bluuebunny/crossref_metadata_2025 huggingface dataset:
Interaction of Rotor and Casing Treatment Flow in an Axial Single-Stage Transonic Compressor With Circumferential Grooves
Authors: Mu¨ller Martin W. , Biela Christoph , Schiffer Heinz-Peter , Hah Chunill
Published: August 2009

The influence of circumferential grooves on the tip flow field of an axial single-stage transonic compressor rotor has been examined experimentally and numerically. The compressor stage provides a strongly increased stall margin with only small penalties in efficiency when the casing treatment is applied. Due to the complex interactions of the grooves with the rotor flow, unsteady measurement techniques have been chosen as an attempt to identify the aerodynamic effects responsible for the operating range extension. Therefore, the casing treatment has been instrumented with piezoresistive pressure sensors in the land between the grooves providing high-resolution static wall pressure measurements at different operating conditions. Data acquisition worked at a sampling rate of 125kHz, providing around 23 static pressure values per blade passage at 11 axial positions at the nominal speed of 20,000 rpm. A comparable dataset, but with 14 sensors, was obtained for the smooth casing. The results show the fluctuation of the tip leakage vortex and shock-vortex-interactions as well as the changed situation with casing treatment. Ensemble-averaged data shows tip leakage vortex trajectories. At near stall conditions with the smooth casing, the vortex hits the front part of the adjacent blade, which indicates the possibility of a spill forward of low momentum fluid into the next passage. Standard deviation values prove a high fluctuation of the pressure field over the tip gap. When the casing treatment is applied, the vortex trajectory maintains alignment along the blade’s suction side, thus preventing the onset of rotating stall. Results are presented as a back-to-back comparison of the smooth casing versus the treated casing at three operating conditions: peak efficiency at a mass flow rate of m˙pe = 16.2kg/s, near stall of the smooth casing at m˙nssc = 14.0kg/s and near stall of the treated casing at m˙ns = 12.6kg/s. Steady and unsteady numerical simulations of the rotor-only flow field have been calculated with and without grooves. These calculations aim at a broad analysis of the occurring flow phenomena at the rotor tip. Tip leakage flow behaviour and vortex trajectories are discussed in detail by summarizing the congruent findings of both numerical and experimental investigations.
Water Film Formation on an Axial Flow Compressor Rotor Blade
Authors: Nikolaidis Theoklis , Pilidis Periclis , Teixeira J. A. , Pachidis V.
Published: August 2009

A numerical approach was used to evaluate the liquid water film thickness and its motion on an axial flow compressor rotor blade under water ingestion conditions. By post-processing blading data and using computer programs to create the blades and their computational grid, the global computational domain of the first stage of an axial flow compressor was built. The flow field within the domain was solved by CFX-Tascflow, which is a commercial CFD code commonly used in turbomachinery. The computational domain consists of an extended inlet, an inlet guide vane, a rotor and a stator blade. Having solved the flow field at Design Point, the inlet guide vane blade was re-positioned to account for changes in idle speed. At that speed, the effects of water ingestion are expected to be more significant on gas turbine engine performance. Several cases with water ingestion were studied, changing parameters like water mass and compressor rotational speed. A FORTRAN computer program was created to calculate the water film height and speed. The extra torque needed by the compressor to keep running at the same rotational speed, was also calculated. The considerable increase in torque was confirmed by experimental observations according to which water ingestion had a detrimental effect on gas turbine operation.
Heat Transfer and Flow Characteristics of a Leading Edge Impingement Cooling System for Low Pressure Turbine Vanes
Authors: Alvarez Jose Javier , de la Calzada Pedro , Krulic Gregory
Published: August 2009

An impingement cooling system for the leading edge of a contra-rotating Power Turbine (PT) representative of a small turboshaft engine was investigated experimentally as well as numerically. The PT vane features a very thin leading edge with high curvature and side channels rapidly facing backwards. Constraints on cooling flow consumption and distribution led to a leading edge (LE) configuration with two rows of staggered jets. This particular configuration was experimentally investigated for three different Reynolds numbers around the design point by using a transient liquid crystal technique. Heat transfer results are presented in terms of surface distributions, impingement rows stagnation line local distributions, streamwise distributions along planes over the impingement stagnation points, span averaged streamwise local distributions and surface averaged values, and then compared to available correlations from existing literature. In order to investigate the flow features of this kind of configuration in more detail, a CFD investigation was also conducted. Results from CFD simulations were compared to the experimental data and then used to discuss in detail the main flow features driving the surface heat transfer.
Assessment of Different Actuator Concepts for Acoustic Boundary Control of a Premixed Combustor
Authors: Bothien Mirko R. , Moeck Jonas P. , Paschereit Christian Oliver
Published: August 2009

