bluuebunny/crossref_metadata_2025 huggingface dataset:
Determination of the Efficiency of a Cooled HP Turbine in a Compression Tube Facility
Authors: De´nos Remy , Paniagua G. , Yasa T. , Fortugno E.
Published: September 2008

The efficiency of a cooled transonic turbine stage was measured in a compression tube facility. The formulation takes into account mechanical losses, coolant flows and leakage flows. The proposed methodology allows computing the efficiency independently from the test rig. Owing to the short testing time (∼0.5 s), specific measurement and data reduction techniques are used. The paper details how the power, the overall mass flow, the mass-averaged inlet quantities and the stage pressure ratio are determined. The measurements of mechanical losses and of the thermodynamic properties of coolant and leakage flows are also described. Finally, results are presented and supported by an uncertainty analysis that identifies the random and systematic contributions to the final error.
Multidisciplinary Integration and Robustness Evaluation Applied to Low Pressure Turbine Casing Design
Authors: De Poli Gian Paolo , Frola Carlo , Gallizio Massimo , Fattore Luca , Mattone Massimiliano
Published: September 2008

In modern aerospace engineering design context one of the most important task is managing and simulate properly the effect of uncertainties on the response and performance of the system. In fact real engineering problems are characterised by random variations of material property, variation of loading conditions, manufacturing tolerances, etc. Different approaches have been developed by the research community to address uncertainties; while reliability methods primarily deal with probability of constraint satisfaction or violation, robust design methods have focused on the variation of system responses due to design parameters random variation. In this paper a robust design (RD) analysis of gas-turbine casing is performed in order to estimate how much uncertainties affect the life of the component. The RD analysis is performed in a multidisciplinary environment since the casing is subjected to thermo-mechanical loads. First thermal steady-state analysis has been performed changing randomly the boundary conditions (heat transfer coefficient and air temperature) and the temperature distribution on the casing is calculated. Then the structural analysis is performed changing geometrical dimensions on the base of defined tolerances and process capability. The result of the Monte Carlo analysis is a statistic distribution of the stress in the critical locations. This information is used to estimate the statistic distribution of the life in each critical location. The obtained result consents to evaluate the risk that some critical location exceeds the life margin limit.
Design and Commissioning of a Transonic Test Section for the Investigation of Shock Wave-Film Cooling Interactions
Authors: Ochs M. , Schulz A. , Bauer H.-J.
Published: September 2008

Transonic turbine flows are usually characterized by trailing-edge shocks. The oblique shock wave generated at the pressure side of an airfoil impinges on the suction side of the adjacent blade/vane, where it interacts with the boundary layer or cooling film. The present study deals with the design of a generic transonic test rig, which reproduces this kind of interaction under conditions similar to those found in real engines. The design of the set-up was strongly influenced by the requirement to use a Schlieren-System for flow visualization and an Infrared-Thermography-System for heat transfer measurements. A supersonic nozzle design is presented, which allows setting the desired Mach-number distribution and controlling the shock formation in the supersonic test section. The geometry of the supersonic part of the transonic test section has been designed using the “Method of Characteristics”. Viscous effects were taken into account by CFD-studies. First experimental results, in terms of static pressure distributions and Schlieren-images, show excellent agreement with the predicted flow behavior.
Blading Aerodynamics Design Optimization With Mechanical and Aeromechanical Constraints
Authors: Li H.-D. , He L. , Li Y. S. , Wells R.
Published: September 2008

A blading design optimization system has been developed using an aeromechanical approach and harmonic perturbation method. The developed system has the capability to optimize aero-thermal performance with constraints of mechanical and aeromechanical integrity at the same time. ‘Aerodynamic mode shape’ is introduced to describe geometry deformation which can effectively reduce the number of design parameters while preserving surface smoothness. Compared to the existing design optimization practices, the present system is simpler, more accurate and effective. A redesign practice of the NASA rotor-67 at the peak efficiency point shows that the aero thermal efficiency can be improved by 0.4%, whilst the maximum static stress has been increased by 33%. Aeromechanical analysis of the optimized blade shows that the aerodynamic damping of the least stable first flap mode is still well above the critical value though the natural frequencies of the first 5 modes have been reduced by 1∼4%. The present finding highlights the need for more concurrent integrations of mechanics, aerodynamics and aeromechanics design optimization.
Experimental and Numerical Investigations of Wake Passing Effects Upon Aerodynamic Performance of a LP Turbine Linear Cascade With Variable Solidity
Authors: Funazaki Ken-ichi , Yamada Kazutoyo , Ono Takahiro , Segawa Ken-ichi , Hamazaki Hiroshi , Takahashi Akira , Tanimitsu Haruyuki
Published: September 2008

