial computational fluid dynamics (CFD) code (FLUENT) directly to an in-house finite element solver used by Rolls-Royce plc. The coupling is achieved by passing metal temperatures from the finite element solver to define CFD boundary conditions, and then passing heat fluxes from the resulting CFD solutions back to the finite element model. This coupling method was applied to a real engine test case, modelling the pre-swirl system of the Rolls-Royce Trent 500 aero engine. The initial phase of the analysis couples an axisymmetric finite element whole engine model to two axisymmetric CFD models. The CFD models define the pre-swirl fluid domain during idle and maximum take off engine running conditions. This demonstrates the code’s ability to accommodate temperature transients through the use of multiple CFD models. A further analysis is then performed coupling the same finite element engine model to a 3D CFD model replicating stabilised maximum take off conditions. This is compared with the axisymmetric to axisymmetric analysis to identify the approximations inherent in using an axisymmetric model to investigate a three dimensional flow structure.
Rayleigh-Be´nard Convection in Open and Closed Rotating Cavities
Authors: King Martin P. , Wilson Michael , Owen J. Michael
Published: November 2008

Buoyancy effects can be significant in the rotating annular cavities found between compressor discs in gas-turbine engines, where Rayleigh numbers above 1012 are common. In some engines, the cavity is ‘closed’, so that the air is confined between four rotating surfaces: two discs and inner and outer cylinders. In most engines, however, the cavity is ‘open’, and there is an axial throughflow of cooling air at the centre. For open rotating cavities, a review of the published evidence suggests a Rayleigh-Be´nard type of flow structure, in which, at the larger radii, there are pairs of cyclonic and anti-cyclonic vortices. The toroidal circulation created by the axial throughflow is usually restricted to the smaller radii in the cavity. For a closed rotating annulus, solution of the unsteady Navier-Stokes equations, for Rayleigh numbers up to 109, show Rayleigh-Be´nard convection similar to that found in stationary enclosures. The computed streamlines in the r-θ plane show pairs of cyclonic and anti-cyclonic vortices; but, at the larger Rayleigh numbers, the computed isotherms suggest that the flow in the annulus is thermally mixed. At the higher Rayleigh numbers, the computed instantaneous Nusselt numbers are unsteady and tend to oscillate with time. The computed time-average Nusselt numbers are in good agreement with the correlations for Rayleigh-Be´nard convection in a stationary enclosure, but they are significantly higher than the published empirical correlations for a closed rotating annulus.