In the design process, new burners are generally tested in combustion test rigs. With these experiments, computational fluid dynamics, and finite element calculations, the burners’ performance in the full-scale engine is sought to be predicted. Especially, information about the thermoacoustic behavior and the emissions is very important. As the thermoacoustics strongly depend on the acoustic boundary conditions of the system, it is obvious that test rig conditions should match or be close to those of the full-scale engine. This is, however, generally not the case. Hence, if the combustion process in the test rig is stable at certain operating conditions, it may show unfavorable dynamics at the same conditions in the engine. In a previous paper (GT2007-27796), the authors introduced an active control scheme which is able to mimic almost arbitrary acoustic boundary conditions. Thus, the test rig properties can be tuned to correspond to those of the full-scale engine. The acoustic boundary conditions were manipulated using woofers. In the present study, proportional valves are investigated regarding their capabilities of being used in the control scheme. It is found that the test rig impedance can be tuned equally well. In contrast to the woofers, however, the valves could be used in industrial applications, as they are more robust and exhibit more control authority. Additionally, the control scheme is further developed and used to tune the test rig at discrete frequencies. This exhibits certain advantages compared to the case of control over a broad frequency band.
Compliant Foil Journal Bearing Performance at Alternate Pressures and Temperatures
Authors: Bruckner Robert J. , Puleo Bernadette J.
Published: August 2009

An experimental test program has been conducted to determine the highly loaded performance of current generation gas foil bearings at alternate pressures and temperature. Typically foil bearing performance has been reported at temperatures relevant to turbomachinery applications but only at an ambient pressure of one atmosphere. This dearth of data at alternate pressures has motivated the current test program. Two facilities were used in the test program, the ambient pressure rig and the high pressure rig. The test program utilized a 35 mm diameter by 27 mm long foil journal bearing having an uncoated Inconel X-750 top foil running against a shaft with a PS304 coated journal. Load capacity tests were conducted at 3, 6, 9, 12, 15, 18, and 21 krpm at temperatures from 25°C to 500°C and at pressures from 0.1 to 2.5 atmospheres. Results show an increase in load capacity with increased ambient pressure and a reduction in load capacity with increased ambient temperature. Below one-half atmosphere of ambient pressure a dramatic loss of load capacity is experienced. Additional lightly loaded foil bearing performance in nitrogen at 25°C and up to 48 atmospheres of ambient pressure has also been reported. In the lightly loaded region of operation the power loss increases for increasing pressure at a fixed load. Knowledge of foil bearing performance at operating conditions found within potential machine applications will reduce program development risk of future foil bearing supported turbomachines.
A Genetic Algorithm Approach to Estimate Performance Status of Gas Turbines
Authors: Li Y. G.
Published: August 2009

Accurate estimation of performance status of a gas turbine engine at certain ambient and operating condition based on measured gas path parameters is very important for both engine designers and users alike. It could be a very challenging task for engine performance engineers to estimate the value of component design parameters in order to match measured gas path parameters when the number of design point component parameters and the number of measurable performance parameters become large. Such status estimation can be used to distinguish the performance difference among fleet engines and build accurate engine models at an artificial design point for individual engines, which is also crucially important for gas path diagnostic analysis. In this paper, a gas turbine design point performance adaptation approach based on the integration of gas turbine thermodynamic performance modelling and a Genetic Algorithm has been developed in order to estimate the design point component parameters and match the available gas path measurements of real engines. In the approach, the initially unknown component parameters may be compressor pressure ratios and efficiencies, turbine entry temperature, turbine efficiencies, air mass flow rate, cooling flows, by-pass ratio, etc. The engine measurable performance parameters may be thrust and specific fuel consumption for aero engines, shaft power and thermal efficiency for industrial engines, gas path pressures and temperatures, etc. The developed adaptation approach has been applied to a design point performance status estimation of an industrial gas turbine engine GE LM2500+ operating in Manx Electricity Authority (MEA), UK. The application shows that the adaptation approach is very effective and robust in producing a model engine that matches the actual engine performance with acceptable computation speed. Theoretically the developed techniques can be applied to different gas turbine engines.
Application of Numerical Analysis Tools for Airframe-Propulsion Integrated Test and Evaluation
Authors: Davis Milt W. , Hale Alan A. , Vining Charles , Cousins William T.
Published: August 2009