This paper deals with experimental and numerical studies on the flow field around a low-pressure linear turbine cascade whose solidity is changeable. The purpose of them is to clarify the effect of incoming wakes upon the aerodynamic loss of the cascade that is accompanied with separation on the airfoil suction surface, in particular for low Reynolds number conditions and/or low solidity conditions. Cylindrical bars on the timing belts work as wake generator to emulate wakes that impact the cascade. Pneumatic probe measurement is made to obtain total pressure loss distributions downstream of the cascade. Hot-wire probe measurement is also conducted over the airfoil suction surface. Besides, LES-based numerical simulation is executed to deepen the understanding of the interaction of the incoming wakes with the boundary layer containing separation bubble.
Analytical Modeling of Radial Fluid Migration in the Boundary Layer of Axial Flow Turbomachinery Blades
Authors: Vad Ja´nos
Published: September 2008

The paper presents an analytical model for qualitative prediction of physical effects stimulating or retarding the radial outward migration of high-loss fluid in the suction side boundary layer of the blades of isolated axial flow rotors. This fluid motion is considered as the main cause of tip stalling. In elaboration of the model, the radial acceleration of fluid particles is expressed, and the Reynolds-averaged Navier-Stokes equation is applied for description of flow. On the basis of the model, guidelines are formulated in order to retard the inception of stall.
Investigation of the Flow Field in a HP Turbine Stage for Two Stator-Rotor Axial Gaps: Part I — 3D Time-Averaged Flow Field
Authors: Gaetani P. , Persico G. , Dossena V. , Osnaghi C.
Published: September 2008

An extensive experimental analysis on the subject of unsteady flow field in high pressure turbine stages was carried out at the Laboratorio di Fluidodinamica delle Macchine (LFM) of Politecnico di Milano. The research stage represents a typical modern HP gas turbine stage designed by means of 3D techniques, characterised by a leaned stator and a bowed rotor and operating in high subsonic regime. The first part of the program concerns the analysis of the steady flow field in the stator-rotor axial gap by means of a conventional five-hole probe and a temperature sensor. Measurements were carried out on eight planes located at different axial positions allowing the complete definition of steady flow field both in absolute and relative frame of reference. The evolution of the main flow structures, such as secondary flows and vane wakes, downstream of the stator are here presented and discussed in order to evidence the stator aerodynamic performance and, in particular, the different flow field approaching the rotor blade row for the two axial gaps. This results set will support the discussion of the unsteady stator-rotor effects presented in paper Part II. Furthermore, 3D time-averaged measurements downstream of the rotor were carried out at one axial distance and for two stator-rotor axial gaps. The position of the probe with respect to the stator blades is changed by means of rotating the stator in circumferential direction, in order to describe possible effects of the non-uniformity of the stator exit flow field downstream of the stage. Both flow fields, measured for the nominal and for a very large stator-rotor axial gap, are discussed and results show the persistence of some stator flow structures downstream of the rotor, in particular for the minimum axial gap. Eventually the flow fields are compared to evidence the effect of the stator-rotor axial gap on the stage performance from a time-averaged point of view.
Engine Design Studies for a Silent Aircraft
Authors: Hall Cesare A. , Crichton Daniel
Published: September 2008

The Silent Aircraft Initiative is a research project funded by the Cambridge-MIT Institute aimed at reducing aircraft noise to the point where it is imperceptible in the urban environments around airports. The propulsion system being developed for this project has a thermodynamic cycle based on an ultra-high bypass ratio turbofan combined with a variable area exhaust nozzle and an embedded installation. This cycle has been matched to the flight mission and thrust requirements of an all-lifting body airframe, and through precise scheduling of the variable exhaust nozzle, the engine operating conditions have been optimized for maximum thrust at top-of-climb, minimum fuel consumption during cruise and minimum jet noise at low altitude. This paper proposes engine mechanical arrangements that can meet the cycle requirements and, when installed in an appropriate airframe, will be quiet relative to current turbofans. To reduce the engine weight a system with a gearbox, or some other form of shaft speed reduction device, is proposed. This is combined with a low-speed fan and a turbine with high gap-chord spacing to further reduce turbomachinery source noise. An engine configuration with three fans driven by a single core is also presented and this is expected to have further weight, fuel burn and noise benefits.
An Experimental Investigation of Secondary Flows and Loss Development Downstream of a Highly Loaded Low Pressure Turbine Outlet Guide Vane Cascade
Authors: Hja¨rne Johan , Chernoray Valery , Larsson Jonas , Lo¨fdahl Lennart
Published: September 2008

This paper presents a detailed experimental investigation of the evolution of secondary flow field characteristics and losses at several measurement planes downstream of a highly loaded low pressure turbine/outlet guide vane (LPT/OGV). The experiments were carried out in a linear cascade at Chalmers in Sweden. Several realistic upstream incidences and turbulence intensities have been investigated for one Reynolds number. Downstream characteristics have been measured with a 5-hole pneumatic probe. This allows for the determination of the mean vortical structures, their development and their interactions. The passage vortex and the blade shed vorticity are clearly visible at different downstream positions. Their intensity is shown to be strongly dependent on the inlet flow angle. The turbulence level seems to play a role on both the mixing within, and between the structures. The measurements also show that the losses along the blade span are dependent on the development of these structures.
Strut Losses in a Diverging Annular Diffuser With Swirling Flow
Authors: Feldcamp G. K. , Birk A. M.
Published: September 2008

An experimental investigation into the overall influence of struts spanning a double divergent annular diffuser followed by a straight cored annular diffuser has been undertaken in order to determine the performance of various strut configurations over a wide range of inlet swirl conditions. Two strut profiles have been investigated in four and eight strut configurations. Results have shown that the presence of struts under no swirl conditions have a relatively small effect on the overall total pressure loss. Increasing the inlet swirl angle to 20° has shown that the struts are able to assist in recovery of the swirling flow such that the pressure recovery nearly approaches that without struts, despite increased total pressure losses. Performance at 40° swirl is highly dependent on strut profile; the higher thickness-to-chord ratio strut configurations show minimal decrease in pressure recovery compared to 20° swirl, while the lower thickness-to-chord ratio configurations experiences a significant decrease as the result of significant flow separation from the struts. The exit swirl number has been shown to correlate strongly with the strut profile shape, while the number of struts had only a secondary influence. The exit velocity profiles show significant distortions at 40° swirl, and as a result the ideal pressure recovery calculated from the inlet and exit profiles change with strut configuration at 40° swirl.
The Performance of Air-Air Ejectors With Triangular Tabbed Driving Nozzles
Authors: McBean S. F. , Birk A. M.
Published: September 2008

This paper describes an experimental investigation into the effects of geometrical variations on ejector system performance when the driving nozzle includes delta mixing tabs. Mixing tabs have been shown to provide good mixing performance with comparable back-pressure penalties to other types of enhanced mixing nozzles. The performance parameters of most interest are pumping, mixing, and back-pressure. Geometric parameters studied include standoff distance, mixing-tube diameter, and tab angle. Experimental testing showed significant performance improvements in mixing and pumping with a decrease in tab angle. Maximum mixing was found to occur with tab angles positioned at 120°. Exceptional mixing was also observed with increased standoff. Back-pressure was shown to increase with increasing standoff and decreasing tab angle, but was not affected by mixing-tube diameter. In addition, a zone of recirculation was identified at the entrance to the mixing-tube. This zone is thought to have an influence on ejector performance.
Numerical Investigation of the Effect of Different Back Sweep Angle and Exducer Width on the Impeller Outlet Flow Pattern of a Centrifugal Compressor With Vaneless Diffuser
Authors: Hildebrandt A. , Genrup M.
Published: September 2008

This paper presents a numerical investigation of the effect of different back sweep angles and exducer widths on the steady-state impeller outlet flow pattern of a centrifugal compressor with a vaneless diffuser. The investigations have been performed with commercial CFD and in-house programmed 1-D codes. CFD calculations aim to investigate how flow pattern from the impeller is quantitatively influenced by compressor geometry parameters; thereby, the location of wake and its magnitude (flow angle and relative velocity magnitude) are analyzed. Results show that the increased back sweep impeller provides a more uniform flow pattern in terms of velocity and flow deviation angle distribution, and offers better potential for the diffusion process inside a vaneless (or vaned) diffuser. Secondary flux fraction and flow deviation angle from CFD simulation are implemented into the 1-D two-zone program to improve 1-D prediction results.
HPT/LPT Interaction and Flow Management in the Inter-Turbine Space of a Modern Axial Flow Turbine
Authors: Yao Jixian , Carson Scott
Published: September 2008

Understanding the unsteady flows of turbine sub-component interaction has been an important aspect in the design of modern aircraft engines to improve component performance. This research is focused on the inter-turbine space and its neighboring blade rows where intensive interactions of the two turbine sub-components occur. Via large-scale parallel numerical simulations, the paper presents the loss mechanism of the HPT rotor shock-wake interaction, the interaction loss transfer, analysis of the 1st LPT nozzle loading, the performance and flow capacity of the 1st LPT stage. Flow management is performed via a reduced-shock HPT rotor design and a 2-in spacer in the inter-turbine space, aimed at reducing interaction loss and wake ingestion loss by the LPT. The effectiveness of these means is analyzed and comparisons are made to baseline configurations.
Effects of Reynolds Number and Tip Clearances on the Performance of a Centrifugal Compressor at Micro Scale
Authors: Isomura Kousuke , Teramoto Susumu , Togo Shin-Ichi , Hikichi Kousuke , Endo Yuki , Tanaka Shuji
Published: September 2008