With increasing emphasis on streamlining the acquisition process, ground-test centers like the Arnold Engineering Development Center (AEDC) are re-evaluating their roles in the development of aerospace systems. Instead of merely providing data from ground-test facilities, the new emphasis challenges AEDC to become a team member that provides knowledge for risk management and decision making during the development and operation of an aerospace system. AEDC has developed an Integrated Test and Evaluation (IT&E) approach to support aerospace system development efforts. Two primary aircraft propulsion subsystems are the inlet and the engine. Traditionally these subsystems have been designed, analyzed, and tested as isolated systems although some integrated testing has been performed. The interaction between the subsystems is modeled primarily through evaluating inlet distortion in an inlet test and then simulating this distortion in engine tests via screens or similar devices. Recently, it has been recognized that significant improvements in both performance and operability can be realized when both the inlet and the engine are designed, tested and analyzed with full knowledge of the other. To aid in the airframe-propulsion integration process, numerical solutions of the forebody-inlet and engine compression system have been coupled and partially demonstrated in a limited manner to show the potential for enhancing the acquisition process.
Automated Methods to Calibrate a High-Fidelity Thrust Deck to Aid Aeropropulsion Test and Evaluation
Authors: Monaco Jeffrey F. , Kidman David S. , Malloy Donald J. , Ward David G. , Gist James F.
Published: August 2009

Modeling and simulation is regularly used to support test and evaluation at the U.S. Air Force aeropropulsion ground and flight test centers. Advanced modeling and simulation tools used to calibrate a high-fidelity thrust program that predicts steady-state and transient operation of a fighter aircraft turbofan engine are discussed in this paper. Maximum likelihood parameter estimation is used to estimate engine-model calibration factors. Empirical modeling techniques are then used to obtain closed-form polynomials characterizing the variations in the calibration factors as functions of engine operating condition. The inferred polynomial relationships are used to constrain the functional form of the calibration factors as part of a final maximum likelihood parameter estimation step. The coefficients of the polynomials are estimated over all ground-test cases simultaneously to minimize the weighted error between ground-test output measurements and thrust-deck predictions. An overview of the modeling and simulation framework is given with examples illustrating how the framework can readily be applied to a diverse range of aeropropulsion test and evaluation tasks. Details of the thrust deck, key variables, and the calibration factors are presented. Calibrated thrust deck predictions are compared with steady-state ground test data; and it is shown how the approach advances the state-of-the-art through methods that are more highly automated and intuitive to the engine analyst. The enhancements that automate the steps to set up a modeling task provide tangible benefits for a variety of test and evaluation applications.
Active Flow Control in a Single-Stage Axial Compressor Using Tip Injection and Endwall Boundary Layer Removal
Authors: Dobrzynski B. , Saathoff H. , Kosyna G. , Clemen C. , Gu¨mmer V.
Published: August 2009

Different active flow control techniques have been investigated in a 1.5-stage axial-flow compressor. Looking at a low-speed single-stage environment, many researchers have shown that highly loaded compressors are tip critical, showing stall inception caused by short length scale disturbances (spikes). It has been shown by several authors that these disturbances are related to the spillage of endwall flow ahead of the blading (spill forward). For the present work, different tip injection configurations were investigated in order to stabilize the near casing flow, increasing the operating range of the compressor. Stall margin improvement and the impact on stage efficiency are compared and discussed. Oil flow pictures of the casing wall above the rotor and of the stator blades as well as traverse data from pneumatic 5-hole probes show the impact of flow control on rotor and stator performance. Another method of energizing the casing wall boundary layer is the removal of low energy fluid by a circumferential slot above the rotor, which was also studied experimentally. Again, the impact on compressor operating range and efficiency, as well as flow field information collected by oil flow visualization and traverse data are discussed. Comparing the different flow control techniques, it is shown that increasing stall margin is not directly linked to stage efficiency. As described in various publications, discrete tip injection is a very powerful technique as far as range extension is concerned, but it also has substantial drawbacks such as the circumferential inhomogeneity of the rotor exit flow. These inhomogeneities may result in poor stator performance, overall resulting in a drop of stage efficiency. This problem does not occur if circumferential boundary layer removal above the rotor is used. This method however shows much less potential for increasing the operating range.
Aero-Thermal Investigation of Tip Leakage Flow in a Film Cooled Industrial Turbine Rotor
Authors: Krishnababu S. K. , Hodson H. P. , Booth G. D. , Lock G. D. , Dawes W. N.
Published: August 2009