The performance of a centrifugal compressor at micro-scale has been assessed by a combination of experiments and numerical calculations. A micro-compressor of 10mm diameter has been fabricated and tested at the condition of 75μm tip clearance. The test results showed higher adiabatic efficiencies than the target value for the compressor to realize a 100W class gas turbine generator. However, this is due to the heat leakage, and a heat leakage model is proposed to be added in the numerical simulations. After matching the calculation results to the experimental results by modeling the heat loss through the wall, calculations of various Reynolds number and tip clearances have been conducted. The study shows that the rate of efficiency reduction due to the increase of the tip clearance is within the range of the data that Pampreen [5] used to construct his model, and the micro-compressor can achieve the target efficiency either by reducing the tip clearance to 15μm, which is practically too small to be achieved. The alternatives methods to achieve the target efficiency are to increase the blade height for about three times, or scale up the compressor. These methods require the increase of the power output of the target gas turbine engine. The study shows that the target efficiency can be better achieved by scaling up the impeller to 15mm diameter to allow increasing the tip clearance to 50μm.
Shape Optimisation of Axial Fan Blades Using Genetic Algorithms and a 3D Navier-Stokes Solver
Authors: Lotfi O. , Teixeira J. A. , Ivey P. C. , Kinghorn I. R. , Sheard A. G.
Published: September 2008

The paper describes the development of an automated design process which was developed to aerodynamically optimise an industrial fan blade geometry taking account of the predicted three dimensional flow. The optimiser employs a genetic algorithm for global optimisation purposes and is coupled to the academic Navier-Stokes solver MULTIP. The optimization task is accomplished by modifying the blade camber line, lean and sweep while keeping the blade thickness distribution and mass flow rate, constant. A number of different configurations have been studied and the behaviour of genetic algorithm tested. Specific interfaces were developed in order to link the optimization code, the automatic grid generator STAGEN, utilised to define the computational meshes, and the three-dimensional Navier-Stokes solver within an automated design loop. The results obtained show that the genetic algorithm when coupled to a CFD tool is not only capable of achieving an improvement in the designs of existing axial fan blades effectively but also that they achieve these results with a minimum amount of user expertise.
Investigations of the Boundary Layer on a Highly Loaded Compressor Cascade With Wake-Induced Transition
Authors: Iseler Jens , Hilgenfeld Lothar , Pfitzner Michael
Published: September 2008

The flow field through a turbomachinery compressor cascades is significantly affected by the unsteady flow originating from the upstream blade rows. The interaction is caused by the wakes from the upstream blades, which affect the properties of the boundary layer of the downstream blades. In addition, pressure fluctuations exist between upstream and downstream blades. These phenomenona play a significant role in the loss generation process on turbomachinery blades because it influences the onset of transition in the boundary layer and has the potential to suppress a boundary layer separation in some cases. Extensive experimental investigations have been performed at the Institute of Jet Propulsion in Neubiberg, where these effects where studied in detail. The measurements were performed on a large scale compressor cascade called V103-220. The chord length of l = 220 mm chosen allowed the unsteady boundary layer development to be studied in great detail and provided high quality data for this complex flow, which can be used for the validation of CFD codes. Unsteady CFD calculations were performed using the RANS-code TRACE developed at DLR Cologne. A modern variant of the Wilcox k-ω turbulence model in combination with a newly implemented transition model was used, allowing a better determination of multimode transition. A multiblock grid with an O-type grid around the blade and a boundary layer resolution of y+<1 was used. Experimental and numerical results confirm that wake passing has a large influence on the unsteady boundary layer development also in this compressor flow case. The premature forced transition is followed by a stable calmed region, which partially suppresses laminar separation due to its higher shear stress level and delays the onset of transition in the path between wakes. In addition, it was found that the leakage from two slots, which are opened in the rig when the wake generator device is installed, changes the flow field considerably. This effect is not fully reproduced by the CFD calculations. To study this effect in more detail, three-dimensional steady and unsteady CFD calculations were undertaken and are being continued.
Performance Impacts Due to Wake Mixing in Axial-Flow Turbomachinery
Authors: Praisner T. J. , Clark J. P. , Nash T. C. , Rice M. J. , Grover E. A.
Published: September 2008

One of the last loss mechanisms remaining to be quantified and correlated for inclusion in meanline predictive systems concerns the mixing of wakes across downstream airfoil rows. Here, we demonstrate that the unsteady losses incurred as turbomachinery wakes mix in downstream rows are a function of the velocity ratio across the downstream row as calculated in the frame of reference of wake generation. Analytical and computational results, compared to measurements of wakes mixing under variable free-stream velocity conditions, reveal that wake-loss modification is primarily a result of an inviscid dilation of the stream tubes that comprise the wake fluid. Further, simulations of wakes exposed to a range of turbomachinery-specific velocity ratios indicate that wake-loss augmentation caused by stream-tube dilation is significantly more pronounced than wake-loss reductions imparted by stream-tube contraction. It is demonstrated that wakes in turbines are dilated in the adjacent downstream row, whether it is a vane or a blade row, through a work extraction process that occurs in the wake-generation reference frame. Finally, comparisons between rig data and CFD simulations suggest that wake-mixing losses, enhanced by downstream rows, can contribute as much as 1.5 percent of lost efficiency in multistage low-pressure turbines.
Experimental Characterization of Vaneless Diffuser Rotating Stall: Part VI — Reduction of Three Impeller Results
Authors: Carnevale E. A. , Ferrara G. , Ferrari L. , Baldassarre L.
Published: September 2008

Vaneless diffuser rotating stall is a major problem for centrifugal compressors since it is a limit to their working range. In the literature some good correlations for predicting stall inception can be found but they do not adequately cover the case of the last stage configuration, especially for very low blade-outlet-width-to-impeller-radius-ratio impellers typically used in high-pressure applications. Extensive research has been performed to define diffuser stall limits for this family of stages: three impellers characterized by different blade-outlet-width-to-impeller-radius-ratios have been tested with different diffuser configurations (different pinch shapes, diffuser widths and diffusion ratios). The basic configuration comprises a 1:1 geometrical scale stage with a return channel upstream, a 2D impeller with a vaneless diffuser and a volute with a constant cross sectional area downstream. Several diffuser types with different widths and diffusion ratios were tested. Detailed experimental results have been reported in previous works [1, 2, 3 and 4]. In this paper experimental data are reviewed in order to analyze impeller influence on diffuser stability and to develop some summarizing consideration on stall behavior of vaneless diffuser for impeller with low blade-outlet-width-to-impeller-radius-ratio.
Influence of Aerodynamic Loading on Rotor-Stator Aerodynamic Interaction in a Two-Stage Low Pressure Research Turbine
Authors: Canepa Edward , Formosa Piergiorgio , Lengani Davide , Simoni Daniele , Ubaldi Marina , Zunino Pietro
Published: September 2008

The unsteady flow within a two-stage low-pressure research turbine equipped with high lift profiles has been investigated in detail for three different aerodynamic loading conditions. Experiments have been carried out at low speed. Velocity and turbulence intensity in the blade-to-blade plane at midspan have been measured by means of a crossed hot-wire probe, upstream and downstream of each blade row. The probe has been traversed circumferentially over 1.5 bladings pitch and the phase-locked data acquisition and ensemble average technique have been used to reconstruct the flow in space and time. The effects of multistage configuration have been identified and analyzed by considering the velocity components and turbulence intensity. Potential interaction from the downstream blading in relative motion, periodic wake perturbations from the upstream blading and preceding stage perturbations make the flow in the second stage extremely complex. Overall the flow downstream of rotors is perturbed in space by upstream and downstream stators, while flow downstream of stators is mostly perturbed in time by rotor effects. As expected, high lift profiles are significantly sensitive to incidence variation, with this effect further enhanced by the multistage cumulative interactions.
Optimisation Techniques Applied to the Design of Gas Turbine Blades Cooling Systems
Authors: Bucchieri Gianlorenzo , Galbiati Massimo , Coutandin Daniele , Zecchi Stefano
Published: September 2008

This paper addresses the methodology used to design the layout of the tip cooling nozzles of a high pressure rotor blade turbine. The methodology used is through a complete CAE approach, by means of a parametric CFD model which is run several times for the exploration of several designs by an optimizer. Hence the design is carried out automatically by parallel computations, with the optimization algorithms taking the decisions rather than the design engineer. The engineer instead takes decision regarding the physical settings of the CFD model to employ, the number and the extension of the geometrical parameters of the blade tip holes and the optimization algorithms to be employed. From CFD validation the final design of the tip cooling geometry found by the optimizer has proved to be better than the base design, which used mean values of all input parameters, and than the design proposed by an experienced heat transfer AVIO engineer, who used standard best practice methods. Furthermore the large number of experiences gained by the simulations run by the optimizer allowed the designer to find laws, functions and correlation between input parameters and performance output, with a further and deeper insight on this specific design problem.
Effect of Wake Passing on Unsteady Aerodynamic Performance in a Turbine Stage
Authors: Yamada K. , Funazaki K. , Hiroma K. , Tsutsumi M. , Hirano Y. , Matsuo A.
Published: September 2008

In the present work, unsteady RANS simulations were performed to clarify several interesting features of the unsteady three-dimensional flow field in a turbine stage. The unsteady effect was investigated for two cases of axial spacing between stator and rotor, i.e. large and small axial spacing. Simulation results showed that the stator wake was convected from pressure side to suction side in the rotor. As a result, another secondary flow, which counter-rotated against the passage vortices, was periodically generated by the stator wake passing through the rotor passage. It was found that turbine stage efficiency with the small axial spacing was higher than that with the large axial spacing. This was because the stator wake in the small axial spacing case entered the rotor before mixing and induced the stronger counter-rotating vortices to suppress the passage vortices more effectively, while the wake in the large axial spacing case eventually promoted the growth of the secondary flow near the hub due to the migration of the wake towards the hub.
Influence of Surface Roughness on Turbine Nozzle Profile Loss and Secondary Loss
Authors: Matsuda Hisashi , Otomo Fumio , Kawagishi Hiroyuki , Inomata Asako , Niizeki Yoshiki , Sasaki Takashi
Published: September 2008