A numerical investigation of the flow and heat transfer characteristics of tip leakage in a typical film cooled industrial gas turbine rotor is presented in this paper. The computations were performed on a rotating domain of a single blade with a clearance gap of 1.28% chord in an engine environment. This standard blade featured two coolant and two dust holes, in a cavity-type tip with a central rib. The computations were performed using CFX 5.6, which was validated for similar flow situations by Krishnababu et al., [18]. These predictions were further verified by comparing the flow and heat transfer characteristics computed in the absence of coolant ejection with computations previously performed in the company (SIEMENS) using standard in-house codes. Turbulence was modelled using the SST k-ω turbulence model. The comparison of calculations performed with and without coolant ejection has shown that the coolant flow partially blocks the tip gap, resulting in a reduction of the amount of mainstream leakage flow. The calculations identified that the main detrimental heat transfer issues were caused by impingement of the hot leakage flow onto the tip. Hence three different modifications (referred as Cases 1 to 3) were made to the standard blade tip in an attempt to reduce the tip gap exit mass flow and the associated impingement heat transfer. The improvements and limitations of the modified geometries, in terms of tip gap exit mass flow, total area of the tip affected by the hot flow and the total heat flux to the tip, are discussed. The main feature of the Case 1 geometry is the removal of the rib and this modification was found to effectively reduce both the total area affected by the hot leakage flow and total heat flux to the tip while maintaining the same leakage mass flow as the standard blade. Case 2 featured a rearrangement of the dust holes in the tip which, in terms of aero-thermal-dynamics, proved to be marginally inferior to Case 1. Case 3, which essentially created a suction-side squealer geometry, was found to be inferior even to the standard cavity tip blade. It was also found that the hot spots which occur in the leading edge region of the standard tip and all modifications contributed significantly to the area affected by the hot tip leakage flow and the total heat flux.
Forced Flame Behaviour in a Spray Combustor
Authors: Arthur David M. , Dowling Ann P. , Cant R. Stewart , Zhu Min
Published: August 2009

The stability of a liquid-fuelled annular aeroengine gas turbine was investigated using a parallelised three-dimensional uRANS code. The simulations were performed on a single sector of an idealised annular gas turbine geometry using a non-orthogonal body-fitted structured grid. A κ-ε turbulence closure model was used and a Lagrangian fuel spray was solved using a Monte Carlo solution of Williams’ spray equation. Mass flow rate perturbations were imposed on the atomiser and the dilution ports of the combustor. The effect of the imposed perturbations on heat release was quantified using flame transfer functions. Excitation of either atomiser or dilution flows led to significant unsteady combustion. With careful design, fluctuations in the rate of combustion due to unsteady flow through the atomiser and dilution ports can cancel, leading to a more stable system.
The Numerical Study of a New Air-Cooled Turbine Blade With Depression Feature at Leading Edge
Authors: Wang Songtao , Wang Kai , Qiang Xiaoqing , Wang Zhongqi
Published: August 2009

A new air-cooled turbine blade with depression feature at leading edge was presented in this paper. The flow field of this blade at the transonic condition was simulated in detail. The change of vortex position induced by the leading edge cooling air injection along the blade profile was observed. The aerodynamic characteristics of the new blade such as film effectiveness, energy loss, pressure field etc. were compared with the traditional air-cooled turbine blade. The result shows that this new blade locks the vortex induced by the leading edge cooling air injection in its depression part. The mixing of cooling air and mainstream becomes weak. Thus, the number of the leading edge cooling air injection holes can be reduced. Correspondingly, the amount of injection can also be reduced.
Simulation and Optimization of an Adjustable Inlet Guide Vane for Industrial Turbo Compressors
Authors: Hensges Michael
Published: August 2009

To investigate the kinematics and dynamics of an adjustable inlet guide vane mechanism (IGV) for industrial turbo compressors, an IGV was modeled as a multibody system (MBS) consisting of elastic interconnections and rigid bodies. Besides investigating the IGV kinematics, its vibrations and structural strength were also verified numerically. The kinematic analyses enabled the design to be optimized in terms of undesirable collisions between the interconnected bodies. The pressure exerted on the guide vanes, which is calculated by CFD simulations, forms a set of forces and torques for each blade. These sets were created for two different performance maps, referred to in the following as Gas I and Gas II. Calculating the desired drive torque, joint reaction forces and the driving ring’s displacements were the essential inputs for the dynamic multibody analyses performed. These investigations showed that the desired torque to drive the mechanism is governed by the sliding element’s friction forces. The gas forces were able to raise the torque by roughly 6% and 32.6% for Gas I and II, respectively. Due to uncertainties in the determination of the friction coefficients, the highest expected values were taken into account for selecting an accurate actuator for the IGV. The strength and vibration analyses were carried out using the Finite Element (FE) Method. All computed critical natural frequencies of the IGV can be empirically considered to be highly damped resonances in the actual system due to joint friction effects. Reaction forces determined by the dynamic multibody analyses were transferred as loads to the FE model. In most cases, the joint reaction forces have been so low that no further investigations were necessary. Hence, verification of strength was carried out using a contact FE analysis for the highest loading condition between the assembly and the pin, which transfers the entire drive force from a lever into the driving ring.
Unsteady Pressure Measurement in a Single Stage Axial Transonic Compressor Near the Stability Limit
Authors: Biela Christoph , Mu¨ller Martin W. , Schiffer Heinz-Peter , Zscherp Carsten
Published: August 2009