The effects of surface roughness of both nozzle and end-wall on a turbine nozzle performance were investigated experimentally using liner cascade wind tunnel facility under the Reynolds number (Re) condition of Re = 0.3∼1.0 × 106. With buffing, milling, sand blasting and shot blasting, the total of seven levels of the model surface roughness were realized. In order to clarify the effect of the nozzle surface roughness on the profile loss, total pressure losses were measured using three-hole probe for different levels of the surface roughness. It became clear the nozzle profile loss increases as Reynolds number increases for larger roughness group. In addition, it appeared the profile loss depends on not only maximum value of the surface roughness but also roughness conditions. In order to examine the effect of surface roughness on the secondary flow loss, spatial total pressure field of the secondary flow region was measured using three-hole probe for the cases of smooth or rough nozzle surface with smooth or rough end-wall. The secondary flow structures were recognized at the 5∼10% span-wise height region of the suction surface of the nozzle for all cases. With increasing the nozzle surface roughness, not only the profile loss but also net secondary flow loss increases, which is defined as the difference between the total pressure loss and the profile loss in the secondary flow region. However, increase of the end-wall roughness has higher effect on the net secondary flow loss increase. Difference of the effect between the nozzle surface roughness and the end-wall roughness on the nozzle secondary flow loss was discussed.
The Effect of Stator-Rotor Hub Sealing Flow on the Mainstream Aerodynamics of a Turbine
Authors: Reid Kevin , Denton John , Pullan Graham , Curtis Eric , Longley John
Published: September 2008

An investigation into the effect of stator-rotor hub gap sealing flow on turbine performance is presented. Efficiency measurements and rotor exit area traverse data from a low speed research turbine are reported. Tests carried out over a range of sealing flow conditions show that the turbine efficiency decreases with increasing sealant flow rate but that this penalty is reduced by swirling the sealant flow. Results from time-accurate and steady-state simulations using a three-dimensional multi-block RANS solver are presented with particular emphasis paid to the mechanisms of loss production. The contributions toward entropy generation of the mixing of the sealant fluid with the mainstream flow and of the perturbed rotor secondary flows are assessed. The importance of unsteady stator wake/sealant flow interactions is also highlighted.
Reducing the Perfomance Penalty Due to Turbine Inter-Platform Gaps
Authors: Reid Kevin , Denton John , Pullan Graham , Curtis Eric , Longley John
Published: September 2008

Individual nozzle guide vanes (NGV’s) in modern aero engines are often cast as a single unit with integral hub and casing endwalls. When in operation there is a leakage flow through the chord-wise inter-platform gaps. A previous paper [1] has shown that these gaps can result in a stage efficiency penalty of 0.5% – 1.5% depending upon how well they are sealed. In this paper, numerical calculations are used to re-design the inter-platform gaps with the aim of minimizing their effect on the mainstream aerodynamics and hence reduce the efficiency penalty. Experiments using a full scale, low speed model turbine and an improved gap-design show that significant performance improvements are possible regardless of how well the gap is sealed.
Numerical Analysis of an Industrial Fan Fitted With Noise Reduction Devices
Authors: Amaral Teixeira J. , Naylor E. , Ivey P. C. , Sheard A. G. , Kinghorn I. R.
Published: September 2008

The reduction of noise emitted by industrial low speed cooling fans, particularly those fitted to air conditioning systems is a concern to fan manufacturers. The market for industrial low speed fans is highly competitive, with fan noise being the major differentiating factor between competing products. Noise reduction strategies are therefore implemented but these can adversely affect the fan’s pressure delivery capability. A reduction of fan speed can also lead to a reduction in fan noise but this is usually accompanied by a corresponding reduction in pressure rise and flow rate. The practical difficulties associated with maintaining fan pressure and flow characteristics while simultaneously reducing fan noise present fan manufacturers with a challenge. Traditional empirical approaches to the reduction of fan noise have almost been exhausted and no longer offer the potential to significantly reduce fan noise. The understanding of the aerodynamic mechanisms that act as broadband noise sources in low speed fans has been the subject of a considerable number of papers over many years. For most fans operating as a single blade row, the main sources of noise, other than those dependent on the incident turbulence levels, depend on the trailing edge and tip gap flow conditions. A range of strategies seeking to control the noise generated by these regions have been proposed over time by various authors and a number of these schemes have reached production status. The current paper details the numerical analysis of an industrial low speed fan, commonly used in conjunction with a cooling matrix, and which incorporates two distinct noise reduction features; trailing edge crenulations and a blade tip fence. Comparisons are carried out between various combinations of blades, with and without the individual features, and a discussion of the aerodynamics of the particular configurations is undertaken from a perspective of their noise reduction capabilities.
Influence of Loading Distribution on the Off-Design Performance of High-Pressure Turbine Blades
Authors: Corriveau D. , Sjolander S. A.
Published: September 2008