With the help of piezoelectric high frequency pressure probes measurements are undertaken to investigate the flow during stable compressor operation close to the stability limit. Fourteen static pressure probes record the static wall pressure and ten total pressure probes record the total pressure at the rotor exit, both in the absolute frame of reference. The data is then visualised as ensemble averaged contour and spectrum plots. With the help of wall and exit pressure, the tip leakage vortex is localised. Oscillations of the tip leakage vortex are seen as well in terms of high relative standard deviation as well as in an excitation of a frequency band around 1/2 BPF. Further investigation of the frequency spectrum with the help of the pseudo-unsteady wall pressure reveal the occurrence of rotating tip leakage vortex disturbances forming a two-passage periodic vortex pattern. The presented measurements were obtained using Rotor-1 from the TU Darmstadt rotor family. With a sampling rate of 125kHz the pressure field is resolved with 23 measurements per passage (at 20.000 rpm, design speed).
Heat/Mass Transfer in Rotating, Smooth, High Aspect-Ratio (4:1) Coolant Channels With Curved Walls in 90° and 45° Orientation
Authors: Sethuraman Eashwar , Nikitopoulos Dimitris E. , Acharya Sumanta
Published: August 2009

The paper presents an experimental study of heat/mass transfer coefficient in 4:1 aspect ratio smooth channels with non-uniform cross-sections. Curved leading and trailing walls are studied for two curvatures and for two different curvature configurations. One configuration has curved walls with curvature corresponding to the blade profile (positive curvature on both leading and trailing walls), and the other configuration has leading and trailing walls that curve inwards into the coolant passage (negative curvature on the leading surface and positive curvature on the trailing surface). The experiments are conducted in a rotating two-pass coolant channel facility using the naphthalene sublimation technique. Only the radially outward flow is considered for the present study at a Reynolds number of 20,000, rotation numbers in the range 0–0.051 and for 90° and 45° orientations with respect to the direction of rotation. In addition to area averaged values, the span-wise mass transfer distributions of fully developed regions of the channel walls are presented in order to delineate the effects of rotation number, channel shape and orientation. The mass transfer data from the curved wall channels are compared to those from a smooth 4:1 rectangular duct with similar flow parameters. Pressure drop along each channel is also measured and put in perspective of the mass/heat transfer results for the 90° orientation, through the presentation of a performance factor.
Relight Envelope of a Military Gas Turbine Engine: An Experimental Study
Authors: Mishra R. K. , Gouda G. , Vedaprakash B. S.
Published: August 2009

A twin spool low bypass turbofan engine under development and its combustor in full-scale were tested independently at altitude conditions to establish the relight envelope of the engine. Demonstration of relight capability and defining its boundary are mandatory for military gas turbine engines and for single engine application in particular. The engine was first subjected to windmill to establish its windmilling characteristics. The full engine was then tested for light-off in an altitude test facility simulating windmilling conditions from 4 to 12 km altitude with flight Mach numbers from 0.2 to 1.0. The relight boundary is defined based on the successful light-off points achieved from engine tests. Similar tests were carried out on the full-scale combustion chamber in a stand-alone mode simulating altitude conditions at engine flame-out. The combustor test has defined the light-off and lean blow out limits of the at each point on the relight boundary. The information of fuel-air ratio at light-off and blow-out is very useful in setting the engine fuel schedule for altitude operation and relight. In this paper an attempt is made to highlight various tests carried out on engine and its combustor to define the relight boundary of the engine. The paper also emphasizes the experience of combustor development and associated problems in meeting the relight challenges of military engines. These problems include the necessity of higher fuel-air ratio at high altitudes, the role of additional localized fuel injection through start-up atomizers, and effect of single igniter on relight characteristics. The relight envelope demonstrated by the engine is very satisfactory to meet the first flight requirement where the flight mission generally concentrate in the zone of 0.6 to 0.8 Mach and altitude does not exceed 10 to 12 km. Combustor and atomizer modification is needed to improve relight performance and to shift the boundary to further left.
Breaking the Swirl With Brush Seals: Numerical Modeling and Experimental Evidence
Authors: Helm Peter , Pugachev Alexander , Neef Matthias
Published: August 2009

Striving for smaller losses in turbomachinery has led to many advancements in the design of seals. Modern sealing concepts such as brush seals hold a great potential to increase the efficiency of both flight engines and stationary turbines. At the same time, in order to maintain stable operating conditions of the rotor, swirl-induced forces must be kept at a minimum in the sealing channels. Therefore, the influence of the permeable and flexible bristle pack of brush seals on the flow around the rotor surface must be known. In this paper the swirl flow in the cavities of two different seal geometries is studied experimentally and numerically. A conventional three-tooth labyrinth serves as a reference. A second seal arrangement with a bristle pack upstream of two teeth is compared with the reference labyrinth. The swirl is evaluated experimentally from total and static pressure measurements in various axial and circumferential positions. Additionally, the axial swirl distribution is calculated using computational fluid dynamics (CFD). Here, the numerical model of the brush seal is based on the porous medium approach and is calibrated using the experimental values of the leakage and the bristle clearance by adjusting the thickness of the bristle pack. The calibrated CFD model is then used to study the impact of the brush seal on the swirl component of the sealing flow. The observed significant decrease of the swirl by the brush seal shows good agreement with the experimental data. The impact of changes in bristle pack clearance on the swirl is also investigated and compared with experimental evidence. The aim is to show that the brush seals have a natural tendency to interrupt seal swirl. They can therefore be used for swirl control in order to create a beneficial impact on the dynamic stability of turbomachines.
Broadband Entropy Noise Phenomena in a Gas Turbine Combustor
Authors: Fischer Andre´ , Bake Friedrich , Ro¨hle Ingo
Published: August 2009

Broadband sound emission in the region of 1 kHz to 3 kHz probably related to entropy noise was found in a model gas turbine combustor. Entropy noise is caused by the acceleration of fluctuating hot spots in the outlet nozzle of aero-engine combustors. It was predicted in literature in the late 70’s. However, the entire relevance of entropy noise concerning the total noise emission of a combustor is not confirmed until now. Disagreements still exist in the dominating propagating frequencies. This paper presents a broadband noise phenomenon appearing from 1 kHz to 3 kHz in a model aero-engine combustor. In this frequency range the sound pressure level depends on the Mach number as the entropy noise theory expected. Experiments are performed for different types of combustion chambers varying in length and material at various power consumptions and equivalent ratios. A scale factor had been chosen to double the length of the combustion chamber, to allow the investigation of different states of temperature mixing in the exhaust flow following the primary flame zone. Due to the decomposition of the acoustic field into forward and backward propagating waves only the emitted sound power can be considered in the analysis. It can be shown that the broadband noise seems to be independent of changes in geometry or material. A similar cold test flow condition was investigated with respect to the broadband noise phenomena and compared to the reactive case.
Optimization of a Channel Roughened by Dimples on Opposite Surfaces for Heat Transfer Enhancement
Authors: Samad Abdus , Lee Ki-Don , Kim Kwang-Yong
Published: August 2009

Shape optimization of a channel with both walls roughened by staggered arrays of dimples is performed to enhance turbulent heat transfer compromising with friction drag. The dimpled channel shape is defined by three geometric design variables, and the design points within design space are selected using Latin hypercube sampling. The shape of the channel is optimized with 3-D Reynolds-averaged Navier-Stokes analysis and surrogate approximation methods. A weighted-sum method for multi-objective optimization is applied to integrate multiple objectives related to heat transfer and friction loss into a single objective. A weighted average surrogate (PBA) model which is constructed by averaging polynomial response surface approximation, Kriging and radial basis neural network surrogate models is used. By the optimization, the objective function value is improved largely, and heat transfer rate is increased much higher than pressure loss increase due to shape deformation. The optimum design produces lower channel height, wider dimple spacing, and deeper dimple. The flow mechanism shows the heat transfer rate is increased mainly in rear portion of the dimple.
The Influence of Dump Gap on External Combustor Aerodynamics at High Fuel Injector Flow Rates
Authors: Walker A. Duncan , Carrotte Jon F. , McGuirk James J.
Published: August 2009

The increasing demand to reduce fuel burn, hence CO2 emissions, from the gas turbine requires efficient diffusion to reduce the system pressure loss in the combustor. However, interactions between pre-diffuser and combustor can have a significant effect on diffuser performance. For example, the consequence of increased fuel injector flow at a dump gap set using conventional design guidelines has been shown [2] to introduce a destabilising interaction between fuel injector and upstream components. The present paper concentrates on examining the effects of increased dump gap. Dump gap ratios of 0.8, 1.2 and 1.6 were employed, with each test utilising the same IGV, compressor rotor, integrated OGV/pre-diffuser and dump geometry. The flow fraction of compressor efflux entering the combustor cowl was set to be representative of lean combustors (50% – 70%). Measurements were made on a fully annular rig using a generic flametube with metered cowl and inner/outer annulus flows. The results demonstrate that, with fixed cowl flow, as the dump gap increases component interactions decrease. At a dump gap ratio of 0.8, the proximity of the flametube influences the pre-diffuser providing a beneficial blockage effect. However, if increased to 1.2, this beneficial effect is weakened and the pre-diffuser flow deteriorates. With further increase to 1.6 the pre-diffuser shows strong evidence of separation. Hence, at the dump gaps probably required for lean module injectors it is unlikely the pre-diffuser will be influenced beneficially by the flametube blockage; this must be taken into account in the design. Further, with small dump gaps and high cowl flow fraction, the circumferential variation in cowl flow can feed upstream and cause OGV/rotor forcing. At larger dump gaps the circumferential variation does not penetrate upstream to the OGV and the rotor is unaffected. The optimum dump gap and pre-diffuser design for best overall aerodynamic system performance from rotor through to feed annuli is a compromise between taking maximum advantage of upstream blockage effects, whilst minimizing any 3D upstream forcing.
Transonic Fan Tip-Flow Features Revealed by High Frequency Response Over-Tip Pressure Measurements
Authors: Shin Hyoun-Woo , Solomon William , Wadia Aspi
Published: August 2009