Linear cascade measurements for the aerodynamic performance of a family of three transonic, high-pressure (HP) turbine blades have been presented previously by the authors. The airfoils were designed for the same inlet and outlet velocity triangles but varied in their loading distributions. The previous papers presented results for the design incidence at various exit Mach numbers, and for off-design incidence at the design exit Mach number of 1.05. Results from the earlier studies indicated that by shifting the loading towards the rear of the airfoil an improvement in the profile loss performance of the order of 20% could be obtained near the design Mach number at design incidence. Measurements performed at off-design incidence, but still at the design Mach number, showed that the superior performance of the aft-loaded blade extended over a range of incidence from about −5.0° to +5.0° relative to the design value. For the current study, additional measurements were performed at off-design Mach numbers from about 0.5 to 1.3 and for incidence values of −10.0°, +5.0° and + 10.0° relative to design. The corresponding Reynolds numbers, based on outlet velocity and true chord, varied from roughly 4 × 105 to 10 × 105. The measurements included midspan losses, blade loading distributions and base pressures. In addition, two-dimensional Navier-Stokes computations of the flow were performed to help in the interpretation of the experimental results. The results show that the superior loss performance of the aft-loaded profile, observed at design Mach number and low values of off-design incidence, does not extend readily to off-design Mach numbers and larger values of incidence. In fact, the measured midspan loss performance for the aft-loaded blade was found to be inferior to, or at best equal to, that of the baseline, mid-loaded airfoil at most combinations of off-design Mach number and incidence. However, based on the observations made at design and off-design flow conditions, it appears that aft-loading can be a viable design philosophy to employ in order to reduce the losses within a blade row provided the rearward deceleration is carefully limited. The loss performance of the front-loaded blade is inferior or at best equal to that of the other two blades for all operating conditions. As such, it appears that there is no advantage to front loading the airfoil for transonic high-pressure turbine blades. The results also provide a significant addition to the data available in the open literature on the off-design performance of transonic HP turbine airfoils.
Effects of 3D Aerofoil Tip Clearance Variation on a 4-Stage Low Speed Compressor
Authors: Tschirner Thomas , Johann Erik , Mu¨ller Ralf , Vogeler Konrad
Published: September 2008

The aerodynamic performance of turbomachinery airfoils and therefore the overall efficiency of an engine are strongly dependent on the design of the near end wall sections of blades and vanes. In addition, good compressor stability can only be achieved if the running clearance is as small as allowed for save operation. In the engine the radial gap varies in size due to thermal effects and deterioration as well as transient maneuvers. Since the width of the running clearance can hardly be reduced, a new aerofoil sectional design for cantilevered vanes has been introduced to improve compressor stability over its whole range of operation. The baseline design and the new improved concept have been tested and analyzed for different clearance widths (TC1…TC3) on the rotor blades and cantilevered stator vanes. A baseline configuration featuring two-dimensional airfoils has been used as a datum to develop a more advanced design applying sweep and dihedral at the stacking axes. The running clearance on rotors and the radial gap on cantilevered stators were increased in three steps. Both numerical and experimental investigations had been carried out to verify the effect of variable running clearances on modified end wall sections. Experimental and numerical investigations have shown the effect of bow and sweep within this low speed application does not fully support the common theory of unloaded end wall sections as discussed in various publications. For the 2D blade design the common theory has been proven by both numerical and experimental evaluations. The 3D blade design configuration (BUILD IX) features a significant difference in efficiency sensitivity due to tip clearance width variation, whereas numerical prediction suggested improved compressor performance and stability. Measurements has shown higher losses at this configuration.
Causes of Acoustic Resonance in a High-Speed Axial Compressor
Authors: Hellmich Bernd , Seume Joerg
Published: September 2008

Non-harmonic acoustic resonance was detected in the static pressure and sound signals in a four-stage high-speed axial compressor when the compressor was operating close to the surge limit. Based on prior research reported in the literature and measurements of the resonance frequency, Mach number of the mean flow, and the axial and circumferential phase shift of the pressure signal during resonance it is shown that the acoustic resonance is an axial standing wave of a spinning acoustic mode with three periods around the circumference of the compressor. This phenomenon occurs only if the aerodynamic load in the compressor is high, because the mode needs a high circumferential Mach number for resonance conditions. Mathematics of existing analyses of acoustic resonances in turbomachinery are complex and have therefore rarely resulted in published examples of good agreement with real engine data. The present paper provides suitable, physically based simplifications of the existing mathematical models which are applicable for modes with circumferential wavelengths of more than two blade pitches and resonance frequencies considerably higher than the rotor speed.
Unsteady Flow Physics and Performance of a One-and-1/2-Stage Unshrouded High Work Turbine
Authors: Behr T. , Kalfas A. I. , Abhari R. A.
Published: September 2008