Data from an array of high frequency response pressure transducers embedded in the casing wall over the tip of a transonic fan are reported. Phase-lock averaging of data from this array was successful in resolving an axial-tangential map of the static pressure rise in the rotor tip, as has been reported by other workers. Phase-lock ensemble RMS processing of the data is shown to be a useful technique that provides insight into the flow physics around the blade tip. Comparison with CFD results allows for more definite identification of features observed in the data. A complex flow field involving the casing wall boundary layer, the blade shock system and the over-tip leakage flow is observed. Differences between CFD data and measurements are explored by way of computational sensitivity studies. Results are reported for a range of throttle settings and speeds.
Determination of Cooling Parameters for a High Speed, True Scale, Metallic Turbine Vane Ring
Authors: Polanka M. D. , Anthony R. J. , Bogard David G. , Reeder Mark F.
Published: August 2009

Film cooling technology has been around for many decades and many significant advances in cooling effectiveness have been made at many different facilities using several different methods. A large proportion of film cooling research is successfully carried out using simplified scaled-up models in wind tunnels coupled with novel measurement techniques. These tests have been very effective in assessing basic film cooling parameters for many cooling hole geometries, patterns, and blowing ratios. In real engines, however, film cooling designs are ultimately subjected to highly unsteady 3-D secondary flows and rotational effects. Few film cooling experiments have quantified these effects on real, true scale turbine hardware in a rotating test environment. The Turbine Research Facility (TRF) at the Air Force Research Laboratory has been acquiring uncooled heat transfer measurements on full scale metallic airfoils both with and without rotation for several years. The addition of cooling flow to this type of facility has provided new capability, and new challenges. The primary two issues being that the film temperature is unknown and that the airfoil is no longer semi-infinite. This makes it more difficult to extract the adiabatic effectiveness and the heat transfer coefficient from the measurements of surface temperature and surface heat transfer since conventional methods used in most other experiments are not valid in this case. In contrast another cooling parameter, the overall effectiveness, is readily obtained from measurements of surface temperature, internal coolant temperature, and mainstream temperature. The overall effectiveness is a normalized measure of metal surface temperatures expected for actual operating conditions. It is the goal of this paper to evaluate how measurements, obtained from a transient blowdown facility like the TRF, can be used to quantify the expected performance of a film cooled turbine airfoil. Additionally, it is imperative to properly correlate these experimental results to the true engine conditions. The data required for this analysis has been collected using an array of surface mounted thermocouples and thin film gauges in a series of experiments where freestream temperatures and coolant temperatures and mass flow rates were varied. The airfoil used in this investigation was a thin walled metallic airfoil with a showerhead cooling scheme and several rows of normal holes on both the pressure and suction sides of the airfoil. The flow is typical of that seen in a modern high pressure turbine — that is an inlet Mach number of about 0.1 accelerating toward sonic at the throat with a high inlet freestream turbulence level of about 20%.
Predicting Separation and Transitional Flow in Turbine Blades at Low Reynolds Numbers
Authors: Sanders Darius D. , O’Brien Walter F. , Sondergaard Rolf , Polanka Marc D. , Rabe Douglas C.
Published: August 2009

There is increasing interest in design methods and performance prediction for aircraft engine turbines operating at low Reynolds numbers. In this regime, boundary layer separation may be more likely to occur in the turbine flow passages. For accurate CFD predictions of the flow, correct modeling of laminar-turbulent boundary layer transition is essential to capture the details of the flow. To investigate possible improvements in model fidelity, CFD models were created for the flow over two low pressure turbine blade designs. A new three-equation eddy-viscosity type turbulent transitional flow model originally developed by Walters and Leylek was employed for the current RANS CFD calculations. Previous studies demonstrated the ability of this model to accurately predict separation and boundary layer transition characteristics of low Reynolds number flows. The present research tested the capability of CFD with the Walters and Leylek turbulent transitional flow model to predict the boundary layer behavior and performance of two different turbine cascade configurations. Flows over the Pack-B turbine blade airfoil and the midspan section of a typical low pressure turbine (TLPT) blade were simulated over a Reynolds number range of 15,000–100,000, and predictions were compared to experimental cascade results. The turbulent transitional flow model sensitivity to turbulent flow parameters was investigated and showed a strong dependence on free-stream turbulence intensity with a second order effect of turbulent length scale. Focusing on the calculation of the total pressure loss coefficients to judge performance, the CFD simulation incorporating Walters and Leylek’s turbulent transitional flow model produced adequate prediction of the Reynolds number performance for the TLPT blade cascade geometry. Furthermore, the correct qualitative flow response to separated shear was observed for the Pack-B blade airfoil. Significant improvements in performance predictions were shown over predictions of conventional RANS turbulence models that cannot adequately model boundary layer transition.
Optimum Use of Existing Sensor Information for Gas Turbine Diagnostics
Authors: Stamatis A. G.
Published: August 2009