This paper presents an experimental study of the flow mechanisms of tip leakage across a blade of an unshrouded turbine rotor. It shows the design of a new one-and-1/2-stage, unshrouded turbine configuration, which has been developed within the Turbomachinery Laboratory of ETH Zurich. This test case is a model of a high work (Δh/u2 = 2.36) axial turbine. The experimental investigation comprises data from unsteady and steady probe measurements, which has been acquired around all the bladerows of the one-and-1/2-stage, unshrouded turbine. A newly developed 2-sensor Fast Response Aerodynamic Probe (FRAP) technique has been used in the current measurement campaign. The paper contains a detailed analysis of the unsteady interaction between rotor and stator blade rows, with particular attention paid on the flow in the blade tip region. It has been found that the pressure field of the second stator row has a influence on the development of the tip leakage vortex downstream of the rotor. The vortex is modulated by the stator profiles and shows variation in size and relative position to the rotor trailing edge when it stretches around the stator leading edge. Thereby a deflection of the tip leakage vortex has been observed, which expresses in a varying circumferential distance between two neighboring vortices of ±20% of a rotor pitch. Furthermore, a significant influence of quasi-stationary secondary flow features of the upstream stator row on the secondary flow of the rotor has been detected. The geometry data of the one-and-1/2-stage turbine will be available to the public domain for validation and improvement of numerical tools.
Preliminary Design Concepts for High Solidity Nozzle-Less Radial Inflow Gas Turbine for Small Capacity Gas Turbine Engine
Authors: Mistry Chetan S. , Channiwala S. A.
Published: September 2008

The present work for the design of nozzle-less radial inflow turbine begins with power requirement of 20 kW. Based on the available parameters like temperature, pressure and mass flow rate required for the design are obtained from cycle analysis initially preliminary design of rotor was developed and from the available loss models the efficiency of the turbine was found. After completion of the preliminary design of turbine, it was felt necessary to optimized the result for best efficiency accordingly an analytical study was undertaken to study the influence of different parameters like inlet absolute Mach number, relative exit Mach number, solidity, relative velocity ratio, hub to shroud radius ratio and rotational speed on efficiency. VISUAL BASIC program is developed to study the effect of different parameters on efficiency, for different speed conditions it can be observed that for same solidity and higher speed gives the compact size with less variation in losses and efficiency. The results obtained from analysis also suggest the value of higher solidity but in practical situation that will restrict the flow through runner that by increasing the losses and reducing the efficiency.
An Investigation Using Wavelet Analysis Into Velocity Perturbations Under the Influence of Elevated Freestream Turbulence at Transition Onset
Authors: Hernon Domhnaill M. , Walsh Edmond J. , McEligot Donald M.
Published: September 2008

The development of streamwise-orientated disturbances at transition onset for zero-pressure gradient boundary layer flow under the influence of 1.3% freestream turbulence intensity is presented. The analysis concentrates on the development of turbulent spots and other coherent structures with the use of wavelet analysis. The turbulent spot structure is shown to change dramatically in shape, sign of perturbation velocity and energy content from the near wall region to the boundary layer edge. An increased number of trubulent structures are observed near the boundary layer edge, all with negative perturbation velocities, compared to those of positive perturbation velocity in the near wall region. The wavelet maps demonstrate some interesting features of turbulent spot development including regions of high frequency disturbance growth prior to the spot passing the sensor. Distributions of peak negative, peak positive and averaged perturbation velocities were obtained at three streamwise positions prior to transition onset. As transition onset approached the magnitude of the negative value far exceeded the positive and their relative positions within the boundary layer changed considerably. The results presented in this report give further insight into the physics of pre-transitional flow illustrating the influence of negative perturbation velocity in the transition process. Furthermore, the importance of peak instantaneous perturbations compared to averaged values is also demonstrated, a feature of the flow that computational techniques will have to model in order to accurately predict transition phenomena.
Low Pressure Turbine Design for Rolls-Royce Trent 900 Turbofan
Authors: Gonza´lez Paloma , Lantero Mikel , Olabarria Victor
Published: September 2008

The Trent 900 engine is part of the Rolls-Royce Trent family, designed to enter into service in the Airbus 380, the largest commercial aeroplane in the world. The LP turbine design was performed in Spain by ITP. Since 1997 ITP has been involved in the Rolls-Royce Trent low pressure turbine design. In the Trent 900, ITP is one of the risk and revenue sharing partners of the project. The design team participated extensively in the low pressure turbine since the conceptual design phase to the engine certification that took place in October 2004 and it will continue through aeroplane civil operation. The aim of this paper is to describe some of the most key characteristics of the LP turbine design, mainly focused on the aerodynamics; The LP turbine concept will be shown, the new aerodynamic features described and how they are supported by former experimental data.