Gas Path Analysis, (GPA), has been proven a powerful tool for Gas Turbine fault detection and isolation. Restriction on GPA effectiveness is mainly due to limited instrumentation imposed by sensor technology and plant integrity. The main concept proposed in the past for overcoming this situation is based on exploiting existing sensor information from different operating conditions. In this paper, potential problems in diagnosis based on methods implementing the above concept are evaluated and a novel method is described. The new method is based on safer assumptions regarding the fault appearance on Gas turbines. The main idea is to exploit different sensor deviation values produced on the same operating point defined through different operating parameters. Application of the proposed method shows significant improvements regarding reliable Gas Turbine Diagnosis.
Design of a New Experimental Rig for Thermal Cyclic Testing of Combustor Liner Tiles
Authors: Mende C. , Liedtke O. , Schulz A. , Bauer H.-J.
Published: August 2009

This paper describes the design and operation of a new test rig, which allows the simulation of real engine operating conditions leading to Low-Cycle Fatigue of combustor liner tiles. The experimental setup will provide high-quality data for the development of damage models. At first the design data of the test rig will be derived from the relevant damage mechanisms in Combustor Liner Tiles (CLT). Then the construction of the test rig and its integration into an existing high temperature high pressure test facility will be elucidated. Finally experimental data of a typical simulated thermal cycle is shown.
Influence of Gap Between Casing and Variable Stator Blade on Axial Compressor Performance
Authors: Ribi Beat , Meyer Michael P.
Published: August 2009

Axial compressors for industrial application are usually assembled from a limited set of standardized stages and casings to meet the various customer specifications. The performance of the entire compressor is then often predicted by a stage stacking method where the characteristics of individual stages (“base characteristics”) are superpositioned. At MAN TURBO these “base characteristics” for a standard blade type are obtained from measurements in a test rig featuring variable stator blades. The geometry is then identical to the one used in compressors designed for fixed speed where the required operating range is obtained by the use of variable stator blades. In contrast, a variable speed already allows for a certain operating range and there is no need for variable stator blades. The fixed stator blades, however, do not include a gap between blade and casing. The question to be addressed in this paper is how this additional gap between stator blade and casing (“stator casing gap”) affects the characteristic of a single stage comprising a rotor and a stator. In order to assess this influence for a variety of geometric parameters such as stator blade settings, chord length and gap height a simple one-dimensional approach for tip leakage loss assessment was adapted and then modified in order to describe the loss mechanisms in a stator casing gap. The relevant mechanism itself and the calibration of the model were inferred from full 3D viscous flow calculations. These calculations were performed for a 8 1/2 stage test compressor, first for the true configuration, i.e., with stator casing gap, and then for a configuration without stator casing gap (as used for compressors with fixed stators). For the former case measurements of some aerodynamic parameters served as check for the CFD results. Despite some difficulties encountered during the calibration by means of CFD, the postulated correction led to a better agreement between the predicted performance by the stage stacking method and the measurements of built compressors.
Experimental Demonstration of the Tip Clearance Flow Resonance Behind Compressor Non-Synchronous Vibration
Authors: Thomassin Jean , Vo Huu Duc , Mureithi Njuki W.
Published: August 2009

Non-Synchronous Vibration (NSV) is a particular type of aero-elastic phenomenon where the rotor blades vibrate at non-integral multiples of the shaft rotational frequencies. NSV behaviour appears similar to off-design stall flutter but with a particular blade tip flow evolution. This paper demonstrates the link between NSV and the resonance induced by the tip clearance flow, based on a proposed hypothesis. At off-design operating conditions, the rotor blade tip clearance shear layer flow can evolve tangentially. It is proposed that this tangential flow becomes a support for an acoustic feedback wave that settles between rotor blades. The feedback wave is driven by the blade vibratory motion. This forms a closed loop system where the feedback wave synchronizes the shear layer vortical structures with the blade vibration frequency. Depending on the blade tip local temperature, and when the feedback wavelength matches within one or two blade pitches, the system becomes resonant and very high vibrations can occur on the blade. An axial stage compressor test rig is used to look into the underlying mechanism behind NSV. The experimental apparatus consists of the first stage of a High Pressure Compressor (HPC) driven by an electric motor. The test section is built to minimize the effects of the adjacent stator blade rows to isolate the role of rotor blade tip clearance flow on NSV. Sensitivity studies are carried out to assess the effects of the rotor blade tip clearance and inlet temperature on NSV. Finally, evidence of the staging phenomena, inherent to the proposed NSV mechanism, is experimentally obtained.